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Aerospace Science and Technology
Journal Prestige (SJR): 0.796
Citation Impact (citeScore): 3
Number of Followers: 382  
  Hybrid Journal Hybrid journal (It can contain Open Access articles)
ISSN (Print) 1270-9638
Published by Elsevier Homepage  [3185 journals]
  • Energy conservation of V-shaped swarming fixed-wing drones through
           position reconfiguration
    • Abstract: Publication date: November 2019Source: Aerospace Science and Technology, Volume 94Author(s): A. Mirzaeinia, M. Hassanalian, K. Lee, M. Mirzaeinia There is currently a growing interest in the area of drag reduction of unmanned aerial vehicles. In this paper, the swarming flight of the fixed-wing drones and a load balancing mechanism during the swarm is investigated. As an example, the swarm flight of EBee Sensfly flying wings is analyzed through the proposed methodology. The aerodynamic drag forces of each individual drone and the swarm are modeled theoretically. It is shown that drones through the swarming flight can save up to 70% of their energy and consequently improve their performance. As swarming drones have different loads and consume a different level of energy depending on their positions, there is a need to replace them during the flight in order to enhance their efficiency. To this end, regarding the number of drones, a replacement algorithm is defined for them so that they will be able to save more energy during their mission. It is shown that there is more than 21 percent improvement in flight time and distance of swarming drones after replacement. This method of replacement and formation can be considered as one of the effective factors in a drag reduction of swarming aerial vehicles.
  • Electric sail trajectory correction in presence of environmental
    • Abstract: Publication date: November 2019Source: Aerospace Science and Technology, Volume 94Author(s): Andrea Caruso, Lorenzo Niccolai, Giovanni Mengali, Alessandro A. Quarta An Electric Solar Wind Sail (E-sail) is an innovative propellantless propulsion system that generates a propulsive acceleration by exchanging momentum with the solar wind charged particles. Optimal E-sail trajectories are usually investigated by assuming an average value of the solar wind characteristics, thus obtaining a deterministic reference trajectory. However, recent analyses have shown that the solar wind dynamic pressure should be modeled as a random variable and an E-sail-based spacecraft may hardly be steered toward a target celestial body in an uncertain environment with just an open-loop control law. Therefore, this paper proposes to solve such a problem with a combined control strategy that suitably adjusts the grid electric voltage in response to the measured value of the dynamic pressure, and counteracts the effects of the solar wind uncertainties by rectifying the nominal trajectory at suitably chosen points. The effectiveness of such an approach is verified by simulation using two-dimensional transfer scenarios.
  • Dynamics and control of de-spinning giant asteroids by small tethered
    • Abstract: Publication date: November 2019Source: Aerospace Science and Technology, Volume 94Author(s): Junjie Kang, Zheng H. Zhu This paper investigates the de-spin control problem of giant asteroids by deploying small tethered spacecraft in the post-capture operation of asteroid redirection or space debris removal. Two simple and novel feedback control laws are developed based on dynamic characteristics of the tethered asteroid-spacecraft system. The first de-spins the asteroid by deploying tether at the tethered spacecraft to dissipate the kinetic energy of asteroid without consuming propellent. Although effective, this control law cannot de-spin the asteroid completely to the static state because of the stability deteriorates. To overcome the limitation, the second control law is developed by de-spinning the asteroid to the static state by deploying tether and firing a thrust at the tethered spacecraft simultaneously. The stability of both control laws is proved. Numerical simulation shows both control laws are effective.
  • The effect of hybridization on high-velocity impact response of carbon
           fiber-reinforced polymer composites using finite element modeling, Taguchi
           method and artificial neural network
    • Abstract: Publication date: November 2019Source: Aerospace Science and Technology, Volume 94Author(s): Mohammad Vahab Mousavi, Hadi Khoramishad In this paper, the effect of hybridization of carbon fiber-reinforced polymer (CFRP) composite laminates was investigated on the high-velocity impact responses of laminates using a validated finite element model, the Taguchi and the artificial neural network methods. The CFRP laminates were hybridized by replacing half of the carbon layers with glass and Kevlar laminae. It was found out that employing the right hybrid materials on the right position of the laminate can considerably change the damage pattern and consequently reduce the projectile residual velocity by increasing the energy absorbed by the CFRP laminate. Introducing four Kevlar laminae on the back of the CFRP composite laminate instead of the four carbon layers resulted the highest improvement in the laminate energy absorption by 67%. The hybrid laminates experienced wider damage zone and more extensive delamination compared to the CFRP laminate. Introducing the hybrid materials with high strength and strain to failure material such as Kevlar on the back of the laminate in the presence of CFRP front face can be considered as an appropriate hybrid composite laminate when exposed to high-velocity impact loading.
  • Experimental study of a bio-inspired flapping wing MAV by means of force
           and PIV measurements
    • Abstract: Publication date: November 2019Source: Aerospace Science and Technology, Volume 94Author(s): Shuanghou Deng, Jun Wang, Hanru Liu This study explores the aerodynamic performance of an X-Wing bio-inspired flapping wing micro air vehicle (MAV) underlying clap-and-fling motion by means of force and flow field measurements. Wind tunnel tests were first conducted to identify the operation region of current flapping wing MAV. The effects of flapping frequency and wing flexibility on force generation and power loading were evaluated using a miniature six-component force transducer. Flow visualization in the chordwise wake of the flapping wings is measured by means of phase-lock particle image velocimetry (PIV) measurement, results revealed the shedding behavior of the vertical structures by the wings during clap-and-fling motion showed a momentum surplus. Additionally, the in-ground-effect (IGE) is also examined aiming to understand the aerodynamic characteristics during take-off and landing of flapping wing MAV.
  • A novel methodology of sequential optimization and non-probabilistic
           time-dependent reliability analysis for multidisciplinary systems
    • Abstract: Publication date: November 2019Source: Aerospace Science and Technology, Volume 94Author(s): Lei Wang, Chuang Xiong Various uncertainties, which are usually time-dependent, affect the reliability of complicated engineering systems seriously. Considering the fact that only limited sample data of the uncertain variables can be obtained in engineering practice during the whole in-service time of multidisciplinary systems, the interval process model is introduced to model the time-dependent uncertain variables, and a non-probabilistic time-dependent reliability estimation model is proposed. In addition, a sequential multidisciplinary optimization and non-probabilistic time-dependent reliability assessment (SMO_NTRA) approach is developed to decouple the time-dependent reliability analysis from the multidisciplinary design optimization (MDO). In the framework of SMO_NTRA, the deterministic MDO and non-probabilistic time-dependent reliability analysis are executed in a sequential manner. Thus the computationally expensive double level optimization problem can be avoided and the efficiency can be greatly improved. Furthermore, the shifting distance of the constraint is calculated by bi-section method. Both numerical and engineering examples are employed to demonstrate the validity of the proposed method.
  • Impacts of jet angle and jet-to-crossflow pressure ratio on the mixing
           augmentation mechanism in a shcramjet engine
    • Abstract: Publication date: November 2019Source: Aerospace Science and Technology, Volume 94Author(s): Zhao-bo Du, Wei Huang, Li Yan, Min-zhou Dong The mixing process between the fuel and the incoming flow should be taken into consideration for the shock-induced combustion ramjet (shcramjet) engine, and this is pivotal for its successful engineering implementation. In the current study, the impacts of the gaseous fuel jet angle and the jet-to-crossflow pressure ratio are investigated numerically in order to make a foundational study and reveal the mixing augmentation mechanisms of different conditions. Flow fields windward and leeward are studied, and some parameters are provided to evaluate the flow field properties quantitatively, namely the mixing efficiency, the total pressure recovery coefficient, the fuel penetration depth and the mixing length. The obtained results predicted by the three-dimensional Reynolds-Averaged Navier-Stoke (RANS) equations coupled with the two equation shear stress transport (SST) k-ω turbulence model show that both the jet angle and the jet-to-crossflow pressure ratio have a great influence on the flow field, and the hydrogen distribution and streamwise vorticity downstream of the injectant orifice are affected seriously. The small jet angle and low jet-to-crossflow pressure ratio are more beneficial for the air-fuel mixing but with some shortcomings, and there needs an optimum result in the future to capture better achievement.
  • Autonomous onboard estimation of mean orbital elements for geostationary
           electric-propulsion satellites
    • Abstract: Publication date: November 2019Source: Aerospace Science and Technology, Volume 94Author(s): Lincheng Li, Jingrui Zhang, Shuge Zhao, Rui Qi, Yanyan Li Mean orbital elements estimation for geostationary (GEO) satellites is important for related studies, including station-keeping, rendezvous and end-of-life disposal. With increasingly limited operational slots in GEO region and the advance of all-electric-propulsion satellites, a fast and accurate mean orbital element estimation tool is necessary. In order to balance estimation precision and mission cost as well as to be independent of the ground station, this paper develops an autonomous onboard estimation method of the mean orbital elements for geostationary electric-propulsion satellites. Natural perturbations in GEO, including Earth's triaxiality, luni-solar attractions, and solar radiation pressure, are considered. Terms of appropriate orders due to these effects are chosen to model the semi-analytical dynamics, where modified short-period variations and differential mean orbital elements are derived. Regarding mean orbital elements as state variables and osculating orbital elements as measurements, a filter as well as analytical Jacobians is formulated to make the accurate estimation. Five scenarios are simulated to validate the accuracy and efficiency of the proposed method in the GEO region.
  • Effects of rotational motion on dynamic aeroelasticity of flexible
           spinning missile with large slenderness ratio
    • Abstract: Publication date: November 2019Source: Aerospace Science and Technology, Volume 94Author(s): Heng Li, ZhengYin Ye The structural rigidity of a spinning missile with large slenderness ratio is usually small, and the structural deformation and rate should not be ignored. Furthermore, rotational motion makes the aeroelasticity more complicated. Therefore, unsteady Euler equations and generalized dynamic aeroelastic equations are coupled simultaneously to simulate the dynamic aeroelastic response of a spinning missile with large slenderness ratio using rigid-motion mesh and radial-basis-function (RBF) morphing mesh techniques. The unsteady Euler equations are solved by computational fluid dynamics (CFD) technique by the in-house code. The Coriolis term and centrifugal loading term due to rotational motion are both considered in the generalized dynamic aeroelastic equations. The rigid-motion mesh and RBF morphing mesh techniques are both based on unstructured mesh, and the rigid-motion mesh is adopted to treat the rigid motion due to rotational motion, while the RBF morphing mesh is employed for flexible structural deformation caused by aeroelasticity. Numerical results of aeroelastic case are well agreed with the experimental results, which validates the numerical method. A missile model with X-X configuration is constructed to investigate the effects of rotational motion on dynamic aeroelasticity. The dynamic aeroelastic responses of the missile with and without rotational motion are simulated, respectively. Comparison results show that the lateral modes and longitudinal modes are coupled together because of rotational motion. In addition, the structural natural frequencies are changed due to rotational motion. In the end, detailed numerical analysis of the generalized dynamic aeroelastic equations used in this paper indicates the mechanism by which the rotational motion leads to the coupling of lateral modes and longitudinal modes and changes the structural natural frequencies.
  • Asymmetric heating and buoyancy effects on heat transfer of hydrocarbon
           fuel in a horizontal square channel at supercritical pressures
    • Abstract: Publication date: October 2019Source: Aerospace Science and Technology, Volume 93Author(s): Xing Sun, Hua Meng, Yao Zheng Heat transfer of the hydrocarbon fuel at a supercritical pressure is a key process in the aerospace engine cooling applications. A CFD model is applied in this paper to study flow dynamics and heat transfer of the aviation kerosene RP-3 in a horizontal square cooling channel under asymmetric heating and buoyancy effects at various supercritical pressures. The turbulent fluid flows are handled by the standard k–ε turbulence model with an enhanced wall treatment, and the strong thermophysical property variations are calculated using the extended corresponding state approaches and a four-component surrogate model of RP-3. The effects of secondary flows and heat flux redistribution induced by buoyancy on supercritical-pressure heat transfer are analyzed. Results indicate that drastic variations of the fuel density with temperature at a supercritical pressure of 3 MPa induce strong buoyancy effect on heat transfer. In the top heated case, secondary flows carry the relatively cold fluid to the side and opposite walls, thereby increasing their fin effectiveness in heat transfer. As the operating pressure increases from 3 to 5 MPa, the density variation of RP-3 is decreased, consequently leading to the weakened buoyancy effect. As the solid thermal conductivity increases from 20 to 100 W/(m-K), the surface heat flux redistribution is dictated by heat conduction in the channel walls, and therefore, the buoyancy effect is significantly reduced. Numerical results herein could help elucidate the heat transfer characteristics of RP-3 in practical engine cooling processes.
  • Preliminary engine design and inlet optimization of the MULDICON concept
    • Abstract: Publication date: October 2019Source: Aerospace Science and Technology, Volume 93Author(s): S. Zenkner, M. Trost, R. Becker, C. Voß In this study, aspects of the propulsion design for a highly swept wing configuration with military application are presented. These include the design of the thermodynamic cycle and the automated optimization of the inlet geometry. First, the thermodynamic design process of a conventional turbofan engine is described, which is intended to meet the requirements of a generic UCAV configuration. In addition to the thrust requirements derived from the mission profile, other constraints such as maximum fuel consumption, available installation space and aerodynamic and structural limits dimension the model in preliminary design. The resulting engine data are used as boundary conditions for the CFD simulation of the associated engine intake. This CFD calculation is part of an optimization process chain for determining the inlet duct geometry with low total pressure loss and low engine visibility. In order to estimate the effects of the different total pressure losses on the mission fuel consumption, the results obtained using CFD calculations are fed back into the performance simulation. This allows quantifying the influence of the inlet geometry on the global engine parameter. Based on the investigations and obtained results, the final configurations for the engine and intake are selected.
  • Wind component estimation for UAS flying in turbulent air
    • Abstract: Publication date: October 2019Source: Aerospace Science and Technology, Volume 93Author(s): C. Grillo, F. Montano One of the most important problem of autonomous flight for UAS is the wind identification, especially for small scale vehicles. This research focusses on an identification methodology based on the Extended Kalman Filter (EKF). In particular authors focus their attention on the filter tuning problem. The proposed procedure requires low computational power, so it is very useful for UAS. Besides it allows a robust wind component identification even when, as it is usually, the measurement data set is affected by noticeable noises.
  • Investigation on flutter stability of the DLR-F19/SACCON configuration
    • Abstract: Publication date: October 2019Source: Aerospace Science and Technology, Volume 93Author(s): Guido Voss, Dominik Schaefer, Cyrille Vidy In the present work, flutter stability studies of an unmanned flying-wing configuration are presented. For this purpose, different fidelity modeling methods (DLM, CFD-Euler and CFD-RANS) are considered. The dependence of flutter speeds on altitude and Mach number is examined, showing that aerodynamic potential-based methods cannot predict aerodynamic phenomena such as flow detachment occurring at high angles of attack. In this respect, it is important that flutter investigations industry-oriented calculation methods are compared with the results obtained by high-fidelity CFD methods.
  • An intelligent design method for actuation system architecture
           optimization for more electrical aircraft
    • Abstract: Publication date: October 2019Source: Aerospace Science and Technology, Volume 93Author(s): Zongxia Jiao, Bo Yu, Shuai Wu, Yaoxing Shang, Haishan Huang, Zhewen Tang, Renlei Wei, Chunfang Li The design of flight control actuation system is facing major challenge due to the development of more electrical aircraft. The task is to find the combinations of power sources, actuators and computers, which becomes more complex because of the new power sources and actuator types of more electrical aircraft. It is impossible to determine optimal architecture by traditional trial-and-error method within acceptable time. Therefore, the need for new methodology for actuation system architecture design emerges. This study proposes an intelligent design method which has steps of design space exploration of actuation system architectures by constraint satisfaction problem (CSP) method, safety assessment process to exclude unsafety solution, multi-objectives optimization to get Pareto optimal front and comprehensive decision for final architecture via analytic hierarchy process. And the design method is implemented in python and a software platform is developed. Furthermore, within the paper a case study for A350 flight control actuation system is presented to testify the application of this methodology. Compared to the traditional hydraulic architecture, the optimal architecture is more competitive in weight, power and cost. At the same time, the optimal architecture is found in less than 30 minutes among 1075 candidates, which greatly reduces the design cycle. This method deals with the problem in the design of flight control actuation system.
  • Aeroelastic global structural optimization using an efficient CFD-based
           reduced order model
    • Abstract: Publication date: November 2019Source: Aerospace Science and Technology, Volume 94Author(s): Dongfeng Li, Andrea Da Ronch, Gang Chen, Yueming Li In aeroelastic structure optimization, the structural model needs to be modified repeatedly to meet all targets. Although computational fluid dynamics (CFD) based reduced order models (ROMs) have been successfully applied to transonic aeroelastic analysis, the existing CFD-based ROMs are at a fixed flight condition for a frozen aeroelastic model configuration. The aerodynamic and structural model have to be reconstructed to ensure accuracy, when a structural modification was made. These reconstructions take a considerable time and the computational costs become prohibitive. To overcome the realistic challenge, we have developed an efficient CFD-based ROM, which is robust to aeroelastic system. This paper presents a new optimization process using the efficient method for aeroelastic global structural optimization. The optimization process employs Genetic Algorithms (GAs) as optimization tool. In order to assess the performance of presented optimization process, the AGARD 445.6 wing model is taken as numerical example. The results show that the most feasible and optimal solutions are effectively obtained by the presented optimization process.
  • Effect of film cooling injection on aerodynamic performances of scramjet
    • Abstract: Publication date: Available online 6 September 2019Source: Aerospace Science and TechnologyAuthor(s): Kuan Zheng, Wei Tian, Jiang Qin, Silong Zhang, Hui Hu With continuous increase in flight Mach number, aerodynamic heating on scramjet isolators becomes increasingly pronounced. Consequently, a thermal protection technique for scramjet isolators is urgently required. In the present study, a numerical investigation was conducted to evaluate the feasibility of applying film cooling on a scramjet isolator. First, the heat transfer characteristics under different coolant flow conditions (i.e., coolant Mach number, coolant total temperature, and injection position) were obtained to evaluate the film cooling efficiency on a scramjet isolator. Next, the characteristics of maximum backpressure and friction drag were analyzed to obtain the effects of film cooling injection on the aerodynamic performances of the scramjet isolator. It was found that the film cooling injection under proper coolant flow conditions could reduce the friction drag and enhance the ability of the isolator to resist backpressure, which is beneficial to the performances of a scramjet engine. In general, the results obtained in this study indicated that film cooling injection could be a practical thermal protection technique for a scramjet isolator.
  • Network topology of turbulent premixed Bunsen flame at elevated pressure
           and turbulence intensity
    • Abstract: Publication date: November 2019Source: Aerospace Science and Technology, Volume 94Author(s): Jinhua Wang, Yaohui Nie, Weijie Zhang, Shilong Guo, Meng Zhang, Zuohua Huang For turbulent premixed flame surface, folded regions will be formed mainly depending on the turbulence-flame interaction and are significant structures which can increase flame surface area at large scale and enhance the local displacement speed. The conventional methods, such as PDF distribution of curvature, may not properly label the folded regions of turbulent flame because these regions intrude deeply into the products, resulting in a small curvature probability. This problem may be aggravated at elevated pressure and turbulence intensity conditions when more folded regions appear at the turbulent flame front. To identify the folded regions, Network topology of turbulent premixed flame front is constructed using the “visibility” method. Results show that this method can convert the spatial signal of turbulent flame to network topology, labeling the folded regions. Compared with curvature PDF, node degree distribution of network can reflect the mechanism of turbulence-flame interaction when the non-dimensional turbulence intensity is increased by different ways. The network structure of turbulent flame will transfer from sparse to condense when the dimensionless turbulence intensity is increased by pressure and perforated plate, as these two methods will extend the turbulence-flame interaction time and promote the interaction intensity, respectively. However, although the dimensionless turbulence intensity will increase with the augment of outlet velocity, the node degree distribution of network structure of turbulent flame front keeps almost constant. This is caused by the reduced turbulence-flame interaction time. It suggests that the turbulence-flame interaction time is an factor as important as the dimensionless turbulence intensity in turbulent premixed combustion. For forced-turbulent premixed flame at elevated turbulence intensity, the “bending phenomenon” will be hidden if the outlet velocity is not taken into account, as the outlet velocity is related to the turbulence-flame interaction time.
  • A novel non-singular terminal sliding mode control-based integrated
           missile guidance and control with impact angle constraint
    • Abstract: Publication date: November 2019Source: Aerospace Science and Technology, Volume 94Author(s): Chao Ming, Xiaoming Wang, Ruisheng Sun A novel impact angle constrained integrated guidance and control (IGC) scheme is proposed with non-singular terminal sliding mode control (NTSMC) technique for missile to intercept a maneuverable target. Firstly, an IGC design model is innovatively established in a strict feedback form which selects the normal overload as the system state in replace of the angle of attack so that the system states both are measurable. On this basis, an IGC law is proposed to guarantee the impact angle converge to the expected command within a finite time and to improve the robustness against the multiple disturbances which include the model parameter uncertainties and the target maneuvering by incorporating the NTSMC method and the extended state observer (ESO) technique. In order to avoid the chattering phenomenon, the smooth derivative of the intermediate control command which is obtained by the introduced nonlinear tracking differentiator (TD) is adopted into the control loop. Then, the stability of closed-loop integrated system is rigorously proved by utilizing the Lyapunov theorem. Finally, extensive contrast simulation results are presented to validate the efficiency, superiority and robustness of the suggested IGC scheme.
  • Research on an active pitching damper for transonic wind tunnel tests
    • Abstract: Publication date: November 2019Source: Aerospace Science and Technology, Volume 94Author(s): Lei Zhang, Yuke Dai, Xing Shen, Xiping Kou, Li Yu, Bo Lu In wind tunnel tests, the sting with a model and a balance installed forms an elastic beam. To reduce the aerodynamic interference, the slenderness ratio of the sting is usually small, which obviously results in low damping of the whole structure. In that case, violent oscillation appears quickly while the sting is exposed to varying aerodynamic loads, which finally leads to inaccurate data and even damage to the structure. In this paper, the vibration of the model-balance-sting system was analyzed experimentally, and the concepts of active pitching damper were detailed in this paper. Based on independent modal space control, linear quadratic regulators were designed with identified state-space models. The active pitching damper was successfully validated by exerting impulse loading in model preparation bay firstly and then tested in a transonic wind tunnel. Experimental results show that the maximum attenuation of vibration reached 91% in standard deviation and extension of the angle of attack was 6°. Moreover, wind tunnel tests proved the effectiveness of the active pitching damper working on different test configurations, which indicates the superiority of the damper.
  • Numerical simulation of fluid forces on moving solid body by the vortex in
           cell method with volume penalization
    • Abstract: Publication date: November 2019Source: Aerospace Science and Technology, Volume 94Author(s): Tomohiro Degawa, Qiang Gu, Tomomi Uchiyama, Kotaro Takamure This study proposes a method for calculating the fluid forces acting on moving solid bodies in flows simulated by the vortex in cell (VIC) method. The body is represented using the volume penalization (VP) method, which introduces the influence of the body on the flow as an external force in the form of a penalization term into the Navier-Stokes equation. The pressure field is computed by solving the Poisson equation, which is derived by taking the divergence of the Navier-Stokes equation. The pressure and viscous stress on the body surface are calculated using a quadratic function along the vector normal to the surface. This study simulates the flow around a cylinder oscillating in still water by the VIC-VP method, and calculates the fluid forces acting on the cylinder using the proposed method. The velocity profiles on four cross-sections at three different time points agree well with the results of a measurement. The in-line force acting on the cylinder also agrees with the existing results computed by a boundary-conforming formulation. These demonstrate the validity of the authors' method for calculating the fluid forces on solid bodies in flows simulated by the VIC-VP method.
  • The performance of droplet evaporation model in predicting droplet
           dynamics and thermal characteristics for R134a single isolated droplet and
           two-phase flashing spray
    • Abstract: Publication date: October 2019Source: Aerospace Science and Technology, Volume 93Author(s): Zhi-Fu Zhou, Guan-Yu Lu, Dong-Qing Zhu, Lu Zhang, Jia-Feng Wang, Bin Chen Flashing spray with low saturation point and high volatility mediums is of great importance in aerospace field. It involves complex droplet dynamics and heat and mass transfer processes in a turbulent, two phase flow. This paper comparatively evaluates the predictive performance of a selected number of droplet evaporation models that focus on convective and blowing effects. The studies span from a single, isolated R134a droplet that evaporates in a convective environment, to a fully turbulent, flashing spray formed through an accidental release of high pressure R134a liquid. An in-house developed code for single isolated droplet evaporation and a modified sprayFoam solver in OpenFOAM for flashing spray are used to calculate droplet and spray behaviors. The results show droplet evaporation model greatly affects the evolutions of droplet diameter, velocity and temperature for single isolated R134a droplet, that the C-R-S model predicts the lowest droplet diameter and velocity, and highest droplet temperature; the H-N-R model predicts the largest droplet velocity and lowest temperature; the A-S and N-G-R-M models predict almost identical results. In contrast to the great impact on droplet evolution for single isolated droplet modeling, droplet evaporation model has little influence on spray and thermal characteristics for R134a two phase flashing spray simulation. However, the A-S model predicts quite different radial profile of droplet temperature at spray periphery compared with other models, which is much lower than the experimental value.
  • DNS analysis of the effects of combustion on turbulence in a supersonic
           H2/air jet flow
    • Abstract: Publication date: October 2019Source: Aerospace Science and Technology, Volume 93Author(s): Yaowei Fu, Changping Yu, Zheng Yan, Xinliang Li In this paper, the effects of combustion on turbulence features are investigated by direct numerical simulation (DNS) of supersonic round turbulent hydrogen jet flows. Two DNS cases are conducted: the one is combustion hydrogen/air jet and the other is non-combustion jet with the same components as a comparison case. In the DNS, the supersonic jet consists of 85% hydrogen and 15% nitrogen in volume with temperature of 305 K and jet velocity of 900 m/s. The ambient air velocity is 20 m/s and temperature is 1150 K. The Reynolds numbers based the jet exit diameter and hydrogen jet velocity both are 22000 and the jet Mach numbers both are 1.2. The DNS results show that occurrence of combustion significantly delays the transition of the jet flow, and results in a 38% decrease in the peak of turbulent kinetic energy, compared with the non-combustion jet flow. The decreasing of the Reynolds number due to the heat release is considered as the main reason. Compared with the non-combustion jet flow, the combustion jet flow has more complete coherent vortex rings in the early stage. In addition, the positive and minus energy fluxes of combustion jet flow are always larger than that of non-combustion jet flow.
  • Surrogate-based aerodynamic shape optimization of hypersonic flows
           considering transonic performance
    • Abstract: Publication date: October 2019Source: Aerospace Science and Technology, Volume 93Author(s): Fei Liu, Zhong-Hua Han, Yang Zhang, Ke Song, Wen-Ping Song, Feng Gui, Ji-Bin Tang Aerodynamic shape optimization of a hypersonic transport aircraft over a wide Mach-number range is challenging. The difficulty is not only associated with the large computational cost of high-fidelity CFD (computational fluid dynamics) simulations but also linked to the reasonable compromise between aerodynamic performances of aircraft at different speed ranges. This article proposes to use efficient global optimization based on surrogate models to address this type of problems. A RANS (Reynolds averaged Navier-Stokes) flow solver is adopted to evaluate objective and constraint functions; Kriging surrogate model combined with a parallel infill-sampling method and a multi-round strategy are employed to find the global optimum. First, a profile optimization with baseline airfoil of NACA64A-204 is conducted to achieve high lift-to-drag ratio (L/D) at both hypersonic (Ma=6.0) and transonic (Ma=0.8) regimes with up to 18 design variables, and significant improvement has been observed. Then, multi-objective wing optimizations of maximizing both hypersonic and transonic L/Ds with up to 54 design variables are performed and multiple optimizations with various sets of weight coefficients are investigated. The optimized wings are further evaluated and compared with the baseline wing at the flow regimes from subsonic to hypersonic speeds. Results show that by using the optimized profile, hypersonic and transonic L/Ds of a typical wing configuration are increased by 13.85% and 7.32%, respectively. After 3-D optimizations, the L/Ds at hypersonic and transonic design points are further improved by 4.47% and 3.19%, respectively, and a better performance over a wide Mach-number range is obtained. It is shown that a multi-round surrogated-based optimization is feasible and effective for aerodynamic shape optimization of hypersonic transport aircrafts over a wide Mach-number range.
  • Three-dimensional numerical study on rotating detonation engines using
           reactive Navier-Stokes equations
    • Abstract: Publication date: October 2019Source: Aerospace Science and Technology, Volume 93Author(s): Li-Feng Zhang, John Z. Ma, Shu-Jie Zhang, Ming-Yi Luan, Jian-Ping Wang Three-dimensional simulations on the rotating detonation engine (RDE) are conducted by solving the Navier-Stokes equations in stoichiometric hydrogen/air mixtures. Free-slip wall and non-slip wall boundary conditions are considered. The emphasis is placed on the effects of the viscosity, mass diffusion and thermal conduction on the detonation wave in the RDE, especially in the boundary layer. The simulation results show that the detonation wave propagates with a lower velocity in the Navier-Stokes solutions compared with that in the Euler solutions. This is mainly due to the effects of mass diffusion and thermal conduction. Meanwhile, the two effects above reduce the pressure of the detonation wave in the mainstream. Moreover, the detonation wave height is higher in the Navier-Stokes solutions than that in the Euler solutions. The temperature behind the detonation wave near the walls is much higher than that inside the chamber due to the existence of the boundary layer. Therefore, the viscosity, mass diffusion and thermal conduction impact the detonation velocity, pressure, the detonation wave height and temperature in the RDE.
  • Thermal flutter prediction at trajectory points of a hypersonic vehicle
           based on aerothermal synchronization algorithm
    • Abstract: Publication date: Available online 5 September 2019Source: Aerospace Science and TechnologyAuthor(s): Tongqing Guo, Ennan Shen, Zhiliang Lu, Di Zhou, Jiangpeng Wu Due to orders of magnitude differences in time scale between structural heat transfer and aeroelastic responses, one-way aerothermal-aeroelastic coupling is adopted to develop a thermal flutter prediction method for a hypersonic vehicle operating along a desired trajectory. In view of the strong dependency of the heat transfer process on the unsteady hypersonic trajectory, an aerothermal synchronization algorithm is established in a non-inertial frame of reference by formulating the governing equations of fluid flow and heat transfer into a unified form. Then the heated free-vibration frequencies and mode shapes are calculated at each trajectory point by using a finite-element analysis. Consequently, the flutter computations are performed on the transiently heated structure at each trajectory point by utilizing a coupled computational fluid dynamics (CFD)/computational structural dynamics (CSD) method. Because of the mass dissimilarity caused by directly increasing the dynamic pressure of a compressible flow, the technique of variable stiffness is introduced to evaluate the flutter dynamic pressure at the point of mass similarity and the stiffness margin of flutter. The present method is applied to the thermal flutter computations of a hypersonic all-movable rudder operating along a given trajectory. The computed temperature differences between the synchronization and conventional partitioned methods, and the significant effects of aerodynamic heating on the structural modes and the flutter characteristics are analyzed in detail.
  • Station-keeping strategy for real translunar libration point orbits using
           continuous thrust
    • Abstract: Publication date: Available online 5 September 2019Source: Aerospace Science and TechnologyAuthor(s): Yi Qi, Anton de Ruiter In this paper, we investigate the station-keeping strategy for translunar libration point orbits (LPOs) using the continuous thrust in the ephemeris model. Ephemeris models with and without the solar radiation pressure (SRP) are proposed for the numerical simulation. Three kinds of translunar LPOs, including halo orbits, vertical Lyapunov orbits, and Lissajous orbits, are used as nominal orbits. A station-keeping strategy based on the backstepping technique is extended to the ephemeris model. Then station keeping under practical constraints, caused by the navigation system and the executive system, is studied in ephemeris models with and without the SRP. Using numerical simulations, influences of the dead-band scheme, navigation errors, the navigation interval time and the area-to-mass ratio on station keeping are discussed, respectively. Furthermore, numerical simulations indicate that if the SRP is taken into account, LPOs obtained in the ephemeris model with the SRP is more preferable as nominal orbits than those without the SRP. More Monte-Carlo simulations testify that our control strategy can be applied to the long-term station keeping for different translunar LPOs.
  • Using approximate similitude to design dynamic similar models
    • Abstract: Publication date: Available online 4 September 2019Source: Aerospace Science and TechnologyAuthor(s): Afshin Banazadeh, Pedram Hajipouzadeh This research deals with the analysis of approximate similitude between the dynamic similar models and the full-scale prototype of an aircraft. Due to physical and technical constraints, a full dynamic similarity is not practically possible and previous works have all neglected one or two similarity criteria like Mach or Reynolds numbers for the sake of Froude number similarity. In this work, it is shown that Mach number has an important effect on aerodynamic characteristics and dynamic response of an aircraft and that neglecting it makes the generalization of the scale-model test data invalid for the full-scale prototype. In order to address this problem, a measurable quantity named approximate similitude parameter is proposed, which takes into account both dynamic similarity criteria and the scale model test objectives. The effectiveness of this parameter is shown by a case study to find out optimal weight, moments of inertia and velocity for the scale model to achieve optimal dynamic similarity.
  • Optimization design method for the cable network of mesh reflector
           antennas considering space thermal effects
    • Abstract: Publication date: Available online 4 September 2019Source: Aerospace Science and TechnologyAuthor(s): Rui Nie, Baiyan He, Shaoze Yan, Xiaofei Ma The mesh reflector antenna is widely used in space satellites for its characteristics of lightweight, large aperture, high precision, and high stiffness. As the form of mesh surface is heavily depended on cable forces and vice versa, the form finding and optimization design of cable networks play an essential role in antenna design. The present methods usually neglect the thermal effects, which is reasonable for ambient temperature. While space thermal loads cause significant variations from the original form finding state, it is essential to investigate and minimize the antenna's on-orbit shape errors caused by the thermal effects. The active shape adjustment on-orbit is far from anticipated, while the preliminary adjustment before launch is also limited by the number of adjustable cables (tension ties) and the engineer's experiences. Under these circumstances, we propose a form finding and design optimization approach for the cable network of mesh reflector antennas considering space thermal effects. In this approach, the cable's thermal deformation and geometric nonlinearity are fully considered at the design stage, and the burden of shape adjustment for thermal errors before launch can be relieved. As all cables instead of adjustable cables are optimized, the improvement of surface accuracy is more obvious than the shape adjustment method. Our work provides a new idea and approach for the optimization design of cable networks and the minimization of on-orbit shape errors.
  • Distributed adaptive vibration control for solar power satellite during
           on-orbit assembly
    • Abstract: Publication date: Available online 4 September 2019Source: Aerospace Science and TechnologyAuthor(s): Enmei Wang, Shunan Wu, Zhigang Wu, Gianmarco Radice The distributed adaptive vibration control for solar power satellite (SPS) during on-orbit assembly is investigated in this paper. Focusing on the platform configuration SPS, the different control units (CUs) and relationship matrices, that describe the topology of whole SPS structure, are firstly defined, and the dynamic model of the CU is proposed through the relationship between the CU and the whole SPS structure. Using the linear quadratic optimal control technique, the adaptive controller of the CU is designed to suppress vibrations, and a compensatory component is included in the controller to deal with the interactions with the whole structure. The closed-loop distributed control system is then developed in the presence of the impact at the specific assembling location, which needs to adaptively update along with the varying dynamic model of the SPS structure during on-orbit assembly. The asymptotic stability of the closed-loop distributed system is investigated, and two representative numerical cases are finally provided. The results demonstrate that the proposed distributed adaptive controllers can significantly suppress the vibration caused by the impact among SPS modules during on-orbit assembly.
  • H -based model following method in autolanding
    • Abstract: Publication date: Available online 4 September 2019Source: Aerospace Science and TechnologyAuthor(s): K. Tamkaya, L. Ucun, I. Ustoglu Probably the most important part during a flight is the landing phase because most of the accidents occur in this phase. Automatic landing systems (ALS) take over the control during this phase to avoid potential pilot-induced risks. However, some external disturbances such as windshear can jeopardize the safe landing. In this paper, the flare part of ALS is handled in a different way. A combination of some useful design methods is brought together to improve the performance of the conventional ALS even under severe weather conditions. Model following method is combined with the H∞ synthesis method to find out the optimal solution for a given cost function. Resultant H∞ optimal control problem is solved using Linear Matrix Inequalities (LMIs) and then a dynamic controller is constructed. On the other hand, the overall system is formed into P-K configuration, thus the system can be reconfigured easily when there exists a change in the system such as addition or removal of disturbance, noise and so on. We achieved significant performance on the system without any disturbance. In addition to that, the robustness takes an important role for the flight systems and needs to be handled correctly. Therefore, two kinds of windshear are taken care of and their effects minimized in a way that the tracking performance remains unaffected. Thus, highly considerable results are obtained using the proposed method even under severe weather conditions.
  • An experimental study on a coaxial flow with inner swirl: Vortex evolution
           and flow field mixing attributes
    • Abstract: Publication date: Available online 4 September 2019Source: Aerospace Science and TechnologyAuthor(s): A. Giannadakis, A. Naxakis, A. Romeos, K. Perrakis, Th. Panidis A 2D particle image velocimetry study of a coaxial flow with inner swirl is presented. An inner swirling jet, produced by tangential injection, interacts with an annular flow generating a recirculating flow field with strong mixing attributes. The characteristics of the cross-plane velocity components of four different test cases are presented (two levels of tangential injection flow rate combined with two levels of annular flow rate) in order to study the mean and turbulent attributes of the swirling vortex. The main features of this complex flow field, which can be considered as the interaction of a typical swirling jet undergoing “vortex breakdown” with an outer annular flow with “backward facing step flow” characteristics, are investigated, focusing on the swirling jet's characteristics. The analysis of the mean and turbulent flow is based on a modified Rossby number, previously proposed by the authors, defined as the ratio of the streamwise velocity jump across the two streams over a typical tangential velocity, which is shown to represent the ratio of the pressure difference due to the streamwise velocity difference and the entrainment of the two flows to that due to the rotation of the swirling vortex. The angular momentum diffusion downstream is evaluated, to assess the mixing between the swirling vortex and the outer flow.
  • Onboard satellite visibility prediction using metamodeling based framework
    • Abstract: Publication date: Available online 4 September 2019Source: Aerospace Science and TechnologyAuthor(s): Xinwei Wang, Chao Han, Pengbin Yang, Xiucong Sun Satellite autonomous systems are employed to address complex space applications through onboard data processing and mission planning. To take advantage of onboard autonomous systems, rapid onboard satellite visibility predictions are necessary for certain decision-making missions, including Earth observation resource allocation and satellite data transmission. We consider this visibility prediction process as a roots-finding problem for a multiple hump function, and design a metamodeling-based framework with a self-adaptive interpolation method. Metamodels are developed as surrogates for visibility prediction functions to reduce expensive computational costs. Our proposed framework has a broad range of applications for all orbital types and orbit propagators. We conduct the experiments using different metamodeling techniques, radial basis functions, Kriging, and support vector regression based upon real China's satellites. Numerical simulations indicate that the proposed framework outperforms existing interpolation methods, efficiently reducing the onboard computational cost.
  • Experimental investigation of gravity and channel size effects on flow
           boiling heat transfer under hypergravity
    • Abstract: Publication date: Available online 3 September 2019Source: Aerospace Science and TechnologyAuthor(s): Xiande Fang, Da Tang, Ling Zheng, Guohua Li, Yuliang Yuan Understanding of flow boiling heat transfer under hypergravity is needed due to its applications to modern flight vehicles. Few investigations on this issue have been reported. The present paper presents the experimental results of R134a flow boiling heat transfer in 1.002 and 2.168 mm ID tubes under hypergravity up to 3.16 g. The results reveal effects of gravity, channel size, heat flux, mass flux, quality, and pressure on and their interrelations in flow boiling heat transfer. Gravity has strong effects on flow boiling heat transfer. The heat transfer coefficients (HTCs) under hypergravity levels up to 3.16 g are normally greater than those under Earth's gravity. The heat transfer characteristics in the 1.002 and 2.168 mm tubes are very different, indicating that gravity effects on flow boiling heat transfer strongly interrelate with channel size. Also, gravity effects on flow boiling heat transfer are influenced by heat flux, mass flux, and pressure, and somewhat related to quality.
  • Control of Aeolian tones from a circular cylinder using forced oscillation
    • Abstract: Publication date: Available online 30 August 2019Source: Aerospace Science and TechnologyAuthor(s): Ruixian Maa, Zhansheng Liu, Guanghui Zhang, Con J. Doolan, Danielle J. Moreau Effects of forced transverse oscillation on the generation of sound from a circular cylinder immersed in uniform flow at a Reynolds number 150 and a Mach number 0.2 is investigated by direct numerical simulation. The cylinder is prescribed to oscillate sinusoidally with a constant oscillating amplitude ratio α of 0.2 of the cylinder diameter and oscillation frequency ratios F=0.2 to 1.4 of the inherent vortex shedding frequency. The impact of the oscillating frequency on the sound pressure is in accordance with that of the forces acted on the cylinder, by which three regimes are identified. In the first regime (F≤0.7), the sound levels are slightly affected by the oscillation. In the second regime, at 0.8≤F
  • Premixing-type liquefied gas bipropellant thruster using nitrous
           oxide/dimethyl ether
    • Abstract: Publication date: Available online 29 August 2019Source: Aerospace Science and TechnologyAuthor(s): Akira Kakami, Atsushi Kuranaga, Yasuyuki Yano This paper describes a nitrous oxide (N2O)/dimethyl ether (DME)-based bipropellant thruster in which the propellant gases are premixed before they enter the thrust chamber. Monopropellant (hydrazine) and bipropellant (nitrogen tetroxide (NTO)/hydrazine) thrusters are used in the onboard propulsion systems on a spacecraft and exhibit high performance. However, hydrazine and NTO are strongly toxic and exhibit high reactivity with respect to various materials. Further, their relatively high freezing points and low vapor pressures make it necessary to use heaters for the storage tanks and tubes and the pressurant feed systems. Hence, we propose an N2O/DME liquefied gas bipropellant thruster. The bipropellant is less toxic compared to hydrazine/NTO and has an adequate vapor pressure and low freezing point, which makes complex temperature management systems unnecessary. We tested a 0.4-N class prototype wherein N2O and DME were made to flow directly into the thrust chamber. The prototype exhibited a low characteristic velocity (C*) efficiency at 60%. Hence, a prototype with a propellant mixer was designed to enhance the C* efficiency and reduce the size of the thrust chamber. A coil-type mixer was developed to mix the fuel and oxidizer based on the secondary flow induced in curved tubes. Thrust measurements showed that this premixing-type prototype had a C* efficiency of 84.5% at an O/F ratio of 3.5.
  • Adaptive prescribed performance control for the post-capture tethered
           combination via dynamic surface technique
    • Abstract: Publication date: Available online 28 August 2019Source: Aerospace Science and TechnologyAuthor(s): Yingbo Lu, Panfeng Huang, Zhongjie Meng In this paper, an adaptive dynamic surface controller with prescribed performance is proposed for the stabilization control of the post-capture tethered combination, where the factors such as measurement uncertainties, model uncertainties, external disturbances and input saturation are considered. Compared with the most prescribed performance approaches, the proposed controller has a simple structure, and needs no error transformations. Firstly, we found a dynamic model of the post-capture tethered combination considering the integrated attitudes of the combination. Then, taken the above-mentioned factors into consideration, the dynamics of the post-capture combination can be transformed to a multi-input-multi-output (MIMO) system with mismatched and matched uncertainties. An adaptive approach is proposed for estimating the upper bound of the uncertainties. Furthermore, an auxiliary dynamic system is adopted to deal with the problem of control input saturation, and the bounded assumption of input error is released. All the signals in the closed-loop system are confirmed to be ultimately bounded by the Lyapunov stability theory. Finally, simulation results validate the performance and robustness improvement of the proposed control algorithm.
  • Roll control for single moving-mass actuated fixed-trim reentry vehicle
           considering full state constraints
    • Abstract: Publication date: Available online 28 August 2019Source: Aerospace Science and TechnologyAuthor(s): Kaixu Dong, Jun Zhou, Min Zhou, Bin Zhao Moving-mass control system applies internal movable mass elements as actuators to shift the center of mass of the vehicle relative to the external aerodynamic forces to generate control torques. Although moving-mass control system provides sufficient control authority, it suffers from its nature of nonlinear and high-coupling. What's more, the design of the moving-mass control system confronts state constraints arising from two aspects: (1) physical limitation: the internal space of the vehicle is limited and the moving-mass element can not translate out of the shell; (2) performance requirements: limiting the maximum velocity of the moving-mass element can decrease the disturbances exerting on the vehicle significantly. Based on the purpose of controlling the roll attitude of a moving-mass actuated vehicle without violating any state constraint, we design a controller using integral barrier Lyapunov functionals to control the system and keep all the states bounded in the meanwhile. The backstepping procedure is adopted and the dynamic surface control scheme is applied to avoid the 'terms explosion' in that procedure. An adaptive law is proposed to diminish the influences caused by uncertainties whose boundaries are unknown. The performance of the proposed controller is illustrated by numerical simulations under various conditions.
  • A three variable refined shear deformation theory for porous functionally
           graded doubly curved shell analysis
    • Abstract: Publication date: Available online 26 August 2019Source: Aerospace Science and TechnologyAuthor(s): Minh-Chien Trinh, Seung-Eock Kim This study develops a three variable refined shear deformation theory to analyze the free vibration and bending behavior of porous functionally graded doubly curved shallow shells subjected to uniform and sinusoidal pressure. Shell displacements are assumed to be caused by extensional, bending, and shear effects. The in-plane displacements produced by bending effects are considered taking the form of the classical plate theory. The in-plane displacements produced by shear effects satisfy the stress-free and strain-free condition at the top and bottom surfaces, eliminating the usage of the shear correction factor in the present study. Two porosity types influence material properties and structure behaviors in different aspects. Hamilton's principle is used to derive Euler–Lagrange equations. Spatial solutions for the differential equation are assumed satisfying boundary conditions and their time-dependent amplitude equations are obtained by applying the Bubnov–Galerkin technique. Natural frequencies and transverse deflections of the shell in different geometry configurations and different porosity types and degrees are obtained and compared. The proposed theory is proved feasible to be applied in the analysis of functionally graded plates and shells with porosity.
  • Mathematical modeling of turbulent boundary layers, modified by
           wall-localized drag reduction techniques
    • Abstract: Publication date: October 2019Source: Aerospace Science and Technology, Volume 93Author(s): Yevhenii Shkvar, Shi-ju E, Andrii Kryzhanovsky Skin friction drag reduction via turbulence manipulating implies a purposeful influence on a very complicated mechanism of vortical structure dynamics with wide range of scales and the disturbances localized in the wall vicinity. The goal of this research is to propose the complex unified approach for semi-empirical modeling of turbulent viscosity with possibility to take into account both simultaneous and separate influence of such factors as: streamlined surface relief (from macro ribs to microgrooves), uniform and intermitted microblowing, polymer additives injection and Large Eddy BreakUp devices. There are three principal advantages of the developed model: 1) It has hybrid structure, joining algebraic (in the wall region) and differential (in the wake region) modeling descriptions, that makes them very flexible in various cases of different turbulent flows prediction; 2) It is built on the base of common principles and its potential generalization to the case of heat conductivity modeling doesn't require the use of Reynolds analogy (incorrect in the vicinity of the rough surface) in the wall region of turbulent boundary layer; 3) The proposed approach is not limited by only RANS level of flow modeling, it can be effectively adapted for improving the near-wall subgrid turbulence model in frames of LES/DES techniques. The elaborated model due to its ability to take into account the mentioned above factors both independently and in different combinations demonstrate universal approach that allows to simulate simultaneous acting the different flow control methods, investigate their non-linear interaction and optimize geometric characteristics and operating conditions for searching the synergistic modes of functioning. The applicability of the developed model is proved by qualitative agreement of wide range of mentioned above kinds of turbulent flows predictions and the experimental results, obtained by different researchers. The proposed modeling approach can be recommended for its application in different research and engineering fields, associated with turbulence vortical structure manipulating.
  • On-line pattern discovery in telemetry sequence of micro-satellite
    • Abstract: Publication date: October 2019Source: Aerospace Science and Technology, Volume 93Author(s): Dong Xu, Gaofei Zhang, Zheng You In micro-satellite engineering, telemetry data is the only basis for the ground staffs to judge the condition of spacecraft in-orbit, and is also a critical reference for testing, operation, and Prognostics and Health Management (PHM). In practice, the analysis and processing of raw telemetry data is very complicated and cumbersome. To maintain the micro-satellites' performance and ensure their reliability, it is important to implement on-line interesting pattern discovery in the ground station monitoring system or Electrical Ground Support Equipment (EGSE) system. The goal of this paper is to monitor evolving telemetry sequence of micro-satellite, and to discover subsequences that are similar to the given query sequence, under the Dynamic Time Warping (DTW) distance. However, in the processing of telemetry sequence, massive amounts of data arrive continuously and it is infeasible to directly use DTW distance. Therefore, this paper improves the DTW and proposes a novel method (eDTW) that can perform pattern discovery on telemetry sequence online. First, the sub-sequence time warping matrix is constructed by taking the telemetry sequence as the X-axis and the query sequence as the Y-axis; Second, calculate the DTW distance from each unit, and the initial position of each path; Finally, determine the current local optimal subsequence and determine the next starting point of the next discovery. Experiments with telemetry sequence of Tsinghua University's NS2 satellite show that eDTW can discover interesting pattern quickly, with no buffering of stream values and without comparing pairs of streams. Our experimental case studies show that eDTW can incrementally capture correlations and discover trends, efficiently and effectively.
  • Design and analysis of a scissors double-ring truss deployable mechanism
           for space antennas
    • Abstract: Publication date: October 2019Source: Aerospace Science and Technology, Volume 93Author(s): Bo Han, Yundou Xu, Jiantao Yao, Dong Zheng, Yongjie Li, Yongsheng Zhao Space deployable mechanism plays an important role in the orbital deployment and stable support of the antenna reflection surface, which is an important component for the spacecraft radar antenna. In order to improve the structural stiffness of the ring truss deployable antenna mechanism when it has a large diameter, a scissors double-ring truss deployable mechanism is proposed in this paper. First, structure analysis of general ring truss deployable antenna and construction of the scissors double-ring truss deployable mechanism are conducted, and the whole mechanism is decomposed to a plurality of mechanism units. Then, degree of freedom (DOF) of the scissors double-ring truss deployable mechanism is analyzed based on screw theory, the result showed that it has only one DOF. Furthermore, based on screw theory, kinematic characteristics of the scissors double-ring truss deployable mechanism are examined, velocities and accelerations of the components in the mechanism are obtained, as well as the Jacobian matrixes. Finally, based on Newton-Euler equation and the principle of virtual work, a dynamic model of the whole mechanism is established, numerical calculation and simulation verification are carried out, and the results verified the correctness of the theoretical analysis. The scissors double-ring truss deployable mechanism proposed in this paper can be well applied in the field of space antennas, and the theoretical analysis method based on screw theory used in this paper can provide insights into other spatial deployable mechanisms.
  • Study of combined flow control strategies based on a quantitative analysis
           in a high-load compressor cascade
    • Abstract: Publication date: October 2019Source: Aerospace Science and Technology, Volume 93Author(s): Shan Ma, Wuli Chu, Haoguang Zhang, Song Yan, Yiming Zhong Micro-vortex generator and a jet flow control technique are combined to improve the flow performance of a high-load compressor cascade, and the injection position is further investigated. The calculated results showed that the reverse flow region of baseline grows with the increment of incidence, and an unstable flow phenomenon rapidly occurs near the leading edge when the incidence rises to +7.9°. Meanwhile, two critical points occur on the end-wall, that is considered to be the main reason for occurring stall. When an injection tube is applied to remove the low energy fluid, the total pressure loss appears drastic reduction, and the occurrence of stall shows a slight delay from +7.9° to +8.7° incidence compared with the baseline. With the micro-vortex generator is introduced in the baseline, the occurrence of stall is delayed from +7.9° to +10.0° incidence due to the mixture of low energy fluid and main flow. The combined application shows an excellent effect in delaying the occurrence of stall and reducing the total pressure loss, and the sudden deterioration of the cascade performance occurs at +9.8° incidence. A quantitative analysis of total pressure loss shows that the secondary flow loss and wake loss contribute to the reduction of total pressure loss, the former is reduced by 31.0% and the latter is decreased by 24.1% at the near stall condition. Moreover, the spanwise injection position can affect both the total pressure loss and stall incidence, and the wider operating range is accompanied by the increment of total pressure loss.
  • Effects of bevelled nozzles on standoff shocks in supersonic impinging
    • Abstract: Publication date: Available online 2 September 2019Source: Aerospace Science and TechnologyAuthor(s): H.D. Lim, T.H. New, R. Mariani, Y.D. Cui Moderately under-expanded jets issuing from a circular baseline and two bevelled circular nozzles impinging upon a perpendicular flat plate were experimentally studied. The effects of nozzle-pressure-ratio and separation distance variations on the standoff shock formations were investigated with schlieren visualizations and a visual hull based three-dimensional (3D) shock reconstruction technique to provide deeper insights into their 3D features. Across all flow configurations arising from the different combinations of these parameters, results indicated that the bevelled nozzles are effective in introducing asymmetry to the standoff shock geometries. Depending on the exact flow configuration, standoff shock locations may also undergo significant upstream displacements. In particular, the single-bevelled nozzle produces highly unsteady standoff shocks with asymmetric oscillation amplitudes along both side of the nozzle lip regions. Changes to the standoff shock key characteristics were observed to be sensitive towards the jet shock structures and reflection point modified by the bevelled nozzle exits. In particular, the strength and relative position of the reflection point are identified as the major contributing factors influencing the upstream static pressure distribution of the standoff shock, hence leading to the observed changes in the standoff shock behaviour.
  • Application of remeshed vortex method for the simulation of tip vortex at
           high Reynolds number
    • Abstract: Publication date: October 2019Source: Aerospace Science and Technology, Volume 93Author(s): Youjiang Wang, Moustafa Abdel-Maksoud The spatial and temporal evolution of a single 3D tip vortex is simulated with the remeshed vortex method. The flow conditions correspond to the tip vortex generated by the rectangular NCA0012 wing with an incidence angle of 5 degrees and the Reynolds number of 530,000. The simulation begins at 5 and ends at 35 chordlengths downstream the wing. The vorticity at the inlet is time independent and determined according to measurement data. A source box is disposed in front of the inlet to realise the vortex inlet boundary condition. The obtained tangential velocity and axial velocity profiles show little diffusion and dissipation. The vortex core parameters correlate well with the experimental measurements for all positions (from 5 to 30 chordlengths downstream the wing).
  • Numerical investigation of flow oscillation in a contracting and expanding
           passage subject to vibration
    • Abstract: Publication date: October 2019Source: Aerospace Science and Technology, Volume 93Author(s): Hao Zan, Weixing Zhou, Xuefeng Xiao, Zhixiong Han, Mantang Chen, Yan Li Stable and accurate fuel supply and distribution is a key issue in flow regulation. Because of the small space of a hypersonic vehicle, the contracting and expanding passage is often used to measure the flow rate. However, we experimentally found that the contracting and expanding passage vibration leads to flow oscillation. The flow oscillation significantly affects the fuel flow measurement, and even threatens the safety of the filling process. To understand the mechanism of flow oscillation, a three-dimensional model of a contracting and expanding passage with hydrocarbon fuel was established. After validating the present model with experimental results, a detailed discussion was presented to study the hydrodynamic characteristics of flow oscillation in a contracting and expanding passage subject to vibration. The behaviors of steady and unsteady flow fields under different conditions were investigated. The results show that axial passage vibration leads to flow oscillation because the vibrating passage causes a change in the fluid velocity field. The recirculatory separation and mainstream regions change periodically. The amplitude of the vibrational velocity and the inlet Reynolds number of the fuel have a considerable influence on flow dynamic behavior.
  • Fast preliminary design of low-thrust trajectories for multi-asteroid
    • Abstract: Publication date: October 2019Source: Aerospace Science and Technology, Volume 93Author(s): Zichen Fan, Mingying Huo, Naiming Qi, Ye Xu, Zhiguo Song Multiple-asteroid exploration with low-thrust propulsion requires the design of transfer trajectory and the selection of the visiting sequence, as well as the estimation of the propellant budget. To this end, this paper presents a method to generate transfer trajectories and visiting sequences rapidly using finite Fourier series (FFS) and Monte-Carlo Tree Search (MCTS). The FFS method can generate the transfer trajectory rapidly, which can provide suitable initial approximations, leading to more accurate trajectory optimizations. This study adopts the MCTS algorithm for asteroid sequence selection in the multi-asteroid exploration. By comparing with the traversal algorithm, the greedy algorithm and the tree search algorithm with the trimming strategy, the numerical results show that the MCTS can be used to obtain a quasi-optimal sequence with higher probability and less computation time. Consequently, a method combining FFS and MCTS can rapidly acquire the quasi-optimal visiting sequences with a large probability and the suitable initial trajectory for a multi-asteroid exploration mission. This is very important for the rapid feasibility assessment of hundreds of flight scenarios at the preliminary mission design stage.
  • Path following of the autonomous airship with compensation of unknown wind
           and modeling uncertainties
    • Abstract: Publication date: October 2019Source: Aerospace Science and Technology, Volume 93Author(s): Jie Wang, Xiuyun Meng, Guanghui Wu The motion control of the airship has been an attractive topic in recent years. This paper investigates the problem of the path following of the autonomous airship. Due to the characteristics of the large volume, slow speed, and low-maneuverability, the airship is sensitive to the wind. Based on the design idea of the line-of-sight (LOS) guidance law, a parameter-observer-based LOS (POBLOS) guidance law is proposed. The unknown and time-varying influence from the wind is estimated by the parameter observer and then compensated in the design of the guidance loop. In order to overcome the disturbances of the modeling uncertainties and drift forces induced by the wind, the method of the command filtered backstepping control (CFBC) combined with the disturbance observer is employed in the control loop. The disturbance observer is constructed based on the formulation of the nonlinear disturbance observer (NDO) and improved by adding the status estimation error feedback, which results in a more robust performance when modeling uncertainties and time-varying disturbances exist. The stability analysis shows that the path following error is ultimately bounded. The results of the simulation indicate that the path following system has a good performance and is robust against the influence of the wind and modeling uncertainties.
  • Reduced-order model-based convergence acceleration of reverse mode
           discrete adjoint solvers
    • Abstract: Publication date: October 2019Source: Aerospace Science and Technology, Volume 93Author(s): Andrew L. Kaminsky, Kivanc Ekici This work presents a new technique to reduce the computational cost of sensitivities calculated using a discrete adjoint solver developed via reverse mode automatic differentiation. A fixed-point iterative method is built for the discrete adjoint sensitivity equations by employing the primal time-stepping adjoint approach. The fixed-point sensitivity solution is then accelerated by building a reduced-order model (ROM) that maps the relationship between the sensitivity solution and its corresponding residual. This model is then used to approximate the converged solution, corresponding to a zero residual. While the approximation might not produce the fully converged solution, it typically provides an improved solution that the fixed-point solver can be re-initialized with, which is one of the novel aspects of the present work. After re-initializing the solution, the ROM acceleration technique can be reapplied until the desired convergence criterion is reached. A key feature of the proposed ROM is that it is formed on the fly during a single sensitivity solution. Additionally, its implementation requires only minor modifications to an existing fixed-point iterative solver. The ROM projection technique is evaluated by considering design optimization of horizontal wind turbine blade profiles and cost reductions of 57 to 80% were achieved.
  • Investigation of pulse detonation combustors — Axial turbine system
    • Abstract: Publication date: October 2019Source: Aerospace Science and Technology, Volume 93Author(s): Vijay Anand, Andrew St. George, Ethan Knight, Ephraim Gutmark A circular array of six pulse detonation combustors is integrated to an axial turbine, to extract power from the detonative combustion process that is ignited cyclically in the array of tubes. The effect of pulsing frequency, fill fraction and rotor revolution frequency on the extracted power is ascertained. A method is presented to estimate the component thermal efficiency using experimentally acquired variables. The resulting thermal efficiency values suggest that for a regime of operating conditions, the detonation cycle is more efficient than the Brayton cycle utilizing deflagration. This range of operating conditions appears to depend on the interaction of the different sectors — that are in different operational phases due to cyclic ignition — with the turbine.
  • Flutter improvement of a thin walled wing-engine system by applying
           curvilinear fiber path
    • Abstract: Publication date: October 2019Source: Aerospace Science and Technology, Volume 93Author(s): Touraj Farsadi, Davood Asadi, Hasan Kurtaran In the present study, the aeroelastic behavior of a wing-engine system modeled as composite Thin Walled Beam (TWB) with curvilinear fiber path is investigated. The variable stiffness is acquired by constructing laminates of TWB with curvilinear fibers having prescribed paths. In order to account the effect of chordwise and spanwise locations, mass, and thrust force of engine on the aeroelastic characteristics of TWB, the novel governing equations of motion are obtained using Hamilton's variational principle. The paper aims to exploit desirable fiber paths with improved aeroelastic properties for different wing-engine configuration. Ritz based solution methodology is employed to solve the equations with coupled incompressible unsteady aerodynamic model based on Wagner's function. Numerical simulation results which conform to previously published literatures are presented for validation purposes. Although different curvilinear fiber paths can be introduced to enhance flutter instabilities for each wing-engine configurations, there exists an ideal placement of engine on the wing considering only the engine mass, and the engine mass and thrust force, simultaneously. A comprehensive insight is provided over the effect of parameters such as the lamination fiber path and the effect of engine positions with different mass and thrust values on the flutter speed and frequency.
  • A strongly S-stable low-dissipation and low-dispersion Runge-Kutta scheme
           for convection diffusion systems
    • Abstract: Publication date: October 2019Source: Aerospace Science and Technology, Volume 93Author(s): Yongle Du, Yanchen Liu Since the A-stability and the order of accuracy of time integration methods have been proven insufficient to guarantee the time-accurate simulations of convection diffusion systems, optimized implicit Runge-Kutta schemes with enhanced stability have been developed for stiff problems, and separately those with low-dispersion low-dissipation errors have been proposed for sensitive wave propagation phenomena. However, an implicit Runge-Kutta method is ideally preferred to address both disparate stiffness and various wave propagation characteristics that may be unknown in advance but often co-exist in complex systems, such as turbulent flows with multi-physical phenomena. Therefore, an optimized three-stage second-order diagonally implicit Runge-Kutta scheme with the strong S-stability and low-dispersion low-dissipation errors is derived in this study. Numerical benchmark tests show that overall this newly derived scheme has comprehensively the best performance among the well-known second-order implicit Runge-Kutta schemes. It reaches a second-order accuracy and produces accurate solutions for wave propagation phenomena, but is also stiffly first-order accurate and remains stable with very large time steps for strongly stiff problems.
  • Multi-objective design optimization of blunt body with spike and aerodisk
           in hypersonic flow
    • Abstract: Publication date: October 2019Source: Aerospace Science and Technology, Volume 93Author(s): Jie Huang, Wei-Xing Yao In order to reduce the aerodynamic drag and aerodynamic heating of the hypersonic blunt body, the aerodisk and spike are installed on the blunt body. In this paper, the influences of the aerodisk on the aerodynamic drag and aerodynamic heating of the hypersonic blunt body are studied by numerical method. The results show that the drag coefficient and maximum heat flux of the blunt body for the configuration with the aerodisk are reduced by 34.09% and 51.06% respectively compared with the configuration without the aerodisk, and the two configurations achieve the drag and heat reduction by reconstructing the flow field. On this basis, multi-objective design optimization of the hypersonic blunt body with the aerodisk and spike is performed by the weighting method and NSGA-II method. The results show that increasing the length of the spike and radius of the aerodisk can reduce total heat flux of the blunt body. The drag coefficient decreases with the increase of the length of the spike. However, with the increase of the radius of the aerodisk, the drag coefficient decreases first and then increases. In addition, the drag coefficient and total heat flux of the blunt body of typical Pareto optimal solution are reduced by 35.90% and 46.97% respectively compared with original design, and adjusting the weighting coefficient of the weighting method can obtain the Pareto frontier calculated by NSGA-II method.
  • Multivariable adaptive control based consensus flight control system for
           UAVs formation
    • Abstract: Publication date: October 2019Source: Aerospace Science and Technology, Volume 93Author(s): Ziyang Zhen, Gang Tao, Yue Xu, Ge Song Formation flight contributes to improving the attack, reconnaissance and survival ability of the multiple unmanned aerial vehicles (UAVs). This paper studies a multivariable adaptive control based consensus flight method for UAVs formation. A majority of existing research is focused on the leader-following consensus problem assuming that only the parameters of followers are uncertain. However, they do not consider the leader dynamic uncertainty and the unknown external disturbances. Therefore, this paper addresses the problem of the UAVs consensus flight control with parametric uncertainties and unknown external disturbances for both the leader and follower. A multivariable model reference adaptive control (MRAC) based consensus flight control scheme is designed for UAVs formation, which enables the follower UAV to track the leader UAV. The stability of the multivariable MRAC based consensus flight control system is analyzed. Simulation results show that the proposed adaptive consensus flight control scheme has stronger robustness and adaptivity than the fixed control scheme.
  • Numerical study of the tip clearance flow in miniature gas turbine
    • Abstract: Publication date: October 2019Source: Aerospace Science and Technology, Volume 93Author(s): Junting Xiang, Jorg Schluter, Fei Duan Miniature gas turbines struggle to obtain the efficiencies of the larger counterparts. One of the issues is the larger tip clearance in the compressor due to manufacturing tolerances. In this study, we have performed two numerical parametric studies to determine the influence of tip clearance on gas turbine compressors in the miniaturization process. Firstly, varied tip clearance values are applied when a constant Reynolds number is maintained to investigate the effect of tip clearance value on compressor performance. Thereafter, varied Reynolds numbers are applied when a fixed tip clearance value is used to investigate the effect of miniaturization on compressor performance. Our results quantify the compressor performance deterioration with the existence of tip clearance and show the performance loss amplifies with the increase of tip clearance value as well as with the miniaturization of the compressor. The flow field near the blade tip region has been studied in detail to identify the occurrence of fluid jet and the transition of fluid jet when tip clearance is applied. This study demonstrates the significant impact of tip clearance on compressor performance and quantitatively shows the effect of miniaturization on compressor performance when blade tip clearance exists.
  • A method to select loss correlations for centrifugal compressor
           performance prediction
    • Abstract: Publication date: October 2019Source: Aerospace Science and Technology, Volume 93Author(s): Chaowei Zhang, Xuezhi Dong, Xiyang Liu, Zhigang Sun, Shixun Wu, Qing Gao, Chunqing Tan The study reviews the main loss mechanisms and corresponding loss correlations, which are used to predict centrifugal compressor performance. There are one or more loss correlations in the open literature for the same loss mechanism. The purpose of this review is to identify a reliable loss correlation set for centrifugal compressor performance prediction. According to the inlet tip relative Mach number and specific speed, a new method to select loss correlations is proposed by testing multiple loss correlations. This method is validated by experimental results on eight centrifugal compressors in public and one in-house. The performance as-predicted by the proposed method is in close agreement with the experimental results. Compared with the conventional set, the proposed method is superior.
  • Flow-induced oscillations of circular cylinder in a narrow channel
    • Abstract: Publication date: October 2019Source: Aerospace Science and Technology, Volume 93Author(s): Peter R. Andronov, Yaroslav A. Dynnikov, Galina Ya. Dynnikova, Sergey V. Guvernyuk The problem of transverse auto-oscillations of a cylinder in a flow of a viscous incompressible fluid in a plane channel is considered. The fluid-structure interaction problem is solved in a complete conjugate formulation, in which a continuous medium and a moving rigid body are described as a general dynamical system without splitting into dynamic and hydrodynamic components. The mesh-free method is the most suitable for solving this problem, since the flow region changes not only the form but also the connectivity at the time when the body touches the wall. Another difficulty is connected with the need to construct an adequate model of elastic collisions of bodies in a liquid. Calculations are performed using the fully Lagrangian method of viscous vortex domains (VVD). A model has been developed for the elastic collision of a body with a fixed wall taking into account the influence of a liquid. The mechanism of self-oscillations is revealed. It is shown that the restitution coefficient characterizing the elastic interaction of colliding bodies is an important parameter of the problem under consideration.
  • Dynamic analysis of stiffened bi-directional functionally graded plates
           with porosities under a moving load by dynamic relaxation method with
           kinetic damping
    • Abstract: Publication date: October 2019Source: Aerospace Science and Technology, Volume 93Author(s): Mostafa Esmaeilzadeh, Mehran Kadkhodayan The principal purpose of this study is investigating the dynamic analysis of porous bi-directional functionally graded (FG) plates reinforced by eccentrically outside stiffeners and subjected to a moving load with a constant velocity. The materials are assumed to be graded in two directions and their effective properties are computed by the rule of mixtures. The FG plates are assumed to have both even and uneven distribution of porosities over the plate cross-section. Using appropriate kinematic relations, the displacements of the plate mid-plane are compatible with those of the stiffeners. The governing differential equations of porous bi-directional FG plates are derived through Hamilton's principle based on the first order shear deformation theory (FSDT) and Von Karman relations for large deflections. Moreover, dynamic relaxation method with kinetic damping (K-DR) coupled with Newmark integration technique are used to solve the plate's time-varying nonlinear equations. The effects of some numerical aspect ratios such as volume fraction, boundary conditions, porosity coefficients and distribution patterns and the existence of stiffeners on dynamic behaviors are investigated. The results show that the stiffness of the porous bi-directional FG plates is highly improved with the aid of eccentric stiffeners; hence, better dynamic behaviors are provided.
  • Constrained dynamic compensation with model predictive control for
    • Abstract: Publication date: October 2019Source: Aerospace Science and Technology, Volume 93Author(s): Fabio A. de Almeida This article extends the benefits of model predictive control for the linear dynamic compensators commonly designed for aerospace systems. The traditional control structures often used in output tracking are synthesized, and the proposed model predictive control scheme modifies the plant input and reference to be tracked, creating artificial demands and control corrections to guarantee the state and satisfaction of the output constraint. The novel predictive controller, with a related reference governor, is presented with assured closed-loop stability, convergence and recursive feasibility for attainable piecewise constant setpoints along the system operation. The modification of the feasible demands and control inputs enlarges the domain of operation of the constrained closed-loop system. The elimination of the quadratic programming solvers running during real-time operation is also presented to reduce the computational burden. The simulation results using a fighter trainer model are finally shown, illustrating the benefits of the proposed technique.
  • A fault-tolerant attitude estimation method for quadrotors based on
           analytical redundancy
    • Abstract: Publication date: October 2019Source: Aerospace Science and Technology, Volume 93Author(s): Shichao Liu, Pin Lyu, Jizhou Lai, Cheng Yuan, Bingqing Wang Inertial sensors and velocity measurement sensors are commonly used for attitude estimation of quadrotors. In the previous fault-tolerant attitude estimation methods, the two types of sensors' faults are treated separately. When one sensor is detected, the other sensor is assumed to be always healthy. In this paper, the situation is improved by introducing the rolling & pitching moment model of quadrotors. The rolling & pitching moment model, x-axis gyro, y-axis gyro and velocity measurement sensors are fused, forming an analytical redundancy based fault-tolerant filter. The real flight experiment data are used for validation, showing that the faults of x-axis gyro, y-axis gyro and velocity measurement sensors can all be detected and isolated, and the state estimation accuracy is improved compared with the traditional method when gyros are faulty.
  • Control device effectiveness studies of a 53∘ swept flying wing
           configuration. Experimental, computational, and modeling considerations
    • Abstract: Publication date: October 2019Source: Aerospace Science and Technology, Volume 93Author(s): Patrick Löchert, Kerstin C. Huber, Mehdi Ghoreyshi, Jacob Allen The present investigation covers studies of different control devices on several 53∘ swept flying wing configurations without a vertical tail plane. The wings considered share the same planform but differ in their spanwise profile shapes. The planform, developed under the NATO AVT-251 Task Group, is referred to as the MULDICON (or MULti-DIsciplinary CONfiguration). The objectives of this article are twofold: (1) to design yaw control surfaces for the MULDICON wing using an experimental/computational approach and (2) to develop aerodynamic models that rapidly and accurately predict the effectiveness of control surfaces over a wide range of flight conditions. The yaw control surface design (position and size) should provide sufficient yaw moment with almost no contribution in roll and pitch moment. To identify such concepts, a number of preliminary experiments on a generic flying wing configuration have been conducted. Two promising concepts from wind tunnel tests were then numerically examined for being implemented in the MULDICON baseline wing. Concepts with spoilers and a split flap were specifically considered. For medium to high angles of attack, the flow topology of the baseline wing is dominated by a vortical flow field on the upper outer wing. This leads to interactions between vortex and the control device which influences the flow and the attitude of the control device on the upper wing side. Furthermore, this work considers developing aerodynamic models for predicting stability derivatives of several MULDICON designs over a wide range of flight conditions. Aerodynamic loads models are only developed for normal force and pitch moment coefficients, however, the developed approach can easily being extended to include lateral aerodynamic coefficients as well. Quasi-steady models of this work are power series expansions of traditional linear aerodynamic models to capture nonlinear effects. Additionally, the models can estimate static and dynamic stability derivatives and the control surface powers.
  • Energy management strategy design and station-keeping strategy
           optimization for high altitude balloon with altitude control system
    • Abstract: Publication date: October 2019Source: Aerospace Science and Technology, Volume 93Author(s): Huafei Du, Mingyun Lv, Lanchuan Zhang, Weiyu Zhu, Yifei Wu, Jun Li The capacity of station-keeping and long endurance of renewable energy system have great impact on the application of the high altitude balloon. In this paper, the station-keeping mechanism based on an altitude control system is proposed to ride the natural wind field with diverse directions and speeds at different altitudes and then retain the balloon within the desired district. Combined with the energy harvesting/consuming model and the dynamic model, the energy management strategy is designed to extend the energy endurance of the balloon. The station-keeping strategy (the pumping/venting states of the air ballonet) based on the energy management strategy is optimized by the genetic algorithm. The optimization result shows that the air ballonet needs to be pumped 6 times and vented 5 times in one day. By pumping and venting the air ballonet, the balloon can be restricted within the district with a radius of 30 km and the maximum horizontal distance between the center of the district and the location of the balloon is about 28 km. In addition, the payload and the altitude control system can be powered by the PV array and the lithium battery sustainably. These results can assist in the design and initial flight tests of the serviceable high altitude balloon.
  • A comprehensive investigation of acoustic power level in a moderate or
           intense low oxygen dilution in a jet-in-hot-coflow under various working
    • Abstract: Publication date: October 2019Source: Aerospace Science and Technology, Volume 93Author(s): Seyed Mahmood Mousavi, Reza Kamali, Freshteh Sotoudeh, Reza Pourabidi, Nader Karimi, In-Seuck Jeung Reducing the level of noise emission is an important requirement in modern propulsion and power generation. This requires gaining a deeper understanding of the underlying physics and identifying the key parameters dominating noise generation in modern combustion technologies. Thus, this paper investigates the effect of several working parameters on moderate or intense low oxygen dilution (MILD) combustion noise. A finite volume solver, GRI-Mech 2.11 with 247 reactions and 49 species, k-ε RNG turbulence model as well as the EDC model are used to develop a computational model of the reactive flow, while the volume fraction and time scale constants are set to 3 and 1, respectively. After validating the numerical method by comparison with the experimental data, MILD combustion is simulated under various conditions to study the noise emission. The results show that modifying the inlet conditions such as changes in species mass fractions, inlet temperature, and inlet Reynolds number alter the acoustic power level through variations in heat release.
  • Optimized diagonally implicit Runge-Kutta schemes for time-dependent wave
           propagation problems
    • Abstract: Publication date: October 2019Source: Aerospace Science and Technology, Volume 93Author(s): Yongle Du, Yanchen Liu, John A. Ekaterinaris High-order, low-dissipation and low-dispersion time-integration schemes are critical for efficient large-eddy simulations of sensitive time-dependent wave propagation problems, such as aeroacoustic phenomena. Based on the relative error analysis of the model equation dy/dt=(μ+iλ)y, a comprehensive analysis on the optimal low-dispersion low-dissipation diagonally implicit Runge-Kutta schemes is presented. Similar studies on optimal implicit Runge-Kutta schemes were proposed in existing publications to minimize the integrated error of the amplification factor in a pre-defined range of λΔt on the imaginary axis. In contrast, this study introduces the concept of “regions with acceptable amplification and phase-shift” of the numerical solutions in the complex (μ+iλ)Δt plane. As a result, the optimization aims to maximize the radii of the regions with acceptable dispersion and dissipation errors. When applied to the 2- and 3-stage diagonally implicit Runge-Kutta schemes, optimal A-stable or not-A-stable schemes with the low-dispersion and low-dissipation property and the 2nd- upto 4th-order accuracy are derived with equally good or better performances as compared to existing implicit Runge-Kutta schemes. Numerical experiments consistently demonstrate that the proposed criteria provide better indicators for the dispersion and dissipation errors of the time-integration schemes. Furthermore, the optimal schemes achieve the design order of accuracy with reasonably large time steps.
  • Effects of steady and pulsed discharge arcs on shock wave control in Mach
           2.5 flow
    • Abstract: Publication date: October 2019Source: Aerospace Science and Technology, Volume 93Author(s): Fan Liu, Hong Yan, Wangjie Zhan, Yunpeng Xue The effects of the steady and pulsed arc discharge on the oblique shock wave control are explored through an experimental study in a supersonic wind tunnel with the maximum design Mach number of 2.5. The oblique shock is formed by a compression ramp with an angle of 7 degree mounted on the floor of the wind tunnel. Four electrodes are placed upstream of the compression ramp and spaced equally in the spanwise direction to generate the discharge arcs. The steady discharge is generated between the electrodes and downstream ramp corner, forming streamwise arcs. The arc length can be modulated between 10 mm and 40 mm by moving the electrodes in the streamwise direction. With a High Frequency Switch (HFS), the pulsed discharge is achievable with a frequency range from 5 kHz to 50 kHz. The pulsed discharge is generated between two adjacent electrodes, forming so-called transversal arcs, which are blown downstream by the main flow. Results show that the steady discharge arcs act like a uniformly distributed conductor. With an increase of the arc length, the arc power increases, and the weakening effect on the shock is enhanced. For the pulsed discharge arcs, the shorter arcs are observed with higher discharge frequency, which implies a lower arc power. Overall, the weakening effect of the steady arcs on the shock is more effective with shock strength reduced by 4%, compared to the pulsed ones with only 0.35% reduction.
  • Appointed-time prescribed performance attitude tracking control via double
           performance functions
    • Abstract: Publication date: October 2019Source: Aerospace Science and Technology, Volume 93Author(s): Zeyang Yin, Afzal Suleman, Jianjun Luo, Caisheng Wei This work investigates the attitude tracking control problem of spacecraft under strong external disturbances and parameter uncertainties. A novel appointed-time stable control scheme is proposed with guaranteed transient and steady-state performance. First, an appointed-time reachable performance function (ARPF) is presented, and its reach time can be arbitrarily selected by the users. Then, a double-ARPFs strategy is introduced, that is, by imposing two ARPFs on the attitude and the system output, respectively, all system states will be appointed-time stable. Furthermore, a robust controller with implementable structure is proposed to guarantee the performance functions under strong external disturbances and parameter uncertainties. And the attitude tracking errors as well as the angular velocity errors are proved to be appointed-time stable. Last, three groups of simulations are organized to verify the effectiveness, robustness and appointed-time stability of the proposed control scheme.
  • Surrogate-based aerodynamic optimisation of compact nacelle aero-engines
    • Abstract: Publication date: October 2019Source: Aerospace Science and Technology, Volume 93Author(s): Fernando Tejero, David G. MacManus, Christopher Sheaf Genetic algorithms are a powerful optimisation technique for the design of complex engineering systems. Although computing power continuously grows, methods purely based on expensive numerical simulations are still challenging for the optimisation of aerodynamic components at an early stage of the design process. For this reason, response surface models are typically employed as a driver of the genetic algorithm. This reduces considerably the total overhead computational cost but at the expense of an inherent prediction uncertainty. Aero-engine nacelle design is a complex multi-objective optimisation problem due to the nonlinearity of transonic flow aerodynamics. This research develops a new framework, that combines surrogate modelling and numerical simulations, for the multi-objective optimisation of aero-engine nacelles. The method initially employs numerical simulations to guide the genetic algorithm through generations and uses a combination of higher fidelity results along with evolving surrogate models to identify a set of optimum designs. This new approach has been applied to the multi-objective optimisation of civil aero-engines which are representative of future turbofan configurations. Compared to the conventional CFD in-the-loop optimisation method, the proposed algorithm successfully identified the same set of optimum nacelle designs at a 25% reduction in the computational cost. Within the context of preliminary design, the method meets the typical 5% acceptability criterion with a 65% reduction in computational cost.
  • Morphing aircraft control based on switched nonlinear systems and adaptive
           dynamic programming
    • Abstract: Publication date: October 2019Source: Aerospace Science and Technology, Volume 93Author(s): Qing Wang, Ligang Gong, Chaoyang Dong, Kewei Zhong This paper investigates the control problem of a morphing aircraft with variable sweep wings based on switched nonlinear systems and adaptive dynamic programming (ADP). The longitudinal altitude motion of the morphing aircraft is first modeled as switched nonlinear systems in lower triangular form. Then, the designed controller is comprised of the basic part and supplementary part. For the basic part, the backstepping technique is applied and a modified dynamic surface is introduced to overcome the ‘explosion of complexity’ problem. Disturbance observers inspired from the idea of extended state observer are designed to obtain the estimations of the internal uncertainties and external disturbances. The common virtual control laws of the backstepping method are developed by the disturbance observers and radial basis function neural networks. On the other hand, for the supplementary part, an ADP approach with the name of action-dependent heuristic dynamic programming is used to further decrease the altitude tracking error, which generates an additional control input by observing the differences between the actual and desired values in the backstepping design. Finally, comparative simulations are conducted to demonstrate the improved control performance of the proposed approach.
  • Parametric design of non-axisymmetric separate-jet aero-engine exhaust
    • Abstract: Publication date: October 2019Source: Aerospace Science and Technology, Volume 93Author(s): John J. Otter, Robert Christie, Ioannis Goulos, David G. MacManus, Nicholas Grech Future civil air vehicles are likely to feature propulsion systems which are more closely integrated with the airframe. For a podded underwing configuration, this close coupling is expected to require non-axisymmetric design capabilities for the aero-engine exhaust system. This work presents the development of a novel parametric representation of non-axisymmetric aero-engine exhaust system geometries based on Intuitive Class Shape Transformation (iCST) curves. An exhaust design method was established and aerodynamic analyses of a range of non-axisymmetric configurations was demonstrated. At typical flight conditions, the introduction of non-axisymmetric separate-jet nozzles was shown to increase the engine net propulsive force by 0.12% relative to an axisymmetric nozzle.
  • Experimental and numerical study of a cryogenic valve using liquid
           nitrogen and water
    • Abstract: Publication date: October 2019Source: Aerospace Science and Technology, Volume 93Author(s): J. Pinho, L. Peveroni, M.R. Vetrano, J.-M. Buchlin, J. Steelant, M. Strengnart In the present study, a cryogenic valve used in launch vehicle liquid propulsion systems is experimentally and numerically characterized. Two independent measurement campaigns are performed with liquid nitrogen and water as working fluids. Two equivalent facilities have been designed and manufactured to perform the cryogenic and the water tests. The characteristic relationship between the volumetric flow rate and pressure drop across the test valve is obtained for both the fluids. As far as cryogenic tests are concerned, temperature measurements at the test valve inlet and outlet are presented as well as visualizations of the flow upstream the test section. The experimental results show that the test valve flow coefficient is independent of the working fluid provided single phase flow conditions. To further validate this result a numerical study is conducted using the commercial code CFD-ACE+. A good agreement between numerical and experimental results is found. Furthermore, test cases in the semi-critical and critical flow conditions are simulated using the so-called full cavitation model. The computed liquid recovery factor is shown to be also independent from the working fluid nature.
  • Numerical identification of separation bubble in an ultra-high-lift
           turbine cascade using URANS simulation and proper orthogonal decomposition
    • Abstract: Publication date: October 2019Source: Aerospace Science and Technology, Volume 93Author(s): Seyed Morteza Sajadmanesh, Mohammad Mojaddam, Arman Mohseni, Ali Nikparto The flow-field inside a gas turbine engine, especially in the low-pressure turbine, is very complicated as it is normally accompanied by unsteady flow structures, strong and rapidly changing pressure gradients, intermittent transition of boundary layer, and flow separation and reattachment, especially during off-design performance. In this article, flow separation and reattachment on the suction side of an ultra-high-lift low-pressure turbine blade is studied and characterized using 3D Unsteady Reynolds-Averaged Navier-Stokes (URANS) equations. For turbulence modeling, transitional-SST method (γ-Reθ) is adopted. The simulations are performed at the exit Reynolds numbers of 200,000 and 60,000, and at a constant isentropic exit Mach number of 0.4. The shape and extent of the separation bubble are primarily dependent on large vortical structures due to the Kelvin-Helmholtz instability and spanwise vortex tube shedding. Therefore, a better prediction of these phenomena could result in a more realistic separation bubble identification and consequently more accurate profile loss assessment. In order to better capture the transitional flow characteristics, which are not often readily available from conventional computational fluid dynamics simulations, the method of Proper Orthogonal Decomposition (POD) is used in this study. Non-coherent structures in the main flow, such as separation bubble, are investigated and studied. The POD modes of pressure-field are analyzed to clarify the generation of spanwise vortex tubes after separation point. In the higher Reynolds number, low-energy small-scale structures in the separation zone and downstream of the trailing edge are observed from the POD analysis. In the lower Reynolds number, high-energy large-scale structures shed from the separated shear layer are identified, which are responsible for increasing turbulent kinetic energy as well as increasing profile losses. This study also shows that the combination of URANS and POD can successfully be used to identify the separation bubble.
  • Flow field and injector heat characteristics of hybrid rocket motor with
           annular-gap injector
    • Abstract: Publication date: October 2019Source: Aerospace Science and Technology, Volume 93Author(s): Chengen Li, Guobiao Cai, Pengfei Wang, Hui Tian A throttleable annular-gap axial injector, which could automatically adjust the injection gap area, is proposed to decrease injection pressure drop variation amplitude during wide range thrust regulation process in hybrid rocket motor. With the annular-gap injector adopted, liquid oxidizer is unhomogeneously injected into combustion chamber through a narrow annular injection gap. In conventional hybrid rocket motor, the oxidizer is usually homogeneously injected through full-inlet injection method. The injection area covers the whole fuel grain port. Difference of the injection methods greatly influences the oxidizer flow characteristics. Consequently, combustion and heat transfer characteristics of the motor are significantly changed. This paper is aimed to analyze the two-phase combustion flow field and coupled injector heat transfer characteristics of a lab-scale hybrid rocket motor with annular-gap injector through two-dimensional axisymmetric steady numerical simulations. The motor adopts 98% hydrogen peroxide and polyethylene as the propellants. Numerical analysis reveals that position of the injection gap influences the regression rate distribution in the first half of solid fuel grain but has little effect on that in the second half. The regression rate is relatively high when the injection gap is close to the fuel inner surface. In addition, the flowing liquid hydrogen peroxide in the injector could cool the chamber head. Sharp turns that produces vortex in oxidizer flow channel decreases the cooling effect and increases the overheating risk. Smooth bend could improve the injector heat transfer characteristics and eliminate the risk.
  • Externally blown elevon applied for the longitudinal control of blended
           wing body transport with podded engines
    • Abstract: Publication date: October 2019Source: Aerospace Science and Technology, Volume 93Author(s): Zhenqing Xin, Zhenli Chen, Wenting Gu, Minghui Zhang, Binqian Zhang Limited longitudinal control authority due to the absence of a horizontal tail and short lever arm is a critical challenge for blended-wing-body (BWB) aircrafts, especially during take-off and landing conditions. To address this issue, an extended external blown elevon is promoted as an alternative. The control capability of this new type of elevon is studied using computational fluid dynamics (CFD) methods. A parameter study is performed on the extended position. The control authority of the blown elevon is more efficient than that of the conventional plain elevon under different flight conditions because of the lengthened lever arm, flow-separation suppression and jet-blowing effects. The effects of a one-engine inoperative (OEI) condition on the control capability is also evaluated. The results indicate that the blown elevon can alleviate the directional moment caused by OEI. It is concluded that the external blown elevon can be a potential option for longitudinal control in practical BWB design.
  • Aerodynamic design assessment and comparisons of the MULDICON UCAV concept
    • Abstract: Publication date: October 2019Source: Aerospace Science and Technology, Volume 93Author(s): Raj Nangia, Mehdi Ghoreyshi, Michel P.C. van Rooij, Russell M. Cummings This article describes the aerodynamic design and assessment of a UCAV wing developed under the NATO-STO AVT-251 Multi-disciplinary Design Task Group. The wing design rationale and process is described, as well as how the idealized wing design produced by the design/optimization process is turned into a realizable aircraft geometry. Both the baseline and designed wings are analyzed in detail using several Navier-Stokes computational fluid dynamics flow solvers. Analysis at low speed and transonic conditions, both in the pitch plane and at sideslip conditions, is presented. Conclusions about the design are drawn and further improvements are suggested.
  • Numerical investigation of the operating process of the liquid hydrogen
           tank under gaseous hydrogen pressurization
    • Abstract: Publication date: October 2019Source: Aerospace Science and Technology, Volume 93Author(s): Jiachao Li, Guozhu Liang, Pingping Zhu, Xi Wang In order to accurately predict the whole operating process of a liquid hydrogen tank under gaseous hydrogen pressurization, a 2-D axial symmetry Volume-of-Fluid (VOF) based numerical simulation method is established. Phase change and turbulence models are included in the numerical simulation. The variations of physical parameters such as the ullage mass, temperature and pressure, are carefully analyzed. The different effects are given based on simulations with and without phase change, and the comparison between feedback pressurization and open pressurization is also given. Compared with the NASA's experiment under the feedback pressurization, the simulation results show that the deviation of pressurant gas masses consumption is 11.0% during the whole operating process. The deviation of the total ullage mass is −0.8%, 1.4% and 7.6% for the ramp period, the hold period and the expulsion period, respectively. The deviation of phase change mass is 7.5% and −21.5% for the ramp period and the expulsion period, respectively. The simulation results also reach an agreement with the experiment on the energy absorption proportions and demonstrate that most of the energy addition from the external environment and the pressurizing gas is absorbed by the tank wall. The liquid gains the least energy during the expulsion period. Temperature stratification appears along the axial direction in the surface liquid region and the ullage region, and the bulk liquid is in a subcooled state. The location of phase change mainly appears near the vapor-liquid interface, where the net condensation appears during the ramp period and the hold period, while the net vaporization appears during the expulsion period. The phase change increases the amplitude of temperature oscillation. The open pressurization has an ullage pressure peak and an average ullage temperature peak, which lead to large impacts on the tank structure, but the control of the inlet mass flow rate is easy to implement. The feedback pressurization could maintain a steady ullage pressure, but more pressurant gas masses are consumed, and the control of inlet mass flow rate becomes more complicated. The simulation results can be used as references for design optimization of the pressurization systems of cryogenic liquid launch vehicles in order to save pressurant gas masses and decrease the ullage pressure peak which could reduce the tank wall thickness and enhance the carrying capacity of liquid launch vehicles.
  • A three-dimensional aircraft ice accretion model based on the numerical
           solution of the unsteady Stefan problem
    • Abstract: Publication date: October 2019Source: Aerospace Science and Technology, Volume 93Author(s): Tong Liu, Kun Qu, Jinsheng Cai, Shucheng Pan This paper develops a three-dimensional aircraft ice accretion model based on the numerical solution of the unsteady Stefan problem. In this model, for both rime and glaze ice situation, the heat conduction within the ice layer is considered as an unsteady process and the evolution of temperature distribution is numerically solved by Variable Space Grid (VSG) method. Unlike the Myers' state-of-the-art model and its extended model which introduces a posteriori correction factor α by assuming a linear temperature profile, current model does not need to tune the parameter α to match the experimental results. A hierarchical overset grid strategy is applied to ease the grid generation and shorten the workload on constructing complex icing configurations. Only the grid around the iced surface needs to be regenerated at each ice accretion step and complex flow characteristics near the ice horn can be captured accurately with such high-quality grid. First, simulations on NACA 0012 airfoil are conducted to validate current ice accretion model. Under rime and glaze ice condition, the predicted ice shapes agree well with experimental results. The predicted temperature evolution at iced surface reveals that ice accretion is a complex process influenced by heat and mass transfer. Then the computations on GLC-305 swept wing are performed to study the ice accretion process on the three-dimensional surface. The results show a good agreement with the experimental data. At the leading edge, the overall convective heat transfer coefficient and droplet collection efficiency increase along the direction from the wing root to the wing tip. This leads to very distinct ice horns at different wing sections.
  • Solar sailing technology challenges
    • Abstract: Publication date: October 2019Source: Aerospace Science and Technology, Volume 93Author(s): David A. Spencer, Les Johnson, Alexandra C. Long Solar sailing technology has been demonstrated in the space environment over the past decade, in Earth orbit and on an interplanetary trajectory. These technology demonstration missions, along with a forty-year history of conceptual studies and laboratory development, have provided a foundation for a new era of missions where solar sailing provides the necessary propulsion to achieve space science and infrastructure goals. Numerous challenges remain on the path to flagship-class missions utilizing solar sails. This paper provides a survey of the current state of the art in solar sailing technology, including a taxonomy of solar sail design. A summary of solar sailing missions is provided, along with description of the larger-scale ground test programs. A set of representative next-generation solar sailing mission concepts is then presented, to establish driving requirements for future applications. To meet the objectives for these future missions, sail areas must increase by a factor of 50–500 relative to the largest solar sail flown to date. Sail loading, sailcraft areal density, characteristic acceleration and lightness number must improve by one to two orders of magnitude. Technology advancements required to meet the future solar sailing performance needs are described, providing a technology roadmap for solar sailing capability.
  • Inlet bent torsional pipe effect on the performance and stability of a
           centrifugal compressor with volute
    • Abstract: Publication date: October 2019Source: Aerospace Science and Technology, Volume 93Author(s): Hanzhi Zhang, Ce Yang, Changmao Yang, Hang Zhang, Leilei Wang, Jiang Chen Inlet bent torsional (BT) pipes are applied to complex turbocharger systems and industrial centrifugal compressors owing to space and weight constraints. To determine the interaction mechanism between the inlet BT pipe and outlet volute, and its effect on the centrifugal compressor performance and stability, three compressor models (model with clean inlet (M0 model) and models with distorted inlets induced by BT pipes (M1 and M2 models)) were chosen to conduct a performance experiment and static pressure measurement of the casing wall. The results show that the M1 and M2 cause a 6.4% increment and 22% reduction in the stable operating range, respectively. At the near-choke and peak efficiency points, both M1 and M2 can change the casing static pressure distribution evidently and thus influence the compressor pressure ratio and efficiency significantly. At the near-stall point, the high static pressure at the inlet region induced by the BT pipe of M1 can weaken the pressure peak strip induced by the volute. Therefore, both the considerable increase in circumferential pressure uniformity and the positive pre-whirl distortion of M1 are responsible for the delay of compressor stall. Finally, the interaction relationship of the inlet/outlet distortions in the entire operating range was sketched.
  • Numerical study on inlet angle of guide vane in recess vaned casing
    • Abstract: Publication date: October 2019Source: Aerospace Science and Technology, Volume 93Author(s): Xiangyi Chen, Wuli Chu, Haoguang Zhang, Xiangjun Li This paper studies the effect of inlet angle of guide vane in the recess vaned casing treatment (RVCT). The smooth casing fan and the casing treated fans with different vane angles have been numerically simulated respectively. The result shows that the application of RVCT brings about noticeable improvements in the fan's stall margin and the modification of guide vane inlet angle leads to various effects. RVCT functions as a bridge that contributes to the extra flow circulation at the blade tip region and mitigates the flow blockage. The inlet angle of guide vane determines the amount as well as the potential in terms of accommodating fluid passing through the RVCT. Based on the evolution of flow topology, it is hypothesized that the separation on the guide vane surface is a prerequisite to the stall of the fan. The scenario that the attachment line totally blocks the vane inlet passage is a criterion of the stall. The deflection of inlet angle from positive to negative enables to enhance the stall margin by delaying the formation of the prerequisite and the criterion of the stall.
  • The effect of angle of attack on the aeroacoustic environment within the
           weapons bay of a generic UCAV
    • Abstract: Publication date: October 2019Source: Aerospace Science and Technology, Volume 93Author(s): David Bacci, Alistair J. Saddington, Derek Bray Cavity flow studies are generally concerned with observing the effect of geometry changes whilst maintaining a fixed zero angle of attack. Cavities employed as weapons bays will, however, experience a range of angles of attack. This paper presents the first known results showing the effect of flight angle of attack on the aeroacoustic characteristics of an internal weapons bay installed in an uninhabited combat air vehicle (UCAV). The UCAV geometry consisted of a Boeing M219-type cavity in a Boeing UCAV1303 airframe. Numerical simulation was conducted using a full-scale detached eddy simulation model and representative transonic flight conditions. As well as the reference case of zero degrees, data for angles of attack of 3.0, 4.5 and 6.0 degrees were analysed. Experimental data was used to validate the reference computational model, which agreed with the overall fluctuating sound pressure level (OAFPL) to within the experimental uncertainty of 4 dB. Data from the computational model was post-processed with frequency-domain and time-frequency-domain techniques showing that the flow structure within the weapons bay was altered significantly by the angle of attack changes, affecting the mean pressure distribution, frequency spectra and resonant modes. Overall, increasing the angle of attack from 0.0 to 3.0 degrees produced an increment in the acoustic load whilst a further increase tended to affect the resonance mechanism and thereby reduce the coherence and the temporal footprints of the resonant modes.
  • Adaptive fast nonsingular terminal sliding mode control for attitude
           tracking of flexible spacecraft with rotating appendage
    • Abstract: Publication date: October 2019Source: Aerospace Science and Technology, Volume 93Author(s): Yue Miao, Inseok Hwang, Ming Liu, Feng Wang This paper investigates the attitude tracking problem for the flexible spacecraft composed of a rigid platform with a rotating appendage. Two robust attitude controllers, each matched to the platform and the appendage, are constructed simultaneously. First, the relative kinematic models are presented, and the dynamic model describing the translation, rotation, and vibration of the spacecraft is proposed using the Lagrange Method. Then, a novel fast nonsingular terminal sliding mode surface is designed, and control strategies are proposed by integrating the aforementioned sliding manifold with the adaptive methodology. The adaptive update laws are employed to estimate the boundaries of various uncertainties, which do not require a priori disturbance information. Finally, illustrative numerical simulations are conducted to demonstrate the performance of the proposed control technique in terms of fast convergence and robustness compared to existing control schemes.
  • Terminal sliding mode control based impact time and angle constrained
    • Abstract: Publication date: October 2019Source: Aerospace Science and Technology, Volume 93Author(s): Zhiwei Hou, Ye Yang, Lei Liu, Yongji Wang In this paper, an impact time and angle constrained guidance (ITACG) is proposed based on the nonsingular terminal sliding mode control (NTSMC) theory. First, the guidance law is derived for stationary targets. The proposed ITACG consists of two parts. One part is designed for a missile to intercept a target from the desired impact angle and the other part aims to achieve the desired impact time. Two different terminal sliding mode surfaces are designed so that the missile can satisfy the impact time and angle constraints simultaneously. Corresponding to the designed sliding mode surfaces, two different Lyapunov candidate functions are proposed and analyzed, and stability conditions are obtained. Then, we extend the proposed guidance for constant acceleration targets. The time-to-go estimation method is modified based on the conception of predicted-intercept-point (PIP) with the constant acceleration targets and stability conditions are also revised. Compared with traditional sliding mode control (SMC) based ITACG, the proposed guidance law is a direct online method. It does not need to design the line-of-sight angle curve off-line, nor does it need to switch between impact time constrained guidance and impact angle constrained guidance. In the end, numerical simulation results show that the proposed impact time and angle constrained guidance law has a good performance even though the missile has a constant acceleration. Salvo attack of multi-missiles against one target is shown in the simulation part by applying the proposed guidance law.
  • Adaptive phase compensator for vibration suppression of structures with
           parameter perturbation
    • Abstract: Publication date: October 2019Source: Aerospace Science and Technology, Volume 93Author(s): Wenchao Niu, Bin Li The constant all-pass filter has been widely used for phase compensation of control systems. However, it is not applicable to compensating phase deviation caused by additional filters to enhance the signal quality in vibration control systems with parameter perturbation. To overcome this challenge, an adaptive phase compensator (APC) is developed by transforming the constant parameters of an all-pass filter into frequency-dependent parameters. In addition, the sources of phase deviations in the control system are analyzed to design the APC, including the additional filters, non-collocated actuator/sensor configuration, and hardware hysteresis. The phase deviations are determined through simulation and experiment. Polynomial fitting is implemented to obtain the APC parameters. To verify the feasibility of the proposed APC, numerical and experimental efforts are undertaken for buffeting suppression of the vertical tail, which is a typical structure with parameter perturbation. Both results demonstrate that the stability and robustness of control system adopting APC are strengthened compared to that of the control system using a constant phase compensator. Moreover, a control system that adopts APC can also effectively reduce the vibration response for structures with parameter perturbation under harmonic and random excitations. This performance improvement indicates that the proposed APC provides more effective compensation performance for the phase deviation of control systems with time-varying perturbations.
  • Effect of voltage and droplet size on electrical ignition characteristics
           of ADN-based liquid propellant droplet
    • Abstract: Publication date: October 2019Source: Aerospace Science and Technology, Volume 93Author(s): Lei Li, Guo-Xiu Li, Hong-Meng Li, Zhao-Pu Yao To remove the dependence of ammonium dinitramide (ADN)-based liquid space engine on the catalyst, a new thermal ignition method was developed for ADN-based liquid propellant resistant to ignition. Experiments of ADN-based liquid monopropellant droplets were carried out in a constant-volume chamber. Ignition delay time and combustion duration were defined to study the electrical ignition characteristics of single droplet. The results show that the ADN-based liquid propellant droplets could be ignited using the current thermal effect of droplets after electrification at appropriate voltage. Droplets were accompanied by bubble formation and gases escape from the droplets during evaporation, decomposition, and combustion. With the increase in voltages, the ignition delay time and combustion duration of ADN-based liquid propellant droplets decreased with the increase in voltage overall when the initial volumes of droplets were 4.189 μL. The ignition delay time of droplets increased with the increase in the initial volume of droplet. The combustion duration of droplets showed fluctuant variation trend with the increase in the initial volume of droplet. When the initial volume of droplet increased to a certain extent, combustion self-sustainment could not be achieved at the later stage of combustion.
  • Aerodynamic layout optimization design of a barrel-launched UAV wing
           considering control capability of multiple control surfaces
    • Abstract: Publication date: October 2019Source: Aerospace Science and Technology, Volume 93Author(s): Zhe Zhu, Hongwu Guo, Jianjun Ma Owing to the size limitations of launchers, foldable main wings and foldable vertical tails are usually used in the design of barrel-launched unmanned aerial vehicles (UAVs), and the folding mechanisms are complex and bulky. Further, the coupling effect between aerodynamic performance and control capability is not typically considered in the aerodynamic layout design of traditional drum-launched UAVs; thus, the overall performance is often improved by serial design and multi-turns optimization, which result in low design efficiency. Based on the available research on barrel-launched UAVs, this paper presents an optimal design scheme for non-vertical tail aerodynamic layout considering the control capability of multiple control surfaces. A flying-wing configuration, with two groups of ailerons and one group of all moving tips, is employed in the design, and its control capability and aerodynamic performance are quantitatively described in terms of the volume of attainable moment subset and lift–drag ratio. The focus of the current design is shifted from improvement of single aerodynamic performance to optimization of overall performance. To overcome the computational complexity observed in direct optimization, the number of design variables is reduced by identifying the key test factors. The concept of a surrogate model matrix comprising cell surrogate models is proposed to effectively fit the aerodynamic characteristics and control capability of the wing. A multi-objective optimization based on the surrogate model matrix is carried out using the neighborhood cultivation genetic algorithm, which effectively improves the optimization efficiency. The entire process is implemented on an ISIGHT platform for automatic optimization design of the wing. At two design points selected on the pareto frontier, the results show that the lift–drag ratio is increased by 28.36%, reachable torque space is increased by 64.29% in state 1, lift–drag ratio is increased by 23.11%, and reachable torque space is increased by 164.29% in state 2. These results provide a reference for the aerodynamic layout design of barrel-launched UAVs for practical control requirements.
  • Robust integrated orbit and attitude estimation using geophysical data
    • Abstract: Publication date: October 2019Source: Aerospace Science and Technology, Volume 93Author(s): Maryam Kiani Geophysical information such as the Earth geomagnetic field and gravity gradient (GG) data can provide a basis for autonomous concurrent orbit and attitude estimation (COAE) of satellites in low earth orbits (LEO), as magnetometers and gravity gradiometer measurements are in general functions of time, position as well as the vehicle's orientation. While gradiometer has recently been investigated just for orbit estimation (OE), the current study is focused on COAE via only utility of the GG data. To this aim, observability conditions are analyzed, where the sensitivity of the proposed COAE approach with respect to various system and roto-translational elements is also examined. Considering the nonlinear nature of the COAE problem, a modified robust unscented Kalman filter is adopted for state estimation that is enhanced by innovation-based fading factors to be robust against the modeling errors. Subsequently, a centralized fusion of GG and three-axis magnetometer (TAM) measurement data is employed to improve the accuracy and reliability of the proposed approach for space navigation. The results obtained from numerical test cases confirm feasibility and effectiveness of the proposed technique for complete COAE of space navigation in LEO satellites.
  • Attitude recovery scheme of magnetically controlled satellite with
           constant thrust
    • Abstract: Publication date: October 2019Source: Aerospace Science and Technology, Volume 93Author(s): Xiwang Xia, Chaoyong Li, Chongbin Guo, Dong Li, Zhao Li When the constant thrust is acting directly on the bias momentum satellite, one component of the angular velocity vector would increment gradually to a larger one or even to infinite, which beyond the controllable region corresponding to B-dot algorithm. And in the following rate damping control process, the corresponding control effort would lead to attitude instability. In this study, under the constant thrust, the satellite's attitude evolution laws is researched and, under the control effort corresponding to B-dot damping algorithm, the satellite's tumbling mechanism is analyzed. Two critical angular velocities, respectively corresponding to the watershed value and the saddle point value, are determined according to the details of attitude control systems, including attitude control cycle and the corresponding time sequence. Theoretical analysis results show that, when the angular velocity is larger than the first critical angular velocity corresponding to the controllable region of B-dot algorithm, the B-dot damping algorithm would fail to de-tumble the satellite. On the contrary, the corresponding control effort would drive the angular velocity to another critical angular velocity, which is the saddle point for B-dot algorithm. Finally, a series of simulation examples are presented to verify the proposed conclusions.
  • A solution for the attitude determination of three-vehicle heterogeneous
    • Abstract: Publication date: October 2019Source: Aerospace Science and Technology, Volume 93Author(s): Pedro Cruz, Pedro Batista This paper proposes a solution to a new attitude determination problem for a three-vehicle formation, where there are restrictions on the detection of other vehicles. Vision-based sensors are considered, which measure line-of-sight (LOS) vectors between different vehicles and inertial vectors that can vary according to the vehicle. There is a constraint on the LOS vectors which are only measured in relation to a chief vehicle. Moreover, each vehicle can measure only one inertial vector. The solution for the different attitude relations is devised geometrically, making use of a multi-stage process. First, two candidates for the relative attitude of each branch are determined. Then, these relative attitude candidates are used to compute inertial attitude candidates for the chief vehicle. The comparison between these results disambiguates the problem, which in general has a unique solution. There are some degenerate solutions, which are also determined. Finally, simulations are carried out considering noise in the sensors. The results are coherent with state-of-the-art approaches to similar problems.
  • Substructuring verification of a rear fuselage mounted twin-engine
    • Abstract: Publication date: October 2019Source: Aerospace Science and Technology, Volume 93Author(s): Braden T. Warwick, Il Yong Kim, Chris K. Mechefske Dynamic substructuring allows for the reduction of large complex structures into substructures to increase computational efficiency and to isolate the local dynamic behaviors of concern. However, errors such as truncation, continuity and rigid body mode errors still limit the applicability of this method experimentally. Additionally, the feasibility of implementing substructuring techniques on finite element models of multi-component fuselage structures has yet to be shown in the literature. The objective of this paper is twofold: first to introduce a feasible substructuring methodology that mitigates the experimental and multi-component limitations with current methods; and secondly, to investigate the modal properties and applicability of substructuring analysis on a rear fuselage mounted twin-engine aircraft. This configuration is not well understood in the literature despite having been shown to have increased interior cabin noise and vibration levels. Experimental validation of the computational model was first performed. A substructuring analysis of the validated computational model produced natural frequencies of the local and global modes that agreed within 6.47% on average, and pseudo-orthogonality terms greater than 0.89 for all modes considered. This methodology proved to be useful for generating an accurate representation of local modes within a global structure for the aircraft configuration studied. This will allow for future work to more thoroughly investigate the local modes using innovative design methods with confidence that the local modes will correlate with the global modes.
  • Cooperative detection based on the adaptive interacting multiple
           model-information filtering algorithm
    • Abstract: Publication date: October 2019Source: Aerospace Science and Technology, Volume 93Author(s): Shuai Zhang, Yang Guo, Zhaoxing Lu, Shicheng Wang, Zhiguo Liu This paper develops a multiaircraft cooperative detection scheme for effectively improving the accuracy when detecting the highly maneuvering target, where the cooperative estimate and observation trajectory planning of multiaircraft are addressed simultaneously. Based on the error compression ratio of the Markov matrix, an adaptive IMM algorithm is proposed for single detection, the convergence characteristic of which is theoretically proven meanwhile. Then, an adaptive IMM-information filtering algorithm is derived by embedding information filtering into the adaptive IMM algorithm for the fusion of the detection information. Next, the coordinated planning of observation trajectories is achieved for multiaircrafts by relying on receding horizon optimization (RHO) and the detection information. The cooperative detection and trajectory planning scheme is illustrated by the simulation results.
  • Vibration control of aero two-blade propeller with input and output
           constraints based on PDE model
    • Abstract: Publication date: October 2019Source: Aerospace Science and Technology, Volume 93Author(s): Xueyan Xing, Jinkun Liu In this study, the two-blade aero propeller is regarded as a distributed parameter system whose dynamics is fully described by partial differential equations (PDEs), taking into account all the system modal information. Then a control approach relying upon the backstepping technology is designed to steer the vibration of the propeller system to a small region even subject to external disturbances and modeling uncertainties. With the help of the proposed control, both input and output constraints of the system can be guaranteed via barrier Lyapunov functions (BLFs) and hyperbolic tangent functions. Utilizing the Lyapunov's direct method, the system stability is verified and the effectiveness of the proposed control scheme is illustrated by numerical simulations.
  • Inertial vector measurements based attitude synchronization control for
           multiple spacecraft formation
    • Abstract: Publication date: October 2019Source: Aerospace Science and Technology, Volume 93Author(s): Zhong Zheng, Mouquan Shen for multiple spacecraft with actuator and angular velocity constraints using vector measurements directly. First, the novel vector measurements-based cooperative attitude synchronization control scheme is presented by using barrier Lyapunov function. Especially, the proposed control algorithm is proven to be robust to time delays in the directed communication links, and the actuator and angular velocity constraints are satisfied simultaneously. Second, further analysis of the initial condition is performed to preclude the undesired equilibrium. It indicates that at least 2 orthogonal inertial vectors are required. Third, the fully distributed attitude synchronization control strategy with adaptive gains is developed. Moreover, the control algorithm is implemented without any global information of the graph. This affords a new approach to optimize the control performance. Finally, simulation results demonstrate the effectiveness and performance of the proposed control schemes.
  • Comparative performance analysis of solid oxide fuel cell turbine-less jet
           engines for electric propulsion airplanes: Application of alternative fuel
    • Abstract: Publication date: October 2019Source: Aerospace Science and Technology, Volume 93Author(s): Zhixing Ji, Jiang Qin, Kunlin Cheng, Fafu Guo, Silong Zhang, Peng Dong Technology progress makes it possible to power airplanes by fuel-flexible and highly efficient solid oxide fuel cells (SOFCs). A turbine-less jet engine is achieved by integrating SOFCs with a compressor and a nozzle. It can be applied to electric propulsion airplanes. The compressor is powered by SOFCs instead of turbines. However, it is uncertain which type of fuel is suitable to the SOFC jet engine because the propulsion system is sensitive to performance and weight. In this paper, preliminary thermodynamic cycle analysis shows that fuel types and pressure ratios are key parameters for the engine. Then, five configurations of the SOFC jet engines when fed by hydrogen, methane, methanol, decane and propane are proposed and comparative performance analysis is accomplished. Main conclusions are as follows: (1) The specific thrust of the engine is nearly regardless of fuel types, which is about 970–1000 N/(kg s−1). The thermal efficiency and the specific impulse of the engine both decrease when methane is replaced by propane or decane. (2) The differences of performance resulting from fuel types increase with the increase of pressure ratios. (3) At the fuel-air equivalent of 1, the specific impulse of the engine fed by methanol is low to 1209 s.
  • Flame propagation and flashback characteristics in a kerosene fueled
           supersonic combustor equipped with strut/wall combined fuel injectors
    • Abstract: Publication date: October 2019Source: Aerospace Science and Technology, Volume 93Author(s): Junlong Zhang, Juntao Chang, Zi'ao Wang, Lin Gao, Wen Bao Flame propagation characteristics in a liquid kerosene fueled supersonic combustor were observed in this paper. The strut/wall combined fuel injection scheme was used to achieve the multi-staged fuel injection, and the thin strut was also adopted as the flame holder. A series of experiments and numerical simulations were carried out in the condition of Ma=2.8. In order to represent the flame characteristics during the flame propagation process, high-speed photography and pressure measurement were taken to record the data. Based on the basic date, the variation of flame area and Mach number distributions during the experimental process were analyzed, and the flowing field characteristics were also discussed according to the numerical results. Results indicated that the core flame would diffuse to the primary flow with the increasing of equivalence ratio, enlarging the flame width. The flame flashback phenomenon of both core flame and wall flame, induced by the flame propagation, was detected and analyzed. Then, the interaction between combustion and flowing in the supersonic combustor was investigated, based on which, the mechanism of flame flashback phenomenon was also revealed. Numerical results showed that the generation of recirculation and low-speed region in boundary layer was the inducing mechanism of flame flashback. With the investigations in this paper, a depth understanding of the flame propagation process in supersonic airflow was achieved, based on which, a further optimization of combustion performance could be conducted.
  • Nonlinear control of unmanned aerial vehicles with cable suspended
    • Abstract: Publication date: October 2019Source: Aerospace Science and Technology, Volume 93Author(s): Ameya R. Godbole, Kamesh Subbarao This paper focuses on the mathematical modeling and control of an unmanned aerial system (UAS) with a payload suspended using a cable. The motion of the payload induces disturbances on the aerial platform and needs to be mitigated for stable operation. The solution to this control problem is presented through the implementation of a passivity based controller, and an extended state observer based active disturbance rejection controller. The implementation of the passivity based controller requires the knowledge of higher time derivatives of the payload oscillations. Assuming only the swing angles of the payload with respect to a UAS are measured, these states (primarily the angular velocity) are estimated using a continuous-discrete Kalman Filter. Alternately since, the payload cable swing angle is difficult to measure, an active disturbance rejection controller is designed and implemented wherein the disturbance induced in the system due to the motion of the payload is estimated using the extended state observer. A comparison between the passivity based controller and the extended state observer based active disturbance rejection controller is performed using a high fidelity numerical simulation.
  • Robust adaptive backstepping fast terminal sliding mode controller for
           uncertain quadrotor UAV
    • Abstract: Publication date: Available online 26 July 2019Source: Aerospace Science and TechnologyAuthor(s): Moussa Labbadi, Mohamed Cherkaoui The problem of controlling the quadrotor orientation and position is considered in the presence of parametric uncertainties and external disturbances. Previous works generally assume that the flight controller parameters are constants. In reality, these parameters depend on the desired trajectory. In this article, a complete mathematical model of a quadrotor UAV is presented based on the Euler-Newton formulation. A robust nonlinear fast control structured for the quadrotor position and attitude trajectory tracking is designed. The position loop generates the actual thrust to control the altitude of the quadrotor and provides the desired pitch and roll angles to the attitude loop, which allow the control of the quadrotor center of gravity in the horizontal plane. The attitude loop generates the rolling, pitching and yawing torques that easily allow the insurance of the quadrotors stability. The outer loop (position loop) uses the robust adaptive backstepping (AB) control to get the desired Euler-angles and the control laws. The inner loop (attitude loop) employs a new controller based on a combination of backstepping technique and fast terminal sliding mode control (AB-ABFTSMC) to command the yaw angle and the tilting angles. In order to estimate the proposed controller parameters of the position and the upper bounds of the uncertainties and disturbances of the attitude, online adaptive rules are proposed. Furthermore, the Lyapunov analysis is used to warranty the stability of the quadrotor UAV system and to ensure the robustness of the controllers against variation. Finally, different simulations were performed in the MATLAB environment to show the efficiency of the suggested controller. The sovereignty of the proposed controller is highlighted by comparing its performance with various approaches such as classical sliding mode control, integral backstepping and second order sliding mode controls.
  • Experimental study of mode transition characteristics of a cavity-based
           scramjet combustor during acceleration
    • Abstract: Publication date: Available online 26 July 2019Source: Aerospace Science and TechnologyAuthor(s): Yu Meng, Hongbin Gu, Jingheng Zhuang, Wenming Sun, Zhanbiao Gao, Huan Lian, Lianjie Yue, Xinyu Chang Experiments were performed in a direct-connect supersonic combustion test facility to simulate scramjet combustor acceleration in high altitude. The combustor inlet flow Mach number increased from 2.40 to 2.94, and the flowrate and total temperature simultaneously changed with Ma. The combustor has two cavities and fuel jets. The fuel used is room-temperature liquid kerosene RP3. Mode transition is seen to occur in the first cavity during acceleration, and pressure fluctuations occur in the transitions from both the ram mode to the transition mode and from the transition mode to the scram mode, indicating that the mode transition process is unstable. When the mode transition occurs upstream, the downstream ram mode has the effect of eliminating instability; therefore, the engine's overall thrust performance is stable. When the downstream is also in the transition mode, the thrust fluctuates, indicating that the mode transition is an unstable process.
  • Studies of thermal deformation and shape control of a space planar phased
           array antenna
    • Abstract: Publication date: Available online 23 July 2019Source: Aerospace Science and TechnologyAuthor(s): Lu Guang-Yu, Zhou Ji-Yang, Cai Guo-Ping, Fang Guang-Qiang, Lv Liang-Liang, Peng Fu-Jun Large planar phased array antenna in space will inevitably occur deformation due to the space thermal environment, which could severely influence the pointing accuracy of the antenna. This paper focuses on the studies of thermal load and shape adjustment method of a planar phased array antenna structure. The finite element method is adopted both to develop the structural model and to obtain the steady temperature field of the antenna structure, based on which a novel shape adjustment method that utilizes cables as actuators is proposed. This method transforms the control problem into the optimization problem of the actuators placement. The optimal placements of actuators and the corresponding control forces are determined by using discrete PSO method and quadratic optimization method. A 100 meter scale antenna structure is presented as example to validate the effectiveness of the proposed method numerically. Simulation results indicate that this method could successfully control the thermally induced deformation of the structure and maintain antenna's shape accuracy.
  • Study on nonlinear vibration of simplified solid rocket motor model
    • Abstract: Publication date: Available online 23 July 2019Source: Aerospace Science and TechnologyAuthor(s): Dongxu Zhang, Shu Cheng, Fangfang Xu, Yumeng Hu, Hong Li The unstable combustion of the large aspect-ration solid rocket in the end phase flight is of great interest and importance to the solid rocket design.The large aspect-ratio solid rocket motor (SRM) tends to produce unstable combustion in flight. As shown in published literatures, it is believed that the main reason for this phenomenon is the internal flow field. However, in terms of the intense pressure oscillation, the contribution of the motor casing should not be ignored. In this paper, the nonlinear natural frequencies of different motor shells in flight are mainly concerned. The SRM model in the working state is simplified to a double-layer time-varying axially moving free-free beam. The governing equation of nonlinear lateral vibration is derived according to the Hamilton principle. The Galerkin method and multi-scale method are adopted to solve the governing equations. Firstly, to verify the accuracy of the presented method, the validating experiments base on the finite difference method are carried out. According to the comparison, the results obtained through the Galerkin method coupled with multi-scale method coincides well with the results obtained by the finite difference method as well as the results in the public literatures. The influence of structural change and axial velocity on the nonlinear coefficients and natural frequencies of the SRM shell is investigated by numerical experiments. The results show that the natural frequency of the SRM shell change during the whole flight and the increase of aspect ratio has a significant influence on the natural frequency. Under the initial excitation, the nonlinear natural frequency will change abruptly. Under some particular excitations, it is possible that the resonance coupling between the shell vibration and internal flow field occurs and it contributes to the amplification of the unstable combustion further.
  • A numerical study of rotating detonation wave with different numbers of
           fuel holes
    • Abstract: Publication date: Available online 19 July 2019Source: Aerospace Science and TechnologyAuthor(s): Qingyang Meng, Ningbo Zhao, Hongtao Zheng, Jialong Yang, Zhiming Li, Fuquan Deng In this paper, three-dimensional numerical investigations are conducted to characterize the slot-hole rotating detonation combustor under various numbers of fuel holes (40, 60, 90 and 120). The effects of fuel holes number on detonation wave are discussed according to three typical stages which are non-reacting stage, detonation formation stage and detonation stable stage respectively. Numerical results show that the number of fuel holes has a significant impact on the formation and propagation characteristics of rotating detonation wave. On detonation formation stage, the general formation processes of detonation wave are similar under different holes numbers, which mainly includes the collision of transmit pressure waves and reactant re-initiation phenomenon. Within the scope of this study, as the increase of holes number, shorter axial length is taken to close to the global equivalence ratio on non-reacting stage. For 90-hole case, it takes the shortest time to establish self-sustained detonation wave and has the highest propagating stability. Additionally, dual-wave mode is obtained when the holes number is 90, while single-wave mode occurs when the holes numbers are 40, 60 and 120. The propagating direction of RDW is reversal when the holes number is 60 and 120, the reason is related to the collision spot of pressure wave and distributions of equivalence ratio. Besides, the establishment of new detonation wave is dependent on the intensity of accumulated pressure wave and the distribution of equivalence ratio.
  • Sub-optimal cooperative collision avoidance maneuvers of multiple active
           spacecraft via discrete-time generating functions
    • Abstract: Publication date: Available online 17 July 2019Source: Aerospace Science and TechnologyAuthor(s): Kwangwon Lee, Hyeongjun Park, Chandeok Park, Sang-Young Park This study presents real-time sub-optimal control for cooperative collision-free transfers of multiple active (actuated) spacecraft in proximity operations. The constrained optimal control problem for collision-free transfers of multiple active spacecraft is decentralized and approximated as an unconstrained optimal control problem for single active spacecraft to mitigate the complexity and difficulty. The new penalty function is proposed by considering relative velocities for cooperative maneuvers between multiple active spacecraft, and is integrated with the quadratic cost function for optimal tracking by continuous-thrust control instead of the inequality constraints for avoiding collision. Then, the infinite-horizon control law applicable to each of multiple active spacecraft is obtained as an algebraic function of the states of both reference solutions and obstacles by employing discrete-time generating functions. Unlike conventional methods based on shooting, the proposed approach does not require repetitive process and initial guesses regardless of the number of active spacecraft. Illustrative examples demonstrate the effectiveness of the proposed approach with the new penalty function especially in simultaneous collision avoidance maneuvers of multiple active spacecraft.
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