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Aerospace Science and Technology
Journal Prestige (SJR): 0.796
Citation Impact (citeScore): 3
Number of Followers: 350  
 
  Hybrid Journal Hybrid journal (It can contain Open Access articles)
ISSN (Print) 1270-9638
Published by Elsevier Homepage  [3157 journals]
  • Output feedback continuous terminal sliding mode guidance law for
           missile-target interception with autopilot dynamics
    • Abstract: Publication date: Available online 17 January 2019Source: Aerospace Science and TechnologyAuthor(s): Zhenhua Zhao, Chuntao Li, Jun Yang, Shihua Li This paper proposes an output feedback continuous terminal sliding mode guidance law (CTSMGL) for three-dimensional missile-target interception with consideration of second-order autopilot dynamics. The proposed output feedback CTSMGL not only guarantees the rates of the line-of-sight angles converge to zero in finite time but also ensures the continuity of control action. Besides, only the relative range and velocity information of guidance system has been utilized in its design process. Simulations of a practical interception process under three types of maneuvering targets are carried out and the simulation results illustrate the effectiveness of the proposed guidance method.
       
  • A fast finite-time convergent guidance law with nonlinear disturbance
           observer for unmanned aerial vehicles collision avoidance
    • Abstract: Publication date: Available online 17 January 2019Source: Aerospace Science and TechnologyAuthor(s): Ning Zhang, Wendong Gai, Maiying Zhong, Jing Zhang In this paper, a fast finite-time convergent guidance law with nonlinear disturbance observer is proposed for the problem of unmanned aerial vehicle collision avoidance. First, a linear time-varying collision avoidance model based on a collision cone is established. Then a fast finite-time convergent guidance law and a nonlinear disturbance observer are designed to ensure the safety of UAV collision avoidance. In addition, the stability of a nonlinear collision avoidance system is proved by the finite-time convergence stability theory, and the stability conditions are used to design the guidance coefficients. The simulation results conclusively demonstrate that this method can achieve collision avoidance in the presence of an unknown acceleration of obstacle. Moreover, the performance of collision avoidance using this method is greater than the collision avoidance method based on normal finite time convergent guidance.
       
  • Buckling and post-buckling analysis of cracked stiffened panels via an
           X-Ritz method
    • Abstract: Publication date: Available online 17 January 2019Source: Aerospace Science and TechnologyAuthor(s): V. Gulizzi, V. Oliveri, A. Milazzo A multi-domain eXtended Ritz formulation, called X-Ritz, for the analysis of buckling and post-buckling of stiffened panels with cracks is presented. The theoretical framework is based on the First-order Shear Deformation Theory and accounts for von Kármán's geometric nonlinearities. The structure is modeled as assembly of plate elements. Penalty techniques are used to fulfill the continuity condition along the edges of contiguous elements and to satisfy essential boundary conditions requirements. The use of an extended set of approximating functions allows to model through-the-thickness cracks and to capture the crack opening and tip singular fields as well as the structural behavior within each single domain. Numerical results for buckling and post-buckling of cracked stiffened panels are compared with finite elements simulations and literature solutions, showing the accuracy and potential of the proposed approach.
       
  • High-power mode control for triaxial gas turbines with variable power
           turbine guide vanes
    • Abstract: Publication date: Available online 17 January 2019Source: Aerospace Science and TechnologyAuthor(s): Tao Wang, Zhao Yin, Chun-qing Tan, Yong-sheng Tian, Qing Gao, Hua-liang Zhang Factors such as the atmospheric environment or component performance influence various matching relationships as well as maximum output power in triaxial gas turbines with variable power turbine guide vanes. This paper proposes a control strategy called High Power Mode (HPM) which solves the matching problem, adjusts fuel flow, adjusts the guide vane to adapt the gas turbine to unpredictable factors automatically, and enhances its maximum output power without exceeding high- and low-pressure rotor rotational speeds or combustor outlet total temperature limits.We first built a component-level gas turbine simulation model to explore the control strategy and the influence of guide vane angle on gas turbine working for the variable geometry triaxial gas turbine. We conducted a series of gas turbine rig tests to validate the simulation model. We then determined the guide vane angle effect law on the gas turbine performance based on the simulations as well as experimental data. HPM, including guide vane angle and fuel flow rate control strategies, was established according to the acquired guide vane angle influence law. We wrote an HPM control program in MATLAB/SIMULINK to find that the program indeed optimizes components and parameters (high- and low-pressure rotor rotational speeds and combustor outlet total temperature) while matching them effectively within unpredictable environments. The maximum gas turbine output power can be increased with the program by more than 4% in certain atmospheric environments.
       
  • Modified near-wakes of axisymmetric cylinders with slanted base
    • Abstract: Publication date: Available online 17 January 2019Source: Aerospace Science and TechnologyAuthor(s): D.S. Bulathsinghala, Z. Wang, I. Gursul Near-wakes of axisymmetric cylinders with slanted base were investigated in wind tunnel experiments for an upsweep angle of 28°. Effects of splitter-plate, cavity, and flaps on the afterbody vortices and separated flow were studied by means of surface pressure and Particle Image Velocimetry measurements. The splitter plate causes more diffused afterbody vortices due to the turbulence ingestion from the separation region. When the slanted base is replaced with a deep cavity, there is weaker roll-up of vorticity due to the lack of streamwise flow and a solid surface. Varying the splitter plate length in the range tested did not have significant influence on the flowfield, apart from affecting the strength of the splitter-plate vortex. With the splitter plate and cavity, unsteadiness is dominated by the flow separation region, as opposed to the afterbody vortices in the baseline case. A pair of vertical flaps attached to the side-edges of the splitter plate can reduce the unsteadiness at the measurement plane immediately downstream of the splitter plate.
       
  • Investigating the effect of engine speed and flight altitude on the
           performance of throttle body injection (TBI) system of a two-stroke
           air-powered engine
    • Abstract: Publication date: Available online 16 January 2019Source: Aerospace Science and TechnologyAuthor(s): Ali Hassantabar, Ahmad Najjaran, Mahmoud Farzaneh-Gord Two-stroke engines are one of the most commonly used engines for unmanned aerial vehicle (UAV) propulsion, which has benefits such as high power, low fuel consumption. Before, most of the two-stroke engines used in UAVs were the carburettor, but nowadays, carburettor engines are not suitable for UAVs because of stringent standards such as high flight endurance, tolerance of extreme environmental conditions. Also, the solution to meet these expectations is the use of electronic fuel injection systems. In this study, the performance of the fuel injection system in the throttle body injection (TBI) of the two-stroke air-powered engine with the capacity of 30 cc used in the UAV is investigated. Therefore, the performance of the TBI system and the engine is simulated using the lotus engine simulation (LES) software. The obtained results (mass flow rate of air and fuel, the pulse width of the injector) are used as the initial and boundary conditions of numerical simulation of the TBI system using computational fluid dynamics (CFD) to simulate and check the airflow and fuel injection in the TBI system. Design features pressure drop in the throttle body, turbulence, speed-pressure distribution in flow, fuel injection and droplet distribution, and mixing them with air, were analysed in the TBI system. First, the air flow is simulated in a Samorfeld rotary burner; where the k-w SST turbulence model is found suitable for airflow modelling in the throttle body. In the next step, to validate the fuel injection modelling, a non-reactive fuel spray was simulated in Sandia burner, and it is observed that the results of fuel injection modelling using the Kelvin-Helmholtz and Riley-Taylor failure model were well-precise to verify the hypothesis and models used. Then the performance of the TBI system was simulated in various operating conditions (1000, 3000, 6000 and 9000 RPM, flying heights of 0 to 20,000 feet). The results of the study of the effect of engine speed on the pressure drop and turbulence intensity of the flow turbulence showed that with the increase in engine speed from 1000 to 9000, the intensity of the turbulence of the flow in the throttle body was increased. Also, with the increase in engine speed, the power of the recycling area was increased. According to the results, the air pressure in the throttle body will decrease with increasing the height. In the study of the effect of the engine speed on the fuel injection characteristic, it was observed that in the 30 degrees throttle and the high speed due to the presence of strong recycling area, the turbulence intensity was high, and the deviation of the droplets from the direction of injection and the dispersion of droplets was also higher.
       
  • Computational study of axisymmetric divergent bypass dual throat nozzle
    • Abstract: Publication date: Available online 16 January 2019Source: Aerospace Science and TechnologyAuthor(s): Yangsheng Wang, Jinglei Xu, Shuai Huang, Yongchen Lin, Jingjing Jiang A bypass dual throat nozzle (BDTN) is a fluidic thrust-vectoring nozzle (FTVN) that is based on the conventional dual throat nozzle (DTN). The BDTN provides high thrust vectoring (TV) efficiency with low thrust loss and does not consume additional secondary flow. A fixed-geometry axisymmetric divergent BDTN (ADBDTN) that has flow adaptive capability and can provide high TV efficiency in the pitch and yaw directions without substantially changing the working state of the engine is investigated in this study. In addition, the TV mechanism and the effects of the expansion ratio, bypass width, and rounding radii at the nozzle throat and cavity bottom on the nozzle performance under the vectored state are presented in detail. Results show that the TV mechanism of the ADBDTN is similar to that of the DTN, which deflects the primary flow through the separation zone inside the recessed cavity. Larger expansion ratios and rounding radii at the nozzle throat contribute to the large pitch thrust-vector angle and high discharge coefficient. Decreasing the expansion ratio and increasing the rounding radius at the nozzle throat can improve the thrust coefficient. As the bypass width increases, the thrust and discharge coefficients increase and the pitch thrust-vector angle reaches its maximum. The effect of the rounding radius at the cavity bottom on the nozzle performance is relatively small. For the configuration with flow adaptive capability, the largest pitch thrust-vector angle is 25.37∘ and the highest thrust and discharge coefficients are 0.957 and 0.983, respectively.
       
  • Effect of acicular vortex generators on the aerodynamic features of a
           slender delta wing
    • Abstract: Publication date: Available online 16 January 2019Source: Aerospace Science and TechnologyAuthor(s): Omid Nematollahi, Mahdi Nili-Ahmadabadi, Hyunduk Seo, Kyung Chun Kim This study investigates the flow structures and aerodynamic performance of a slender delta wing that has a sweep angle of 65 degrees and is equipped with acicular vortex generators. The results were compared with those of a smooth wing. Particle image velocimetry (PIV) was used to visualize the flow at six different sections of the delta wing in both the chord-wise and span-wise directions at an angle of attack (AOA) of 30 degrees and a Reynolds number of Re=2.4×105. The effect of the acicular vortex generators on the vortex diameters, locations of vortex breakdown, linear convolution integral patterns, circulation, and separation were studied. The Q-criterion was used for vortex boundaries for the span-wise sections. The results show the acicular vortex generators have positive effects on the aerodynamic performance of the wing. Furthermore, the acicular wing delays the vortex breakdown in comparison with a smooth wing and increases the flow momentum near the upper wing surface and behind the wing.
       
  • Control of compressor tip leakage flow using plasma actuation
    • Abstract: Publication date: Available online 16 January 2019Source: Aerospace Science and TechnologyAuthor(s): Zhang Haideng, Wu Yun, Li Yinghong, Yu Xianjun, Liu Baojie Control of the compressor tip leakage flow using plasma actuation is studied in this paper based on a compressor cascade model. Three types of plasma actuation layouts are designed: the axial plasma actuation which induces body force along the axial flow direction, the normal plasma actuation which induces body force normal to the interface between the tip approaching and leakage flow, and the stagger angle plasma actuation which induces body force normal to the blade chordwise direction. Influences of plasma actuation on the compressor cascade tip total pressure loss and blockage as well as the flow structures are first studied numerically and the corresponding flow control law is obtained. The numerical results are then verified through experiments. At last, the flow control mechanism behind the suppression of compressor tip leakage flow using plasma actuation is investigated. It is found that the key element of plasma actuation in suppressing the compressor tip leakage flow lays in improving the axial momentum of compressor cascade tip flow. The optimal flow control effect in suppressing tip leakage flow is achieved with the axial plasma actuation that is the most effective in improving the axial momentum of compressor tip flow.
       
  • Nonlinear dynamic analysis near resonance of a beam-ring structure for
           modelling circular truss antenna under time-dependent thermal excitation
    • Abstract: Publication date: Available online 16 January 2019Source: Aerospace Science and TechnologyAuthor(s): W. Zhang, R.Q. Wu, K. Behdinan The analysis on the nonlinear dynamics of the circular truss antenna is an important problem for its design and control since its large flexible structure. This paper investigates the nonlinear dynamic behaviors of a beam-ring structure modelling the circular truss antenna subjected to the periodic thermal excitation. Based on describing the displacements and the nonlinear strains of the beam-ring structure, the kinetic energy and potential energy are calculated for the system. Using Hamilton's principle, the nonlinear partial differential governing equations of motion and the boundary conditions are derived for the beam-ring structure. Applying Galerkin's approach, the nonlinear partial governing differential equations of motion for the beam-ring structure are truncated into a two-degree-of-freedom system of the ordinary differential equation including the quadratic nonlinearities. The four-dimensional averaged equation is obtained for the case of the primary resonance and 1:2 internal resonance by using the method of multiple scales. From the averaged equation, numerical simulations are presented to investigate the frequency-response curves and the influences of the thermal excitation on the nonlinear dynamic responses of the beam-ring structure. The numerical results demonstrate that there exist the complex nonlinear dynamic phenomena of the beam-ring structure. The finding phenomena of this paper are helpful for controlling the nonlinear vibrations of the antenna.
       
  • Discharge voltage characteristic in ablative pulsed plasma thrusters
    • Abstract: Publication date: Available online 15 January 2019Source: Aerospace Science and TechnologyAuthor(s): Guorui Sun, Zhiwen Wu, Hang Li, Linghan Zeng Ablative Pulsed Plasma Thruster (APPT), as the earliest electric thruster, has been applied in engineering applications over 50 years. The voltage characteristic, as the most direct parameter, is mostly measured in all APPT experiments to measure the performances. However, almost all researchers applied the capacitor voltages instead of the electrode voltages to estimate the performances (i.e., propellant utilization efficiency, plasma flow between electrodes). Due to the real-time variation of the plasma impedance, there were some deviations in these estimations. In this study, voltage characteristics of the accelerated electrodes were analyzed experimentally and theoretically: in the single discharge process, the trends of phase delay and amplitude difference between electrode voltage and capacitor voltage are related with the variations of the resistance and inductances of the plasma bulk; as the APPT discharge voltage increases, the phase delay and amplitude difference between electrode voltage and capacitor voltage increase in the whole trend, the estimated deviations become larger. This study provides a reference for the numerical model analysis in ablative pulsed plasma thrusters.
       
  • Numerical simulation on thermal and mass diffusion of MMH–NTO
           bipropellant thruster plume flow using global kinetic reaction model
    • Abstract: Publication date: Available online 3 December 2018Source: Aerospace Science and TechnologyAuthor(s): Kyun Ho Lee A space propulsion system has a crucial role to perform several mission operations of a spacecraft successfully in an orbit. When a thruster is fired, the exhaust plume gas can have effects on the performance of a spacecraft because the expanded plume gas molecules directly collide with the spacecraft surfaces in the vacuum environment. Thus, the present study investigated more realistic plume flow behaviors using a global kinetic reaction model for an actual combustion process of a fuel and an oxidizer. To achieve this, the 4-step global combustion model of monomethylhydrazine and nitrogen tetroxide was incorporated for the first time in the plume flow analysis to reflect a more practical firing condition of a bipropellant thruster. Then, thermal and mass diffusion predictions of the plume flow were compared with the chemical equilibrium condition to examine the distinct differences between the proposed and conventional approach. For efficient numerical calculations, the Navier–Stokes equations and the DSMC method were combined to deal with a continuum flowfield inside the thruster and a rarefied plume gas flow outside the nozzle together. With the present analysis results, major differences in the thermal and mass behaviors of the plume gases were compared between the two reaction models, and their influences are also discussed.
       
  • Mechanism/structure/aerodynamic multidisciplinary optimization of flexible
           high-lift devices for transport aircraft
    • Abstract: Publication date: Available online 24 October 2018Source: Aerospace Science and TechnologyAuthor(s): Yun Tian, Jianchong Quan, Peiqing Liu, Doudou Li, Chuihuan Kong Mission adaptive variable camber wing in both chord-wise and span-wise directions that can improve the aerodynamic performance during takeoff, landing and cruising flight, will be the state-of-the-art high-lift system for next generation airliners. Based on NASA TrapWing model released on the 1st AIAA CFD High-lift Prediction Workshop, a smart high-lift system with “Flexible Droop Nose & Single Slotted Hinge Flap combined with spoiler deflection & Flexible Trailing Edge Flap” is proposed in this paper. The Flexible Droop Nose is actuated by kinematic chains mechanism, the Single Slotted Hinge Flap is actuated by simple hinge mechanism and the Flexible Trailing Edge Flap is actuated by link/track mechanism.A mechanism/structure/aerodynamic multidisciplinary optimization platform based on iSIGHT software is constructed for this smart high-lift system. This platform consists of stress analysis, high-lift configuration generation, high-lift configuration structure grid generation, computational fluid dynamics and optimization algorithm modules. The optimal takeoff and landing configurations with comprehensive performance of mechanism, aerodynamic and structure is then obtained after multidisciplinary optimization. Finally, the CFD results show that the aerodynamics performance of this smart high-lift system is more effective than the original NASA TrapWing model.
       
  • Coverage-based three-dimensional cooperative guidance strategy against
           highly maneuvering target
    • Abstract: Publication date: Available online 25 August 2018Source: Aerospace Science and TechnologyAuthor(s): Wenshan Su, Kebo Li, Lei Chen A coverage-based cooperative guidance strategy is proposed for multiple missiles to cooperatively intercept a highly maneuvering target. The scenario of interest is that several missiles separate with each other to cooperatively intercept the target at the beginning of terminal guidance phase, and both the interceptors and target are assumed to have limited maneuverability. The guidance goal is formulated to cooperatively cover the target evasion region by coordinating the reachable sets of different interceptors. To achieve this goal, firstly, the concept of virtual aiming points is introduced to bias the reachable set of true proportional navigation guidance law (TPN) as expected. Then, to maximally cover the target maneuvering range, we transform the cooperation problem into an optimization problem to find a set of proper virtual aiming points for different missiles. The main advantage of the proposed guidance strategy is that it enables multiple missiles without maneuverability superiority to intercept a highly maneuvering target. Numerical simulations clearly show the effectiveness of the proposed cooperative strategy.
       
  • The effect of mismatching between combustor and HP vanes on the
           aerodynamics and heat load in a 1-1/2 stages turbine
    • Abstract: Publication date: Available online 11 January 2019Source: Aerospace Science and TechnologyAuthor(s): Yifei Li, Xinrong Su, Xin Yuan In modern aero-engine and gas turbine, the hot streak and residual swirl caused by the combustor have strong impact on both aero and thermal performances of the high-temperature turbine. Due to different design philosophies of combustors and turbine, one combustor usually matches two or more vanes and the resultant mismatching effect is numerically investigated with unsteady simulation where one nonuniform core faces two Nozzle Guide Vane (NGV) blades. Results indicate that the nonuniformity clocking position dominates the averaged-temperature, while the swirl orientation dominates the radial distribution. The mismatching effect strongly impacts both time-averaged and instantaneous performances, in that neighboring blades of NGV have averaged-temperature difference of 10.5 K, and it also halves the base frequency and amplifies unsteady variation in the rotor passage. By analyzing both aero and thermal results from a matrix of four cases, the LE-pos configuration has the best performance. The complex transportation of the nonuniformity in the LE-pos case is detailed and the unsteady interactions between the inlet nonuniformity and the complex flow topology are analyzed. The residual swirl carries the hot streak fluctuating radially in rotor passages and redistributes the temperature on rotor surface. Results from this work highlight the necessity of considering the mismatching effect in gas turbine design optimization.
       
  • Three-dimensional terminal angle constrained robust guidance law with
           autopilot lag consideration
    • Abstract: Publication date: Available online 11 January 2019Source: Aerospace Science and TechnologyAuthor(s): Yi Ji, Defu Lin, Wei Wang, Shaoyong Hu, Pei Pei Based on sliding mode control theory and back stepping design technique, this paper proposes a robust guidance law considering autopilot dynamics for maneuvering target interception in three-dimensional environment and terminal angle constraints. The proposed guidance law is summarized as an “observer-controller” system. More specifically, to estimate the knowledge of target maneuvers, an adaptive second-order sliding mode observer is presented whose design parameters can be adjusted autonomously according to the estimation error, and a nonsingular switchable sliding manifold based finite time convergent controller is proposed to drive the line-of-sight angle error and angular rate to a small region around zero, further, enable the missile accurately intercept the target in finite time. Finally, by formulating the autopilot lag as a first- order or second-order dynamics, the higher-order comprehensive model is built and complex guidance laws are presented. The simulation results demonstrated the proposed properties.
       
  • Three-dimensional nonlinear gravity assisted aiming point guidance
    • Abstract: Publication date: February 2019Source: Aerospace Science and Technology, Volume 85Author(s): Shashank H. Ramesh, Radhakant Padhi A novel three dimensional aiming point guidance law is presented in this paper, which eliminates the need for gravity compensation, resulting in engagements occurring with zero acceleration commands, in turn leading to smaller overall control effort and an excellent zero effort miss behavior. The conditions necessary for collision in the presence of gravity are derived using model-based prediction and computationally efficient shooting method. This is followed by employing differential-geometric guidance philosophy for achieving collision with the target. The efficacy of the proposed guidance law is then verified using simulation studies for terminal phase ballistic missile interception, and is compared with proportional navigation, aiming point guidance and compensated-weave guidance. Results indicate that the proposed guidance law ensured collision with desired level of accuracy and outperformed the aforementioned guidance laws used for comparison, in terms of miss distance, zero effort miss behavior and control effort.
       
  • Autonomous Reduced-Gravity Enabling Quadrotor Test-bed: Design, Modelling
           and Flight test Analysis
    • Abstract: Publication date: Available online 11 January 2019Source: Aerospace Science and TechnologyAuthor(s): Kedarisetty Siddhardha This work establishes conventional quadrotors as viable platforms to produce non-zero reduced-gravity. Toward this, a 1-D vertical manoeuvre with time-varying acceleration is proposed, which when followed by the quadrotor, the on-board payload experiences desired reduced gravity level for a specified time period. Since the proposed manoeuvre involves high axial accelerations and velocities, the standard steady-state thrust model cannot be used. A CFD tool—RotCFD—is used to study the thrust variation of a quadrotor in axial flight, and develop a thrust model. The developed model is experimentally validated through series of axial flight tests. As a consequence of the time varying forces—propeller thrust as well as drag—acting on the quadrotor, a control structure with fixed gains is neither capable of stabilizing the attitude nor maintaining the desired accelerations. A novel method of gain compensation is proposed which stabilizes the quadrotor and ensures quick acceleration convergence to the commanded value. Flight tests results for various reduced gravity levels from 0.8g–0.3g for a time interval of 3 seconds are presented to demonstrate the repeatability and reliability of the proposed control and automation strategies. Experimental results show that g-quality of the order 10−3g is achieved.
       
  • Full-length visualisation of liquid oxygen disintegration in a single
           injector sub-scale rocket combustor
    • Abstract: Publication date: Available online 11 January 2019Source: Aerospace Science and TechnologyAuthor(s): Dmitry I. Suslov, Justin S. Hardi, Michael Oschwald This work presents results of an effort to create an extended experimental database for the validation of numerical tools for high pressure oxygen-hydrogen rocket combustion. A sub-scale thrust chamber has been operated at nine load points covering both sub- and supercritical chamber pressures with respect to the thermodynamic critical pressure of oxygen. Liquid oxygen and gaseous hydrogen were injected through a single, shear coaxial injector element at temperatures of around 120 K and 130 K, respectively. High-speed optical diagnostics were implemented to visualise the flow field along the full length of the combustion chamber. This work presents the analysis of shadowgraph imaging for characterising the disintegration of the liquid oxygen jet. The large imaging data sets are reduced to polynominal profiles of shadowgraph intensity which are intended to provide a more direct means of comparison with similarly reduced numerical results. Comparing half-lengths of these profiles across operating conditions show clear groupings of load points by combustion chamber pressure and mixture ratio. All load points appear to collapse to an inverse dependence of length on impulse flux ratio.
       
  • Static performance modeling and simulation of the staged combustion cycle
           LPREs
    • Abstract: Publication date: Available online 10 January 2019Source: Aerospace Science and TechnologyAuthor(s): Mahyar Naderi, Liang Guozhu The objective of the current research is to analyze the steady state performance of staged combustion cycle liquid propellant rocket engines (SCCLPREs) and to provide a comprehensive case study with all the required technical input data. For this purpose, using the modular concept, engine components are classified into 14 modules and a modeling and simulation platform for SCCLPRE integrated system simulation has been developed in MATLAB Simulink. Space shuttle main engine (SSME) has been selected as a case study, and its static performance has been studied. A total of 34 elements has been taken into account, and using 110 linear/non-linear equations the engine's system modular model has been established. Most of the equations are solved by direct substitution of the initial guesses and iteration through eleven nested iterative loops and a few nonlinear complex equations are solved by Newton-Raphson method. The SSME's throttled performance at four different throttling levels has been simulated, and the effect of throttling has been studied on engine major elements. It is found that by throttling, the pressure of preburners, combustion chamber and high-pressure turbopumps are influenced more than other elements. The simulation mean error at 109% rated thrust level (RTL), 100% RTL, 60% RTL and 50% RTL is 1.63%, 5.76%, 12.91% and 15.28% respectively. Comparing to the previous studies, the simulation accuracy has been enhanced 3% to 7%. The developed model is a basis for further research over SSME and other SCCLPRE.
       
  • Numerical Investigation of Aeroacoustics Damping Performance of a
           Helmholtz Resonator: Effects of Geometry, Grazing and Bias Flow
    • Abstract: Publication date: Available online 10 January 2019Source: Aerospace Science and TechnologyAuthor(s): Gang Wu, Zhengli Lu, Xiao Xu, Weichen Pan, Weiwei Wu, Jun Li, J. Ci In this work, the effects of the grazing flow and its geometric dimensions on the aeroacoustics damping performance of a Helmholtz resonator are numerically evaluated. The grazing flow tangentially passing through the neck of the resonator is characterized by Mach number. And it is varied from 0 to 0.1. 2D numerical simulations are conducted by solving linearized Navier-Stokes equations in frequency domain via COMSOL 5.3. The numerical model is first validated by comparing with experimental and theoretical data available in literature, as a low Mach number grazing flow is present only. Good agreement is obtained. The model is then used to examine the effect of the grazing flow and the resonator geometric dimensions on transmission loss (TL) performance. Four key parameters are identified, and they include 1) the neck length, 2) neck diameter, 3) cavity volume and 4) the grazing flow Mach number. It is found that as the grazing flow Mach number is greater than 0.07, increasing the neck length leads to a decreased TLmax. Same observation is found for increased cavity volume Vr. Smaller Helmholtz resonator Vr or shorter neck length Ln is found to be involved with a larger TLmax in the presence of the grazing flow. Further study is then conducted, when there is a joint bias-grazing flow. The bias flow could be injected with respect to the grazing flow in 3 different directions, 1) parallel, 2) perpendicular and 3) counter flow. The bias flow injection direction is found to lead to 5 dB transmission loss difference. Parallel or counter injection of the bias flow is showed to be associated with improved TL performance, since there are counter-rotating vortices produced in the cavity. Finally, increasing the Mach number of the grazing flow leads to the noise damping effectiveness being deteriorated in general, even in the presence of a bias flow. The present work shed lights on the aeroacoustics design of Helmholtz resonators in the presence of a grazing and bias flow.
       
  • Lagrangian analysis of mass transport and its influence on the lift
           enhancement in a flow over the airfoil with a synthetic jet
    • Abstract: Publication date: Available online 9 January 2019Source: Aerospace Science and TechnologyAuthor(s): Shengli Cao, Ya Li, Jiazhong Zhang, Yoshihiro Deguchi The mass transport and lift enhancement in flows over the NACA0015 airfoil with a synthetic jet (SJ) are studied numerically using Lagrangian coherent structures (LCSs). Trajectories of synthetic particles belonging to 9 different groups in a vicinity of the airfoil are tracked to analyze the transport. The influence of mass transport on the aerodynamic performance of the airfoil and on the surface pressure coefficient distribution is studied. The difference in transport properties for the airfoil with and without the SJ is analyzed to understand the reasons of the lift enhancement. The results show that the LCSs could be successfully used to describe transport and mixing in the complex flows in the blowing and suction phases. More, the fluid originating in the upper reaches is continuously sucked out of the flow field in the suction phase, then the jet injected by the jet slot moves to the lower reaches of the jet slot in the blowing phase. Compared with the original airfoil, the controlled airfoil could bring about the fluid adhering to the surface moving constantly and the lift coefficient improved apparently, whether in the suction or blowing phase. Resulting from the SJ, the high momentum fluid is injected into the flow field near the airfoil. Therefore, the boundary layer separation point of the airfoil is pushed downstream, and the area of the flow separation zone is greatly reduced. Generally, utilizing LCS to study mass transport process can provide a new perspective for the study of the SJ.
       
  • Aerodynamic characteristics of flexible wings with leading-edge veins in
           pitch motions
    • Abstract: Publication date: Available online 9 January 2019Source: Aerospace Science and TechnologyAuthor(s): Yeong Gyun Ryu, Jo Won Chang, Joon Chung To demonstrate the effects of wing deformations on aerodynamic performances during the wing reversal, aerodynamic force/torque and flow vector-fields were measured. Wing models consisted of wing planes with various thicknesses and two leading-edge veins, which obstructed spanwise deformations (Case 1 as a rigid wing and Cases 2 to 4 as flexible wings). They also underwent three different pitching periods (t/TR,α=0.1,0.2, and 0.4) to determine the distinct changes in vortical structures and corresponding aerodynamic characteristics. Flexible wings generally showed a negative camber in the stroke motion, causing poor aerodynamic performances rather than a rigid wing. This was caused by the positive camber and camber change to the negative during the pitch motions. After the start of the stroke, the positive camber caused separated and weakened tip vortex (TV) structures, hindering leading-edge vortex (LEV) formations around the wingtip. Depending on the LEV shedding, the flexible wings had less aerodynamic forces than the rigid wing in stroke motion. During the pitch motions, on the other hand, the dynamic cambers influenced rotational mechanisms such as the wing-wake interaction and rotational force, causing higher or less initial lift increment. As the t/TR,α increased, the amount of the lift augmentation decreased due to the weakened wing-wake interaction, instead showing a downwash. At t/TR,α=0.4, however, the higher initial lift peak was found because of the absence of the U_TEV2, leading to the decline of the downwash. The delayed U_LEV1 dispersal and the U_TEV2 traces generated the similar induced flows to the wing-wake interaction, resulting in the higher lift augmentation. Furthermore, Case 2 achieved the higher lift augmentations than the other cases in all pitching periods. The slight wing deformations not only reduced the distance between the vortices and the wing surface, but also caused the delayed vortex dispersals. These induced the stronger rapid flows toward the wing surface, causing the higher initial lift peaks. Case 2 also had the highest C‾L/C‾P,t at t/TR,α=0.1 than the other cases. These results provide a good agreement to have higher aerodynamic efficiency in its specific chordwise flexibility range.
       
  • Evaluation of the relationship between the aerothermodynamic process and
           
    • Abstract: Publication date: Available online 9 January 2019Source: Aerospace Science and TechnologyAuthor(s): Trung Hieu Nguyen, Phuong Nguyen Tri, Francois Garnier Gaining a full understanding of the aerothermodynamic (AT) process in the high-pressure turbine of an aircraft engine is a complex task. The data in the literature on this topic are very limited, particularly that on tridimensional simulation. We report herein, for the first time, the designs, computer simulations, tridimensional calculations of the interactions of AT parameters under various operation conditions (takeoff, cruise) of an aircraft engine by using a refined mesh system containing approximately 6 million polyhedrons and more than 1 million interactions to solve equations in a high-pressure turbine (HPT). Our research provides, for the first time, an overview of high-resolution topographic images of the distribution of AT parameters in multi-row turbomachinery. Our calculations indicate that the relationship between these parameters was convoluted and depended on each operating condition. For example, our findings show that the thermal boundary conditions and rotor speeds strongly affected the flow temperatures (14%) and flow velocity (31%). On the other hand, the cooling system appeared not to affect AT parameters. The temperature nonuniformities in the turbine in the axial and radial directions were also observed.
       
  • Improved artificial potential field based lateral entry guidance for
           waypoints passage and no-fly zones avoidance
    • Abstract: Publication date: Available online 9 January 2019Source: Aerospace Science and TechnologyAuthor(s): Zhenhua Li, Xiaojun Yang, Xiangdong Sun, Gang Liu, Chen Hu In this paper, a novel improved artificial potential field based lateral guidance algorithm for waypoints passage and no-fly zones avoidance is proposed. By introducing the improved attractive potential field and the improved repulsive potential field, the proposed algorithm converts the waypoints passage problem and the no-fly zones avoidance problem into the reference heading angle determination problem. Based on the heading error threshold and heading angle constraints, the reference heading corridor can be obtained in real time, which is adaptively updated during the entry flight according to the waypoint and no-fly zone constraints. Then the bank reversal logic is employed to regulate the sign of the bank angle, control the lateral motion of the vehicle, and steer the vehicle successfully to the desired terminal zone without any collision of the waypoint and no-fly zone constraints. The simulation results for verification and comparison cases show that the proposed algorithm has better adaptability, applicability and robustness than the alternative methods, and it can be successfully applied in lateral entry guidance of hypersonic glide vehicles with waypoint and no-fly zone constraints.
       
  • Velocity behavior downstream of perforated plates with large blockage
           ratio for unstable and stable detonations
    • Abstract: Publication date: Available online 9 January 2019Source: Aerospace Science and TechnologyAuthor(s): Bo Zhang, Hong Liu, Bingjian Yan Due to the excellent thermal propulsion performances of detonation, it has been applied for the purpose of aerospace propulsion devices, e.g., pulse detonation engines (PDEs), rotating detonation engines (RDEs), and oblique detonation wave engines (ODWEs). However, it remains challenging for developing those new-concept propulsion devices, mainly because it is a formidable task to establish a steady and self-sustained detonation in the hypersonic flow and combustible mixture. One of the fundamental problems is to understand the interaction of diffractions and the detonation waves within the engines. In this study, perforated plates with various blockage ratios are seated at the beginning of the detonation propagation to explore the diffractions that generated from the plates on the detonation propagation mechanism. Two explosive mixtures of C2H2 + 5N2O and C2H2 + 2.5O2 + 70%Ar are studied to illustrate the difference in propagation velocity behavior of detonation after it suffers from large-scale diffractions. The results show that diffractions have less effect on the propagation of highly unstable mixture, and the effect becomes obvious as BR increases to 0.962; this phenomenon occurs because detonation has an irregular cellular pattern and sub-structures are characterized by highly unstable detonation, in which the detonation instabilities are amplified by the large perturbations of obstacles, leading to an augment in forming more cellular cells of detonation in its front that cover the weakening effects from diffractions. In contrast, the diffractions significantly affect the stable mixture, manifested by a remarkable increase of the critical pressure for a self-sustained detonation downstream of the obstacle with the augment of blockage ratio; this phenomenon can be attributed to the diffractions being distributed along the curvature over the detonation surface, thereby causing more excessive curvature of the entire detonation front, which in turn exacerbates the failure of detonation.
       
  • Active aeroelastic output feedback control with partial measurements by
           the method of receptances
    • Abstract: Publication date: Available online 8 January 2019Source: Aerospace Science and TechnologyAuthor(s): Kumar Vikram Singh, Charlene Black, Raymond Kolonay The method of receptances allows the design of active aeroelastic control for suppressing the flutter instabilities and for augmenting the flutter boundaries by pole placement. This method is purely based on available receptance transfer functions, which can be available for controller design from the embedded sensors and actuators on the aircraft. In recent years, receptance based single-input and multiple-input full state feedback control has been developed, which requires the displacement and velocity measurements associated with each degree of freedom, which may not be available in practice. The design of a controller using partially available measurements in these scenarios is related to the static output feedback design, which leads to the solution of associated nonlinear problems and/or the design of complimentary state estimators or observers. In order to overcome such restrictions, in this paper a method for feedback control design using the receptance method is presented which can utilize the available partial measurements for control gain computation. This is achieved by introducing redundancy of acceleration measurement and adding associated acceleration feedback in the controller loop. This leads to a linear constrained least-square problem for the design of the controller. The working of this approach is shown with numerical examples associated with various wing models. It is demonstrated that prescribed aeroelastic modes can be controlled to achieve flutter suppression and flutter boundary extension, by designing a feedback controller using partial measurements from the available sensors. Parametric studies are presented to demonstrate that such output feedback control allows the optimal selection of measurements and the corresponding sensor location to minimize the controller norm.
       
  • NUMERICAL INVESTIGATION INTO THE UNDERLYING MECHANISM CONNECTING THE
           
    • Abstract: Publication date: Available online 8 January 2019Source: Aerospace Science and TechnologyAuthor(s): Yanhui Wu, Guangyao An, Bo Wang This paper presents a series of systematic multi-passage unsteady RANS simulations on a transonic axial flow compressor rotor (Rotor 35). The objective is to have a better understanding of the underlying flow mechanism which connects the phenomena of the tip leakage vortex (TLV)'s breakdown to the appearance of flow unsteadiness. It has been revealed that both bubble-type and spiral-type breakdown of the TLV can result in a self-sustained flow unsteadiness at high-loading flow conditions. The origin of such unsteadiness lies in that the vorticity region redistributed by the vortex breakdown is capable of affecting the pressure distribution on the pressure side of a passage. Once this threshold event is met, the swirl intensity of the TLV and the strength of shock wave, which are key factors controlling the vortex breakdown, varies accordingly, thus leading to an instantaneous rather than a stationary vortex breakdown occurring in the confined rotor-passage system. As compared to the bubble-type breakdown of TLV, the spiral-type breakdown of TLV exerts more severe impact on the pressure distribution on the PS of the passage. As a result, a significant change in the scale of the breakdown region occurs. This gives rise to not only an appearance of a new vortex structure but also a blockage transfer across the passage against the rotor turning direction. The new vortex structure, termed as tip secondary vortex (TSV), is essentially a vortex segment arising from the spiral-type breakdown of TLV. The necessary condition for the inception of the rotating wave like RI is the blockage transfer induced by the spiral-type breakdown of TLV and its resultant interaction with the tip leakage in the adjacent passage.
       
  • Forecasting-based data-driven model-free adaptive sliding mode attitude
           control of combined spacecraft
    • Abstract: Publication date: Available online 8 January 2019Source: Aerospace Science and TechnologyAuthor(s): Han Gao, Guangfu Ma, Yueyong Lv, Yanning Guo In this paper, a novel forecasting-based data-driven model-free adaptive sliding mode attitude control (FMFASMC) method is proposed for the postcapture combined spacecraft in the presence of unknown mathematical model and external disturbance. First, a forecasting-based data-driven model-free adaptive controller (FMFAC) is developed to ensure its adaptation to various working conditions by simultaneous adjustment of controller parameters with online and offline input–output measurement data. Then, a sliding-mode-based supplementary controller is introduced to improve the tracking performance of FMFAC in terms of robustness. Compared with existing works, the designed control scheme only utilizes the attitude angle and attitude angular velocity of the combined spacecraft, not requiring corresponding mechanism model, thereby dramatically curtailing the complexity and difficulty of relevant controller design. Finally, simulation comparisons between the FMFASMC, FMFAC, and traditional model-free adaptive control (MFAC) are presented for two groups of illustrative examples. The simulation results verify the effectiveness of the proposed control method.
       
  • Experimental test of a 3D parameterized vane cascade with non-axisymmetric
           endwall
    • Abstract: Publication date: February 2019Source: Aerospace Science and Technology, Volume 85Author(s): Zhendong Guo, Hongyan Bu, Liming Song, Jun Li, Zhenping Feng Experimental test of a 3D parameterized vane cascade with non-axisymmetric endwall was conducted by using the particle image velocimetry (PIV) and five-hole probe. Both the straight annular cascade with axisymmetric endwall (denoted by Ref) and the negatively bowed cascade with non-axisymmetric endwall (denoted by Opt) were tested. Here, the Opt design was obtained through optimization based on Computational Fluid Dynamics (CFD), by combining the techniques of section profiling, compound lean and non-axisymmetric endwall [1]. The results showed that, the pressure distributions of blade surface at 10%, 50% and 90% spans of CFD results and experimental data are in good agreement. The trends of total pressure coefficient at cascade outlet (denoted by Cpt) and outlet flow angle (denoted by α1) were well matched; the Cpt of Opt design was shown increased by 0.47%, and 0.41% in experiment and CFD simulations, respectively. The contours of Cpt showed that the profile loss of Opt design is significantly reduced and the secondary loss region gets closer to hub and shroud. By tracking the secondary vortices near the hub of cascades, the affected area of secondary flow of Opt design was confirmed to be smaller when compared to that of the Ref design; but the size of passage vortex near the cascade exit was increased. The results of PIV and five-hole probe were consistent; they both confirmed that the combined parameterization of both blade and endwalls is effective in reducing the affected area of secondary flow and the profile loss of the low-aspect-ratio vane cascade.
       
  • Design methodology using characteristic parameters control for low
           Reynolds number airfoils
    • Abstract: Publication date: Available online 7 January 2019Source: Aerospace Science and TechnologyAuthor(s): Sen Zhang, Huaxing Li, Afaq Ahmed Abbasi In the past century, low Reynolds number airfoil design for both military and civil applications has continued to capture the interest of practitioners of applied aerodynamics. A new design methodology for airfoils has been developed, which controls the airfoil profile through 8 parameters. The airfoil aerodynamic characteristics were predicted through the XFOIL code, to optimize airfoil under mode FRONTIER environment. The reliability verification result predicted that the XFOIL code can be used to determine the airfoil performance in the linear segment of the lift coefficient curve. Application of this methodology was demonstrated through two cases. For Case-I, airfoils design were based on E387 airfoil profile and synthetic aerodynamic performance at different angles of attack for Re=2.0×105 were considered. While for Case-II, designed airfoils were based on PSU 94-097 airfoil with synthetic aerodynamic performance at different angles of attack for Re=4.0×105. Results of both cases showed enhanced aerodynamic properties for the designed airfoils.
       
  • Hydrodynamic performance in a sloshing liquid oxygen tank under different
           initial liquid filling levels
    • Abstract: Publication date: Available online 7 January 2019Source: Aerospace Science and TechnologyAuthor(s): Zhan Liu, Yuyang Feng, Gang Lei, Yanzhong Li Fluid sloshing usually results in some unintended consequences, including additional sloshing force and thermodynamic imbalances, it is necessary to conduct in-depth investigations on fluid sloshing hydrodynamics to ensure the safe operation of spacecraft. To predict the sloshing hydrodynamic in a non-isothermal liquid oxygen tank, a numerical model is built with fluid sloshing and heat transfer coupled. Both the external environmental heat leak and the interface phase change are considered in detail. The external sinusoidal excitation is realized by user defined functions. Coupled with the mesh motion treatment, the volume of fluid method is used to capture the movement of liquid-vapor interface during sloshing. Compared against the experiment results, the standard k–ε model has been proven to have great prediction accuracy. Based on the numerical model, effect of the initial liquid filling level on the sloshing hydrodynamic performance is studied. The vapor and liquid pressure, the sloshing force and the related sloshing moment, and the dynamic responses of the liquid-vapor interface are discussed and analyzed. With some valuable conclusions arrived, the present study is significant to better understanding of the non-isothermal sloshing dynamic process.
       
  • Flight performance simulation and station-keeping endurance analysis for
           stratospheric super-pressure balloon in real wind field
    • Abstract: Publication date: Available online 7 January 2019Source: Aerospace Science and TechnologyAuthor(s): Huafei Du, Jun Li, Weiyu Zhu, Zhipeng Qu, Lanchuan Zhang, Mingyun Lv For the absence of the propeller system to resist wind load, the balloon will float everywhere with unstable position. In order to fix the balloon within a specific district and extend the station-keeping endurance, an effective approach that utilizes the wind from different directions by altering the float altitude was proposed. A program was developed to simulate the flight performance and the station-keeping endurance of the stratospheric super-pressure balloon in the real wind field. The influences of the initial Helium volume ratio and the float altitude on the travelled horizontal distance of the balloon were analyzed. The results indicate that the altitude of the balloon fluctuates with the change of the volume of the balloon during the day-night cycle, providing the possibility to catch the wind from different directions. By choosing the appropriate Helium volume ratio and float altitude, the travelled horizontal distance of the balloon can be shorten greatly. Our study can provide guidance for investigation on active control algorithm for station-keeping of the balloon that can fix the balloon within the desired district for months and years.
       
  • Comparison of Wind Speed Models within a Pitot-Free Airspeed Estimation
           Algorithm using Light Aviation Data
    • Abstract: Publication date: Available online 7 January 2019Source: Aerospace Science and TechnologyAuthor(s): Matthew B. Rhudy, Mario L. Fravolini, Marco Porcacchia, Marcello R. Napolitano This paper presents an analytical redundancy-based approach to the estimation of the true aircraft airspeed without the use of a Pitot probe. Measurements from an Inertial Measurement Unit (IMU), Global Positioning System (GPS), and angle of attack and sideslip angle vanes are used within a sensor fusion algorithm utilizing the kinematic model of a fixed-wing aircraft to co-estimate the airspeed, wind speed, and attitude of an aircraft. The presented model does not use information relative to the aircraft dynamic model, and therefore, is totally independent of the specific aircraft. Due to the necessity of the algorithm to determine the aircraft wind speed, a predictive model of the wind behaviour is necessary. This work compares two different stochastic wind models - the random walk (RW) and the Gauss-Markov (GM) - within the context of the airspeed estimation problem using flight data of a light aviation aircraft. The use of light aviation data is an important new consideration since previous work has focused on unmanned aircraft. The results of this study indicated that the RW model provides better performance over the GM model for airspeed estimation. Additionally, the light aviation data further reinforces the cross-platform capability of the considered algorithm and, furthermore, demonstrates the effectiveness of this method within manned flight.
       
  • Sandbraking. A technique for landing large payloads on Mars using the
           sands of Phobos
    • Abstract: Publication date: February 2019Source: Aerospace Science and Technology, Volume 85Author(s): Francisco J. Arias, Salvador De Las Heras The basis of a novel braking technique by using the Phobos sands for landing large payloads on Mars is outlined. Here consideration is given to the utilization of the Phobos or Deimos regolith as material for aerobraking by discharging a load of sand at certain distance in front of the spacecraft during the descent manoeuver. Although immediately after getting rid the load of sand in front of the spacecraft they have a null relative velocity with the spacecraft, however, because the stronger atmospheric drag acting on the tiny particles of sand they will be promptly decelerated. As a result, the particles of sand will impact onto the front of the spacecraft with a velocity close to the terminal velocity of the spacecraft itself. By using a pusher-disc – or akin damping system, in front of the spacecraft the momentum exchange from the sand collisions will result in a braking force acting on the spacecraft. Due to the very small delta-v budget required to lift material from the surface of Phobos or Deimos to their transfer orbits, then a small amount of dedicated rocket chemical propellant brought from Earth could be transformed into a huge amount of sand lifted from the surface of Phobos of Deimos to their transfer orbits. The large thrust generated by the Sandbraking makes this technique propitious for landing of planetary bodies struggling against gravity.
       
  • Composite block backstepping trajectory tracking control for disturbed
           unmanned helicopters
    • Abstract: Publication date: February 2019Source: Aerospace Science and Technology, Volume 85Author(s): Xiangyu Wang, Xin Yu, Shihua Li, Jiyu Liu In this paper, the position and yaw angle trajectory tracking control problem is studied for unmanned helicopters subject to both matched and mismatched disturbances. To achieve the trajectory tracking goal, a feedforward-feedback composite control scheme is proposed based on the combination of the generalized proportional integral observer and the block backstepping control techniques. The controller design process mainly consists of two stages. In the first stage, some generalized proportional integral observers are developed for the helicopter system to estimate the mismatched, matched disturbances and their (higher-order) derivatives. In the second stage, the composite controller is designed by integrating the block backstepping control method and the disturbance estimates together. The proposed composite scheme guarantees asymptotic tracking performances for the position and yaw angle of the helicopter to the desired trajectories even in the presence of fast time-varying disturbances. Numerical simulations demonstrate the effectiveness of the proposed composite control scheme.
       
  • Investigation on cooling effect with a combinational opposing jet and
           platelet transpiration concept in hypersonic flow
    • Abstract: Publication date: February 2019Source: Aerospace Science and Technology, Volume 85Author(s): BinXian Shen, Liang Yin, XiaoLi Zhang, WeiQiang Liu A platelet transpiration is introduced to strengthen the cooling effect in reattachment region for opposing jet in hypersonic flow. This introduced means can strengthen the local thermal protection performance without leading to an overall enhanced cooling intensity, being helpful on saving the total cooling gas. The two-equation SST k–ω turbulence model has been utilized to study the heat reduction and flow fields of the simplified combinational nose-tip in hypersonic flow. At first, the influence of individual pore are analyzed. The obtained results show that transpiration gas forms a cooling film adhering to the body surface which isolates the serious aerodynamic heating. The transpiration pores which are located on upstream margin of the reattachment point exhibit the best cooling effect. Then, the combination nose-tip is improved. A limited amount of pores are arranged in the reattachment region merely to strengthen the local cooling efficiency. The cooling capacity is obviously promoted with transpiration gas in contrast to that without transpiration. The peak heat flux reduces more than 12% with the mass flux of total cooling gas only promotes by 7.5%. Finally, the study demonstrates that the simplified model with limited amount of pores promote the cooling efficiency in contrast to that without transpiration.
       
  • CFD Based Criteria of Stall Onset in Centrifugal Compressors
    • Abstract: Publication date: Available online 4 January 2019Source: Aerospace Science and TechnologyAuthor(s): Adel Ghenaiet, Smail Khalfallah The determination of stall onset is very useful for the control and safe operation of centrifugal compressors at high pressure ratios. This is very delicate and intricate task since the details of the flow need to be acquired at several measuring planes. The state-of-the-art CFD tools have contributed substantially in the analysis of the flow structures and the design of centrifugal compressors. The present paper demonstrates the potentiality of numerical simulations to map the full flow field and assess some criteria to predict the stall onset, typically those relating the boundary layer growth and stability. Through the entire impeller and diffuser, the quantification of the blockage factor in addition to the modified loading criteria (diffusion rate) which combines the blade-to-blade and hub-to-shroud Richardson numbers seem more suitable in predicting the stall onset and stall cells positions. The well known models of centrifugal compressors operating at low-speed (LSCC) and at a high-speed (DLRCC) served for the validations. Besides, the present CFD based criteria are useful since they may easily integrate the design optimization chain.
       
  • Scaled Sequential Threshold Least-Squares (S2TLS) Algorithm for Sparse
           Regression Modeling and Flight Load Prediction
    • Abstract: Publication date: Available online 4 January 2019Source: Aerospace Science and TechnologyAuthor(s): Shengwei Zhu, Yi Wang This paper presents a Scaled Sequential Thresholded Least Squares (S2TLS) algorithm to construct sparse regression models for flight load prediction. The combined use of a sparsification parameter λ and a magnification factor χ is proposed to tune both the model complexity and the regressor complexity. A bumpiness function is introduced to preferentially select simple regressors to improve model prediction. A cost function J consisting of the estimation residual and the bumpiness is then proposed to determine parameters (χ, λ) satisfying balanced performance. Parametric analysis is undertaken to investigate the effect of χ and λ on sparse regression performance. It is found that the optimal solution is hidden within a trench-like region of the (χ, λ) domain. Two methods using different optimization variables and algorithms are then presented to search for optimal combinations of λand χ. Case studies are performed, and model results are compared against the test and CFD data of flight loads. Excellent agreement is observed, and the data (even with significant complications) is well bounded by the 95% confidence interval. Importantly, the underlying load-driving factors can be successfully identified. The new S2TLS algorithm represents a robust, efficient, and accurate method for flight load modeling and prediction.
       
  • Adaptive actuator failure compensation control for hypersonic vehicle with
           full state constraints
    • Abstract: Publication date: Available online 3 January 2019Source: Aerospace Science and TechnologyAuthor(s): Yong Xi, Yao Meng This paper addresses the longitudinal control problem of air-breathing hypersonic vehicles subject to actuator faults, full state constraints and backlash-like hysteresis. To accomplish the full state constraints, the modified nonlinear mappings are proposed to transform the altitude subsystem to a new system without state constraints. An adaptive backstepping controller is developed for this new system, in which the command filters are embedded to eliminate the problem of “explosion of terms”. Then, the controller parameters are updated to compensate the dysfunction of failed actuators and attenuate the unfavorable effects resulted from the backlash-like hysteresis. It is proved that the proposed controller can guarantee that the tracking errors converge to an arbitrarily small residual set and the full state constraints are not violated. Simulation results are carried out to demonstrate the effectiveness of the proposed control scheme.
       
  • Investigation on wall shear stress measurement in supersonic flows with
           shock waves using shear-sensitive liquid crystal coating
    • Abstract: Publication date: Available online 3 January 2019Source: Aerospace Science and TechnologyAuthor(s): Jisong Zhao Wall shear stress (skin friction) is one of the fundamental surface quantities in aerodynamics but its measurement remains challenging. This paper studies the measurement of wall shear stress vector fields in supersonic flows with shock waves using the shear-sensitive liquid crystal coating (SSLCC) technique. A previously developed SSLCC technique was improved and used for the measurement of wall shear stresses in supersonic flows. Experimental results show that the flow structures including shock waves in a supersonic jet flow were visualized by the SSLCC technique; furthermore, the shear stress vector field over a planar surface induced by the supersonic jet flow was measured by the SSLCC technique. The shock waves and the compression-expansion repetitions in the supersonic jet flow were captured, respectively, by analysing the shear direction field and the skin friction lines. This work demonstrates the capabilities of the SSLCC technique to visualize and measure wall shear stresses in supersonic flows with complex shock waves.
       
  • Assessment of transition modeling for high Reynolds flows
    • Abstract: Publication date: Available online 3 January 2019Source: Aerospace Science and TechnologyAuthor(s): Konstantinos Diakakis, George Papadakis, Spyros G. Voutsinas Variants of the eN method and γ–Reθ transition model are thoroughly evaluated against measured data obtained for four airfoils in conditions that range from Re=3 to 40 millions. For the eN method, the option of obtaining the necessary boundary layer thickness properties by integrating the boundary layer equations is compared to that of retrieving this information directly from the flow-field data. It is found that the two options agree when the former uses the Falkner–Skan profiles and the latter utilizes a significantly finer than the baseline grid. Regarding the γ–Reθ model, the effect of inlet conditions and the significance of limiting turbulence production on the accuracy of transition prediction are investigated and found to be important. Regarding the correlation to measurements, in all of the cases considered, even though the two model sometimes agreed qualitatively, the variant of the eN variant that is based on the Falkner–Skan profiles outperformed the γ–Reθ model, especially for Reynolds numbers higher than 5 millions.
       
  • Circumferentially propagating characteristic dominated by unsteady tip
           leakage flow in axial flow compressors
    • Abstract: Publication date: Available online 3 January 2019Source: Aerospace Science and TechnologyAuthor(s): Jichao Li, Shaojuan Geng, Juan Du, Hongwu Zhang, Chaoqun Nie The unsteady tip leakage flow (UTLF) is experimentally studied to observe the circumferential propagation in axial flow compressors. The evolutionary process of the circumferential propagation is captured by utilizing a collection of time-resolved pressure transducers on the casing with circumferential and chord-wise spatial resolution. Results show that the circumferential propagation dominated by the UTLF exists and occurs only after the emergence of the UTLF in the throttling process. Since then, the propagating speed and the scale of disturbance gradually augment, until the transition to the propagating speed and scale of stall inception. The tip air injection is applied to further verify the aforementioned circumferential propagation. Two kinds of injected momentum ratios, namely, micro and macro injection adopted in single-rotor compressor, and two types of injected methods, i.e., external air source and self-recirculation applied in single-stage and three-stage compressor demonstrate that either micro injection with a small stall margin improvement (SMI) or macro injection with a large SMI, or the different injected types in single- and three- stage compressor, with the increasing stability-enhancing capability of tip air injection, the UTLF is effectively weakened, and the circumferentially propagating speed and the scale of the disturbance dominated by UTLF accordingly shows a decreasing trend. The above results provide guidance for understanding the relationship between the UTLF and stall inception in terms of circumferential propagation that can be used to predict the compressor stability.
       
  • Fusing wind-tunnel measurements and CFD data using constrained gappy
           proper orthogonal decomposition
    • Abstract: Publication date: Available online 3 January 2019Source: Aerospace Science and TechnologyAuthor(s): M. Mifsud, A. Vendl, L. Hansen, S. Görtz In this article gappy proper orthogonal decomposition (POD) is used to fuse wind-tunnel measurements and computational fluid dynamics (CFD) data to provide a consistent and more comprehensive output of greater utility. The technique is used to fuse a very limited and ‘gappy’ set of wind-tunnel surface pressure measurements with computational surface pressure and friction data. By integrating the resulting fused surface data over a complete aerodynamic surface, the corresponding aerodynamic coefficients are obtained. A comparison between these integrated aerodynamic coefficients and the wind-tunnel balance measurements show a better agreement than that obtained between CFD and the wind-tunnel experiment. In addition, the original gappy POD algorithm is modified and extended to provide an accurate and consistent match between the integrated coefficients evaluated from the fused data set and the wind-tunnel balance measurements, thus producing a set of data distributed over the entire aerodynamic surface that conforms to experiment. This is achieved by considering the wind-tunnel tunnel balance measurements as equality constraints in the gappy POD least-squares problem formulation. The technique is demonstrated for steady flow around an RAE 2822 airfoil and an industrial wing-body aircraft configuration. Also, an important feature of the methodology that is essential for its use in an industrial environment, namely regularization, is discussed.
       
  • Flow characteristics of monopropellant micro-scale planar nozzles
    • Abstract: Publication date: Available online 3 January 2019Source: Aerospace Science and TechnologyAuthor(s): Daniel T. Banuti, Martin Grabe, Klaus Hannemann We investigate the flow in planar microscale nozzles and find that design and analysis paradigms based on the assumption of a dominant isentropic core with moderate viscosity corrections are not valid. Instead, the flow downstream of the throat is dominated by boundary layers that may choke the flow to subsonic velocities. The geometrical expansion ratio is found to be essentially irrelevant, instead, the length from throat to exit plane is found to be a much more important design parameter. Full 3D simulations are required to predict the flow topology; thermophysical modeling of the expanding gas has a noticeable impact on predicted performance. An analytical estimation of the Knudsen number in the expanding flow is given, allowing to determine its values from the expansion pressure ratio. An axial thrust analysis suggest truncation of the nozzle, resulting in a predicted 20% increase in thrust and 30% increase in specific impulse compared to the baseline configuration. The work has been carried out within the European Commission co-funded PRECISE project which was focused on designing and testing a micro chemical propulsion system thruster prototype using catalytically decomposed hydrazine as propellant.
       
  • Research on gravity vertical deflection on attitude of position and
           orientation system and compensation method
    • Abstract: Publication date: Available online 3 January 2019Source: Aerospace Science and TechnologyAuthor(s): Zhuangsheng Zhu, Bo Zhao, Yiyang Guo, Xiangyang Zhou Position and Orientation System (POS) is the key sensor of remote-sensing system. It is a typically Strapdown Inertial Navigation System (SINS) and Global Navigation Satellite System (GNSS) integrated measurement system, which provides high-precision position and attitude parameters. The gravity disturbance vector is one of the main error sources of high-precision SINS. The error propagation of POS is analyzed. It shows that the vertical component of the gravity disturbance vector is introduced into the POS horizontal channel through the vertical deflections, which leads to the error of the attitude measurement of the POS. In this article, a new real-time gravity compensation method is proposed, which includes the gravity disturbance as the error states of POS Kalman filter, and the accurate gravity disturbance model is constructed by a “Block-Time Variation” Markov Model (B-TV-MM) based on high-precision gravity map, whose resolution is enhanced by a new interpolation method based on Gaussian Process Regression (GPR). A flight experiment was conducted to evaluate the efficiency of the proposed method and the results showed that the proposed method performs better compared with other real-time gravity compensation methods.
       
  • Improvements of performance and stability of a single-stage transonic
           axial compressor using a combined flow control approach
    • Abstract: Publication date: Available online 3 January 2019Source: Aerospace Science and TechnologyAuthor(s): Jiaguo Hu, Rugen Wang, Danqin Huang The performance and the stability of axial compressors are seriously concerned as it's widely used in aero-engines and power machines. A new combined flow control approach was developed previously using blade slot and vortex generator, which was used to reduce flow loss and improve flow stability. The approach was experimentally proved to be very effective in linear compressor cascades. To further evaluate its availability in actual compressors, the new approach is implemented in a single-stage transonic compressor, which introduces a part-span slot into the rotor while a full-span slot and a vortex generator into the stator. CFD simulations are performed to evaluate the gains of compressor performance, and flow structures are analyzed to explain the corresponding mechanisms. Results show great improvements of performance and stability of the compressor. The averaged total pressure ratio and isentropic efficiency improve by 1.82% and 0.88% respectively due to the great reduction of separations in both the rotor and the stator. Because the compressor instability is initiated by the stator corner stall, the delay of the corner separations leads to a great increase of the stable mass-flow range by 26.1%. The compressor flow structures are also well improved because the detrimental backflows and large-scale vortices are mostly removed by the combined approach.
       
  • Active control method for restart performances of hypersonic inlets based
           on energy addition
    • Abstract: Publication date: Available online 2 January 2019Source: Aerospace Science and TechnologyAuthor(s): Hongkang Liu, Chao Yan, Yatian Zhao, Sheng Wang The restart performance of the inlet is crucial to a hypersonic air-breathing propulsion vehicle. An active flow control method based on energy addition for the restart of hypersonic inlets is brought forward in this paper. Using Reynolds Averaged Navier–Stokes equations method, the effects of energy addition parameters on the restart performances of hypersonic inlets are investigated and the restart process is presented. This study verifies that the present method is feasible to improve the restarting capability of hypersonic inlets. The restart performances rely heavily on the energy addition parameters. Using energy addition properly, it usually yields a better inlet performance and more importantly enables the inlet restart again; otherwise, the large separation bubble still exists. Investigations on energy addition parameters further reveal that energy addition always increases the mass flow rate, and its center far from the cowl facilitates restarting the inlet and decreasing the stagnation pressure losses. It also appears advantageous to work with energy addition at a moderate power consumption and effective radius. Besides, the restart process indicates that the parabolic shock induced by heated region plays an important role on the restart performance. It alters the separation bubble, broadens the flow passage and enables more deflected air into the inlet. The adverse pressure gradients in the entrance are greatly changed as well. Finally, results also suggest that an appropriate effective time for energy addition is sufficient to restart the inlet. The variations of hysteresis loops show that energy addition prompts the separated bubble to disappear and the inlet to restart in advance. As a result, the restarting Mach number declines and the operation range of hypersonic inlets prominently enlarges.
       
  • Vortex dynamic mechanism of curved blade affecting flow loss in compressor
           cascade during corner stall process
    • Abstract: Publication date: Available online 2 January 2019Source: Aerospace Science and TechnologyAuthor(s): Xiaoxu Kan, Songtao Wang, Ling Yang, Jingjun Zhong In general, understanding the cascade flow loss mechanism is crucial for the design and optimisation of the cascade from the basically essential. The research objective of this study was to develop a highly loaded compressor linear cascade, and a numerical simulation was performed to obtain both the vortex structures and the flow loss of the cascade during the corner stall process. The results indicate that the curved blade affects the influence range of the vortex structures by changing the pressure gradients of the cascade, and subsequently changes the transport process of the vortex structure to low-energy fluid clusters, thus affecting the weight distribution of the flow loss. Additionally, the weight coefficients of the passage vortex and concentrated shedding vortex are merged together, and accounts for half of the total flow losses. Finally,the innovation of this paper is to propose a topological analysis method based on the accurate and quantitative identification of the singular points' positions serves as a useful research method to reveal the vortex dynamic mechanism of the weight distribution of the flow loss affected by the curved blade, and the vortex criterion of the corner stall is proposed additionally.
       
  • A visual/inertial integrated landing guidance method for UAV landing on
           the ship
    • Abstract: Publication date: Available online 2 January 2019Source: Aerospace Science and TechnologyAuthor(s): Yue Meng, Wei Wang, Hao Han, Jingxuan Ban This paper presents a visual/inertial integrated guidance method for UAV shipboard landing. The airborne vision system is utilized to track infrared cooperated targets on the ship and output their center coordinates in the image. Meanwhile, the attitude information of the UAV is obtained from airborne Inertial Measurement Unit (IMU). Extended Kalman Filter (EKF) is chosen to fuse the visual and inertial information. The UAV's position, attitude, and velocity relative to the runway and the ship motion information are estimated. The ship motion information is utilized to predict the position of the intended landing point at the touchdown moment through the Autoregressive algorithm. And the motion information of the UAV is used to calculate the deviation from the intended landing path. All the information is delivered to the flight control system to calculate control command. Simulation shows that the visual/inertial integrated landing guidance system achieves satisfied estimation and prediction results.
       
  • Aerodynamic inverse design using multifidelity models and manifold mapping
    • Abstract: Publication date: February 2019Source: Aerospace Science and Technology, Volume 85Author(s): Xiaosong Du, Jie Ren, Leifur Leifsson Aerodynamic inverse design is proposed using multifidelity models and the manifold mapping (MM) technique. Aerodynamic inverse design aims at achieving a target performance characteristic, such as a pressure coefficient distribution of an airfoil or local lift distribution of a wing. Due to the high computational cost of accurate aerodynamic models and the large number of design variables, the overall cost of inverse design can be prohibitive. The MM-based optimization algorithm leverages the speed of the low-fidelity model to accelerate the optimization process, but refers back to the high-fidelity model to ensure an accurate solution. In this work, the MM technique is applied to the characteristic distribution under consideration in each application. In particular, the pressure coefficient distribution is modeled with the MM technique in the case of airfoil inverse design, and the sectional lift distribution in the case of wing design. The proposed method is tested and evaluated on six airfoil inverse design cases and one rectangular wing inverse design case. In the two-dimensional cases, parameterized with eight design variables, direct aerodynamic inverse design using pattern search required 700 to 1200 high-fidelity model evaluations, which took 300 to 700 hours in total. The MM-based design algorithm required less than 20 high-fidelity simulations and 1000 to 2000 low-fidelity evaluations, which took 30 to 90 hours. In the three-dimensional case, parameterized with three design variables, direct aerodynamic inverse design took around 50 hours, whereas the MM-based design needed around six hours.
       
  • Field-of-view constrained two-stage guidance law design for
           three-dimensional salvo attack of multiple missiles via an optimal control
           approach
    • Abstract: Publication date: February 2019Source: Aerospace Science and Technology, Volume 85Author(s): Xiaolin Ai, Linlin Wang, Jianqiao Yu, Yuanchuan Shen This study considers the cooperative guidance problem for multiple missiles against a stationary target with the field-of-view constraint. A two-stage guidance scheme is proposed to realize salvo attack of multiple missiles without using any information on time-to-go or its estimation. In the first guidance phase, this study integrates the cooperative guidance problem, state tracking problem and field-of-view constraint into a unified optimal control framework. A non-quadratic field-of-view constraint cost is innovatively constructed via an inverse optimal control approach, which leads to an analytical, distributed, and optimal guidance law to provide favored initial conditions for the latter phase. The stability and optimality of the closed-loop system is also proven. With respect to the second guidance phase, missiles disconnect from each other and are governed by the typical pure proportional navigation guidance law to eventually realize the salvo attack mission. Numerical simulations are performed to validate the theoretical results.
       
  • Interaction of a plane shock wave with an area of ionization instability
           of discharge plasma in air
    • Abstract: Publication date: February 2019Source: Aerospace Science and Technology, Volume 85Author(s): T.A. Lapushkina, A.V. Erofeev, O.A. Azarova, O.V. Kravchenko The conducted study is devoted to the research area of control of high speed flows and shock-wave configurations. Experimental and numerical results on the interaction of a plane shock wave (M = 5–6) with the area formed by ionization instability in gas discharge plasma at a pressure of 3–7 Torr and gas-discharge current of 100–400 mA are presented. The region of ionization instability in plasma of high degree of ionization and non-equilibrium (gas temperature T ∼ 300–600 K, electron temperature Te∼5⋅103 K) has been obtained experimentally in air. This instability is characterized by spherical strata of changing shape. Two types of a shape of the ionization instability have been obtained: large-scale type and small-scale type. Also, the transitional type from large-scaled strata to low-scaled ones has been obtained. The influence of the ionization instability was established to cause the distortion of an initially plane shock wave up the complete disappearance of its front. Numerical modeling the interaction of the initially plane shock wave with ionization-unstable plasma using the Euler system of equations has been conducted. The area formed by ionization unstable plasma was modeled via a set of thermal layers with varying densities (and temperatures). Changing the physical–chemical properties of the gas medium was taken into account varying the ratio of specific heats. Layered structures with wavy shock-wave and contact discontinuity fronts have been observed. The multiple generations of shear layer instabilities and the Richtmyer–Meshkov instabilities have been obtained as the result of the interaction of a shock wave with thermal strata. It has confirmed that the shapes of the both discontinuities change acquiring an unstable character. Examples of the flow modes with the disappearance of the shock wave front causing by the origination of a structure with cellular order of shear layer instabilities have been presented. A good agreement at the qualitative level with the experimental Schlieren images has been demonstrated.
       
  • Nonlinear vibration of metal foam cylindrical shells reinforced with
           graphene platelets
    • Abstract: Publication date: February 2019Source: Aerospace Science and Technology, Volume 85Author(s): Yan Qing Wang, Chao Ye, Jean W. Zu This paper performs nonlinear vibration analysis of metal foam circular cylindrical shells reinforced with graphene platelets. An improved Donnell nonlinear shell theory is employed to formulate the present model. The graphene platelet reinforced material properties are evaluated by the Halpin–Tsai equation. Different types of porosity and graphene platelet (GPL) distribution are taken into account. Governing equations are derived via Hamilton's principle and then they are transformed to ordinary differential equations using the Galerkin method. Afterwards, nonlinear frequencies of the system are solved by using the multiple scale method. Our findings demonstrate that GPL reinforced metal foam (GPLRMF) shells exhibit hardening-spring vibration characteristics. The nonlinear to linear frequency ratio of the shell closely relates to the porosity distributions and GPL patterns. The effect of geometrical size of graphene platelets on nonlinear vibration characteristics of GPLRMF cylindrical shells is also highlighted.
       
  • Control-oriented unsteady one-dimensional model for a hydrocarbon
           regeneratively-cooled scramjet engine
    • Abstract: Publication date: February 2019Source: Aerospace Science and Technology, Volume 85Author(s): Jicheng Ma, Juntao Chang, Junlong Zhang, Wen Bao, Daren Yu A control-oriented unsteady 1-D model for regeneratively-cooled scramjet engine which is composed of flow-combustion module, regenerative cooling module, is presented in this paper. It can reflect unsteady process in real-world where flow-combustion module couples with regenerative cooling module at each time step. The coupled 1-D model is quantitatively validated under steady and unsteady conditions. By the analyses of steady characteristics using this coupled 1-D model, it can be concluded that the allowable maximum first stage fuel equivalence ratio of regeneratively-cooled scramjet engine is lower and thrust is larger than no-cooled scramjet engine for the same equivalence ratio under the same flight Mach number. The difference of overtemperature boundary between regeneratively-cooled and no-cooled scramjet engine is also pointed. Moreover, regenerative cooling process results in moving toward less of fuel equivalence ratio of combustion mode transition. Unsteady characteristics analyses are investigated subsequently. Results indicate that there exists thermal inertia after kerosene flowing in cooling channels so that the flow rate of kerosene of entering into combustor cannot reflect kerosene changing signal in time. It leads to that thrust response time of regeneratively-cooled scramjet engine is longer than that of no-cooled scramjet engine. Some special attentions should be paid that combustion mode transition can be induced and affected by regenerative cooling process in some conditions. As a whole, the low computation cost makes this coupled 1-D model suitable for control system design and overall performance evaluation for regeneratively-cooled scramjet engine.
       
  • Ice accretion and aerodynamic effects on a multi-element airfoil under SLD
           icing conditions
    • Abstract: Publication date: February 2019Source: Aerospace Science and Technology, Volume 85Author(s): L. Prince Raj, J.W. Lee, R.S. Myong The impingement behavior of large water droplets, their interactions with the solid wall and the subsequent ice accretion and aerodynamic effects have become a key issue in in-flight aircraft icing. In this study, ice accretion and aerodynamic effects on a multi-element airfoil were investigated under the recently introduced Appendix O icing envelope. Supercooled large droplet (SLD) dynamics were taken into account by employing a unified computational approach. Ice accretion was simulated using a partial differential equation (PDE) based solver, instead of the commonly used control volume method. The numerical solver of the SLD impingement was built on the droplet deformation and droplet–wall interaction splash models. The unified solvers for clean air, large droplet impingement, ice accretion, and the aerodynamic analysis of ice effects—all of which are based on a single unstructured upwind finite volume framework—were first validated using available experimental data and then applied to investigate ice accretion and the resulting aerodynamic effects on multi-element airfoils for various flight conditions and, in particular, near-freezing SLD icing conditions. Interestingly, two counter-intuitive results were found when comparing the ice accretion and associated aerodynamic degradation for non-SLD and SLD cases. Moreover, considering runback ice was shown to be essential in the design of an ice protection system (IPS) for the multi-element wing.
       
  • Simulation of helicopter ditching using smoothed particle hydrodynamics
    • Abstract: Publication date: February 2019Source: Aerospace Science and Technology, Volume 85Author(s): Mark A. Woodgate, George N. Barakos, Nigel Scrase, Tim Neville This paper explores the potential use of smoothed particle hydrodynamics methods for helicopter ditching. The method appears suitable for this task since it is mesh-free and can accommodate the interaction between a floating object and the free-surface of water. Simple cases of objects dropped on water were first studied to establish confidence on the method, and quantify the effect of the numerical parameters of SPH including the boundary condition between the water and solid, the effect of the number and type of smoothed particles as well as the generation of different sea-states for the ditching. Once confidence on the method was established, experiments for the ditching of a model-scale helicopter were used for validation. The smoothed particle hydrodynamics method provides good agreement with experiential data for the position and velocity of the helicopter fuselage.
       
  • A method to minimize stage-by-stage initial unbalance in the aero engine
           assembly of multistage rotors
    • Abstract: Publication date: February 2019Source: Aerospace Science and Technology, Volume 85Author(s): Yongmeng Liu, Maowei Zhang, Chuanzhi Sun, Ming Hu, Danyang Chen, Zewei Liu, Jiubin Tan The paper proposes a method to minimize stage-by-stage initial unbalance in the aero engine assembly of multistage rotors based on the connective assembly model. The process of the mass eccentric deviation propagation in the assembly is analyzed. The initial unbalance of the final assembly is improved by properly selecting the assembly orientations of multistage rotors. A constrained nonlinear programming model is extracted from the optimal assembly strategy by choosing the initial unbalance as the objective function, and choosing the assembly orientations as the nonlinear constraint. The globally optimized solution of the constrained nonlinear programming model is solved using a genetic algorithm. The proposed method is used on an experimental set-up with the horizontal balancing machine and the effectiveness of the proposed method is verified by the assembly of the multistage rotors using the optimal assembly strategy. Compared to the direct assembly strategy, which the assembly orientations without consideration, the initial unbalance of final assembly using the optimal assembly strategy are reduced by 45.8%, 63.1% and 72.7% for two, three and four rotors assembly, respectively. The proposed method can improve the assembly quality of multistage rotors in the aero engine assembly and be used for assembly guidance, tolerance allocation, and so on, especially for the assembly with a large number of rotors.
       
  • Unscented Kalman filters for range-only cooperative localization of swarms
           of munitions in three-dimensional flight
    • Abstract: Publication date: February 2019Source: Aerospace Science and Technology, Volume 85Author(s): Bradley T. Burchett Range-only cooperative localization for formations of projectiles is investigated using extended and unscented Kalman filters. Observability for range only cooperative localization is reviewed. Consequences of observability allow the user to choose many features of the localization scheme including number of measurements, ground references, and required on-board sensing. Several Kalman filters are designed including centralized filters using 3 degree of freedom plant models using both extended and unscented schemes, and decentralized unscented Kalman filters using 3 degree of freedom plant models. Results are shown for several scenarios to illustrate the effectiveness of each estimation scheme.
       
  • Robust optimization of impulsive orbit transfers under actuation
           uncertainties
    • Abstract: Publication date: February 2019Source: Aerospace Science and Technology, Volume 85Author(s): Mohammad Sadegh Mohammadi, Abolghasem Naghash Having some initial and final conditions, it is possible to find a solution for an impulsive orbit transfer, using various trajectory optimization methods. However, there is no guarantee for solution to correspond to the real world condition since there are many possible uncertainties like unpredictable malfunctions in actuation of thrusters that may cause mission failure. The approach of this paper is to develop a method to find a trajectory less sensitive to uncertainties and the motivation is to make a good tradeoff between optimization accuracy and run time. In this paper, novel methods are introduced to make the optimization fast, globally optimized, constraint free and full perturbed. In addition, there is no need to guess an initial guess for the solution because of our use of genetic algorithm. The presented method has two parts for each step of optimization. First part calculates an injection guaranteed trajectory using Lambert's method and second part handles uncertainties over this trajectory. Monte-Carlo method is used for sampling. To reduce the computational cost, a surrogate model is calculated and used in each step of optimization. The result of this robust optimization is a trajectory in which uncertainties make less injection error variance.
       
  • Integrated relative position and attitude control for spacecraft
           rendezvous with ISS and finite-time convergence
    • Abstract: Publication date: February 2019Source: Aerospace Science and Technology, Volume 85Author(s): Yu Wang, Haibo Ji This article is concerned with the relative motion control for spacecraft rendezvous in the presence of external disturbance. First, to ensure accurate relative position tracking and attitude synchronization between a pursuing spacecraft and a target spacecraft, a novel integrated six-degrees of freedom (6-DOF) dynamics model is established with consideration of thrust layout. In particular, based on line of sight coordinate frame, the relative position dynamics is more suitable for actual measurement circumstances and applicable for arbitrary orbital forms. Subsequently, to implement the rendezvous mission, two groups of control schemes are designed using backstepping method with input-to-state stable property and finite-time control technique, respectively. In Lyapunov theoretical framework, the designed controllers are proved to guarantee the convergence of relative position and relative attitude tracking errors. Finally, simulations results are carried out to validate the effectiveness of the proposed control approaches.
       
  • Gas breakdown mitigation in satellite slip rings
    • Abstract: Publication date: February 2019Source: Aerospace Science and Technology, Volume 85Author(s): F. Avino, P. Martens, A.A. Howling, D. Bommottet, I. Furno Direct current (DC) gas breakdown is experimentally investigated for high-voltage circular conductors and insulators that reproduce the main features of a satellite slip ring. Breakdown voltages are measured in the pressure range 10−3–10 mbar, where gas breakdown is predominant with respect to vacuum breakdown. The measured breakdown curves show clear similarities with Paschen's curve, which is generally associated with parallel plate electrodes. It is shown that the low-pressure branch of the measured curves is determined by breakdown between the high-voltage ring and the grounded vacuum chamber, whereas the high-pressure branch is due to discharges between the high-voltage ring and the adjacent grounded rings. A technical solution is introduced to inhibit the gas discharges at low-pressures in a slip ring assembly: The diameter of the grounded conducting discs is extended, strongly increasing the measured breakdown voltages by modifying the electric field distribution. The safe operating pressure range of the satellite slip ring is thereby increased by two orders of magnitude.
       
  • Kriging surrogate model applied in the mechanism study of tip leakage flow
           control in turbine cascade by multiple DBD plasma actuators
    • Abstract: Publication date: February 2019Source: Aerospace Science and Technology, Volume 85Author(s): Jianyang Yu, Zhao Wang, Fu Chen, Jianing Yu, Cong Wang The dielectric barrier discharge (DBD) plasma actuator, in which electrodes are asymmetric arranged, has already demonstrated its ability in flow control. In the present work, the configuration of multiple plasma actuators is placed at the suction side of the cascade top to realize the tip leakage control. However, massive configurations appear when the number of plasma actuators increases, resulting in the investigation of actuator configuration for tip leakage flow control becomes a challenge. The surrogate modeling approach provides a cheap and efficient method for investigating the effect of multiple plasma actuators on the tip leakage flow control. By constructing an approximation model, tip leakage mass flow rates of all configurations are obtained in the present work. What's more, the flow structures in the tip clearance controlled by the plasma actuators are explained in the process of topological analysis. The results show that the tip leakage mass flow rate is decreasing with the number of active plasma actuators increasing. However, the decreasing would reach its limits in the process of adding plasma actuators. In the analysis of flow topology, single actuator would generate a small vortex at the suction side to cause an obstacle in the tip leakage flow. While the continuous arrangements of plasma actuator is beneficial to generate a larger obstacle in the tip clearance, which resulting in a significant reduction of m˙CFD and the promotion of Cp.
       
  • Robust trajectory tracking controller for quadrotor helicopter based on a
           novel composite control scheme
    • Abstract: Publication date: February 2019Source: Aerospace Science and Technology, Volume 85Author(s): Juqian Zhang, Dawei Gu, Zhaohui Ren, Bangchun Wen Robust trajectory tracking problem is dealt with for the under-actuated quadrotor system, which is decomposed into two subsystems, i.e., the inner-loop rotational subsystem perturbed by the time-varying disturbances (e.g., parameter variation, model mismatches, payload changes, wind gusts), and the outer-loop translational subsystem subjected to both the uncertain disturbances and time-varying measurement-delay simultaneously. The novel composite control scheme integrating a nominal controller with robust compensators is hierarchically developed. Composite observers consisting of time-varying disturbance observer (TDO) and time-varying measurement-delay observer (TMO) are incorporated into the nominal backstepping control (NBC) as robust compensators, to compensate for the time-varying uncertainties by estimating the disturbance dynamics and system states. Stability of the closed-loop control system is analytically proved based on the Lyapunov–Krasovskii theorem, and the proposed controller theoretically guarantees that the tracking error converges to a small neighborhood around the origin. Comparative flight performances under different controllers for the quadrotor helicopter, in the presence of uncertain disturbances and time-varying measurement-delay, are presented to demonstrate the effectiveness and superiority of the proposed controller.
       
  • Pressure characteristics of a ram-RDE diffuser
    • Abstract: Publication date: February 2019Source: Aerospace Science and Technology, Volume 85Author(s): Douglas A. Schwer, K. Kailasanath, Tom Kaemming This work focuses on dynamic simulations of a representative ram RDE subsonic diffuser to study ram RDE compatibility with a supersonic inlet. Of particular interest is the propagation of pressure perturbations due to the detonation waves in the RDE combustion chamber. Three-dimensional CFD simulations were completed using a notional diffuser for a ram RDE device operating at 42,000 ft and Mach 2.5. A previous RDE code was used to provide an exhaust forcing function assuming a slot injector plate. A total of 8 different exhaust forcing functions were used, representing RDE air inlet area ratios between 0.3 and 0.6 in single and dual wave operation. Simulations were then completed for the diffuser with the Propel code for each of these exhaust forcing functions. Analysis indicates that the amount of pressure variation seen near the diffuser inlet is very sensitive to the shape of the forcing function at the diffuser exit. The broad pressure rise of the lower area ratios tend to remain intact through the diffuser to the intake with minimal dissipation, while the sharper pressure peak of the higher area ratios tend to be substantially dissipated. For single-wave operation, the broad pressure rise yields pressure variations at the intake that vary from 20.5% (area ratio of 0.3) to 39.1% (a=0.6). For the dual-wave operation, the sharper peak of the pressure rise exhibited even more dissipation. Thus, for dual-wave operation, the amount of pressure variation at the intake varies from 21.0% (a=0.3) to 21.9% (a=0.6) for the different area ratios.
       
  • Experimental investigation of influence factors on flame holding in a
           supersonic combustor
    • Abstract: Publication date: February 2019Source: Aerospace Science and Technology, Volume 85Author(s): Yanhua Wang, Wenyan Song In this paper, combustion and flame stability are studied experimentally in a kerosene fueled dual-mode scramjet combustor at various air and fuel conditions. The lean and rich blowout limits of liquid kerosene injected from the wall upstream of a cavity flameholder are examined in Mach 2.0 airflow. The effects of injector configurations on the blowout limit are investigated. It is found that injector configuration has significant effect on the flame stabilization. The lean blowout limit is much lower when using injector C, which has a configuration of 2×Φ0.35 mm. But for the rich blowout limit, injector B which has a configuration of 3×Φ0.35 mm performs better than others. Therefore, the diameter and distribution of fuel injection holes should be considered fully in combustor design. Furthermore, the influence on blowout limits of combustor entrance total temperature and total pressure are also investigated. As the total temperature increasing, the equivalence ratio range between rich and lean blowout limit become larger. While, compared with the total temperature, the effect of total pressure on the blowout limit is slighter.
       
  • Roughness effect on shock wave boundary layer interaction area in
           compressor fan blades passage
    • Abstract: Publication date: February 2019Source: Aerospace Science and Technology, Volume 85Author(s): Ryszard Szwaba, Piotr Kaczynski, Piotr Doerffer The main objective of this research is to study experimentally the effect of transition on the flow structure of the shock wave boundary layer interaction in the blades passage of a compressor cascade. For this purpose a model of a turbine compressor passage was designed and assembled in a transonic wind tunnel. In the experiment the distributed roughness with different heights and locations was used to induce transition upstream of the shock wave. Three locations of distributed roughness in form of standard sandpaper strips were chosen to apply on the blade. The average roughness height comprised in a range of 8 to 16 μm. All together 8 different flow cases were investigated. The paper focuses on the influence of the boundary layer transition induced by different roughness values and locations on the flow pattern in the blades passage of a compressor cascade. Very promising results were obtained in the roughness application for the boundary layer transition control, demonstrating a positive effect in changing the nature of the interaction.
       
  • Trim investigation for coaxial rigid rotor helicopters using an improved
           aerodynamic interference model
    • Abstract: Publication date: February 2019Source: Aerospace Science and Technology, Volume 85Author(s): Ye Yuan, Renliang Chen, Pan Li The coaxial rigid rotor helicopter has been proposed as a future high-performance rotorcraft concept. However, the aerodynamic interference of this helicopter is complicated because it couples with the unique flapping feature of the rigid rotor, which further alters the trim characteristics of coaxial rigid rotor helicopters. Thus, a multi-point vortex ring element (MVRE) model is developed to simulate the aerodynamic interference between rotors. The method for establishing this MVRE model is illustrated, and a wind tunnel experimental dataset is used to assess its precision hover and forward flight states. Next, a flight dynamics model of the coaxial rigid rotor helicopter is built based on the MVRE aerodynamic interference model and the flapping feature of the rigid rotor. The influence of the rotor wake on the fuselage and the horizontal and vertical tails can also be calculated using this model. The trim characteristics of this helicopter are evaluated with flight test data for speeds ranging from 0 m/s to 80 m/s, and the results confirm that this model can reflect the trim characteristics with satisfactory precision. In addition, the calculation process demonstrates that the MVRE model provides a much faster computing rate. Considering the aerodynamic interference and rigid rotor characteristics, the trim results of the coaxial rigid rotor helicopter present unique features: aerodynamic interference in the coaxial rotor system not only increases the collective pitch and the collective differential but also adds a negative gradient under the forward speed in the longitudinal cyclic pitch in the low-speed forward flight range. Moreover, the rotor wake effect on the other parts of the helicopter is distinct from the corresponding effects on conventional helicopters in terms of the trim characteristics.
       
  • Capturing transition with flow–structure–adaptive KDO RANS
           model
    • Abstract: Publication date: February 2019Source: Aerospace Science and Technology, Volume 85Author(s): Jinglei Xu, Ding Xu, Yang Zhang, Junqiang Bai By extending Bradshaw's assumption from the free shear flows to the wall-bounded flows, the Kinetic energy Dependent Only turbulence model (KDO) established a new Reynolds stress constitution. The Bradshaw function (τ12/k) and the coefficient of the dissipation term are the only two empirical coefficients. They were both calibrated with the turbulent Reynolds number, Rek=ρk1/2d/μ, which depends on the wall distance. Once the wall distance is replaced with “flow–structure–adaptive” parameters, such as the eddy viscosity ratio, r=μt/μ, the model could naturally capture various transition phenomena. The improved KDO model is assessed by some test cases including the classic bypass transition of T3A and T3B boundary layers, the natural transition of the T3A boundary layer, and the separation bubble induced transition of Aero-A airfoil. The assessments show that the improved KDO model is not capable of capturing the precise process of the laminar-turbulent flow transition, but the model can accurately predict the transition onset locations. The model does not include specific transition mechanisms; however, for high Reynolds numbers and complex flows with different types of transition, the predictions are reliable.
       
  • 3D shell model for the thermo-mechanical analysis of FGM structures via
           imposed and calculated temperature profiles
    • Abstract: Publication date: February 2019Source: Aerospace Science and Technology, Volume 85Author(s): S. Brischetto, R. Torre The exact three-dimensional (3D) shell model proposed in the present paper is able to perform the thermal stress analysis of simply-supported Functionally Graded Material (FGM) spherical and cylindrical shells, cylinders and plates. The model is based on the 3D equilibrium equations for spherical shells developed using an orthogonal mixed curvilinear coordinate system. The use of this reference system allows the investigation of cylindrical shells, cylinders and plates as particular cases of spherical shells by means of simple considerations on the radii of curvature. The 3D shell model uses a layer-wise approach and the exponential matrix method to calculate the general and the particular solutions through the thickness direction z. The system of second order differential equations in z is not homogeneous because of the thermal terms which are externally defined. The system is reduced to a group of first order differential equations in z simply redoubling the number of variables. The solution is in closed form in the in-plane directions α and β because of the hypotheses of simply-supported boundary conditions, harmonic forms for displacement and temperature fields, and isotropic behavior in the in-plane directions for functionally graded materials. In order to define the equivalent thermal load, the temperature profile through the thickness is separately defined by means of three possible ways. Using the hypothesis of temperature amplitudes imposed at the top and bottom external surfaces in steady-state conditions, the temperature profile can be: imposed as linear through the entire thickness direction, calculated by solving the 1D version of the Fourier heat conduction equation, or calculated by solving the 3D version of the Fourier heat conduction equation. The effects of different temperature profiles on the displacement and stress analyses of FGM plates and shells are here remarked. The first order differential equation system in z has not constant coefficients because of the presence of radii of curvature for shells and through-the-thickness variable elastic and thermal coefficients for the FGM layers. An appropriate number of mathematical layers is introduced to calculate the curvature influence for shells and the elastic and thermal material coefficients for FGM layers. Therefore, the system can be considered as differential equations with constant coefficients. The proposed results allow the evaluation of thickness ratio, geometry, lamination scheme, thickness material law and temperature profile effects in the related thermal stress analysis of single-layered and sandwich FGM plates, cylinders, spherical and cylindrical shells.
       
  • A hollow combustor that intensifies rotating detonation
    • Abstract: Publication date: February 2019Source: Aerospace Science and Technology, Volume 85Author(s): Yuhui Wang, Jialing Le Annular combustors are traditional, while hollow combustors without centerbodies are becoming popular for rotating detonation engines. Both schemes are studied experimentally in this paper to determine which combustor demonstrates a better detonation performance and pressure gain performance. The hydrogen and air sources are at room temperatures of 280–285 K. The annular and hollow combustors have the same scheme, although the centerbody is removed from the hollow combustor. The combustors are made optically accessible by embedding a piece of quartz glass in the outerbody. The hollow combustor channel has an outer diameter of 100 mm, and the detonation channel of the annular combustor has an inner diameter of 60 mm. Fuel and air are injected into the combustor from 150 cylindrical orifices with a diameter of 0.8 mm and a circular channel with a throat width of 1 mm, respectively. The results show that the hollow combustor demonstrates a better detonation performance and a higher pressure gain than the annular combustor. High-speed images show the rotating detonation is propagating with a lap time which is in good agreement with that obtained from pressure traces.
       
  • Stability enhancement using a new hybrid casing treatment in an axial flow
           compressor
    • Abstract: Publication date: February 2019Source: Aerospace Science and Technology, Volume 85Author(s): Jichao Li, Juan Du, Fan Li, Qianfeng Zhang, Hongwu Zhang A hybrid slot–groove casing treatment (CT) composed of front axial slots and rear circumferential grooves is experimentally studied in a low-speed axial compressor in consideration of stability and efficiency. Results show that the hybrid slot–groove CT can improve the stall margin by 19.79% with an efficiency penalty of 1.5%. Compared to full-groove and full-slot CTs, the hybrid slot–groove CT's stall margin improvement and minimization of efficiency loss are excellent. Detailed measurements indicate that, unlike smooth casing (SC), the hybrid slot–groove CT can unload the blade tip, weaken the tip leakage vortex, and improve the disturbance observed downstream to postpone stall. Given the influence of serious passage blockage derived from the hub region after improving tip flow capability under the hybrid slot–groove CT, the stall route and inception captured in the casing wall and different downstream radial positions demonstrate that, unlike the SC with a typical spike-type inception (short length-scale with 2–3 blade passages), the hybrid slot–groove CT creates a long length-scale stall inception (6–8 blade passages). In addition, a new type of instability inception with a different frequency band from the stall cell is found in the hub region under the hybrid slot–groove CT. This inception appears more than hundreds of revolutions before the fast trigger of spike-type inception and can be used as an early stall warning signal under high hub loading.
       
  • Preliminary aeroelastic design of composite wings subjected to critical
           gust loads
    • Abstract: Publication date: February 2019Source: Aerospace Science and Technology, Volume 85Author(s): D. Rajpal, E. Gillebaart, R. De Breuker Including a gust analysis in an optimization framework is computationally expensive as the critical load cases are not known a priori and hence a large number of points within the flight envelope have to be analyzed. Model order reduction techniques can provide significant improvement in computational efficiency of an aeroelastic analysis. In this paper, after thorough analysis of 4 commonly used model order reduction methods, balanced proper orthogonal decomposition is selected to reduce the aerodynamic system which is based on potential flow theory. The reduced aerodynamic system is coupled to a structural solver to obtain a reduced-order aeroelastic model. It is demonstrated that the dominant modes of the aerodynamic model can be assumed to be constant for varying equivalent airspeed and Mach number, enabling the use of a single reduced model for the entire flight envelope. A dynamic aeroelastic optimization method is then formulated using the reduced-order aeroelastic model. Results show that both dynamic and static loads play a role in optimization of the wing structure. Furthermore, the worst case gust loads change during the optimization process and it is important to identify the critical loads at every iteration in the optimization.
       
  • Adaptive control of underactuated flight vehicles with moving mass
    • Abstract: Publication date: February 2019Source: Aerospace Science and Technology, Volume 85Author(s): Jianqing Li, Sai Chen, Chaoyong Li, Changsheng Gao, Wuxing Jing The configuration of internal moving masses is a key challenge for applying moving mass control technology to flight vehicle control. A novel configuration with a large mass ratio moving mass and reaction jets is proposed for bank-to-turn control. The control system of the proposed configuration consists of the attitude dynamics and the moving mass dynamics, which are coupled by the additional inertia moment of moving mass. To deal with the coupling, the integrated control of an attitude-servo system and a lateral underactuated control based on immersion and invariance theory is presented. To overcome the uncertainties in the flight vehicle model, immersion and invariance theory is employed to design an estimator for the unknown aerodynamic parameters. The estimator has an additional nonlinear term which adjusts the performance of the estimation error. The simulation results show that the proposed attitude-servo controller for the longitudinal subsystem can enhance the response of the system and the underactuated controller of the lateral subsystem can reduce fuel consumption.
       
  • Dynamic response of axially functionally graded beam with
           longitudinal–transverse coupling effect
    • Abstract: Publication date: February 2019Source: Aerospace Science and Technology, Volume 85Author(s): Ke Xie, Yuewu Wang, Tairan Fu This study investigates the dynamic response of an axially functionally graded (AFG) beam with longitudinal–transverse coupling effect. The beam is acted by a moving transverse harmonic load and a moving longitudinal harmonic load simultaneously. Both the classical beam theory (CBT) and the Timoshenko beam theory (TBT) are employed and written in a unified form. The system of equations of motion is obtained using Lagrange's equations. The nonlinear formulations of the longitudinal–transverse coupling AFG beams are presented. These formulations are solved using the Newmark method in conjunction with a direct iteration procedure; subsequently, the free and forced dynamic behavior characteristics of the AFG beams are investigated. A comparison of the results obtained using the TBT and CBT is performed. The nonlinear effects caused by different excitation frequencies and amplitudes both in the transverse and longitudinal directions of the moving loads on the dynamic responses of the AFG beam are discussed. Further, different nonlinear dynamic behaviors of the AFG beam caused by different longitudinal–transverse coupling coefficients are revealed. The numerical results indicate that the longitudinal–transverse coupling effect plays an important role in the dynamic behavior of the AFG beam.
       
  • Optimization and control application of sensor placement in
           aeroservoelastic of UAV
    • Abstract: Publication date: February 2019Source: Aerospace Science and Technology, Volume 85Author(s): Weiqi Yang, Hui Yang, Shuo Tang In order to suppress aeroservoelastic in high-aspect-ratio flexible UAV, in the present work an advanced sensor placement criterion is developed using Cuckoo search algorithm in combination with an enhanced active control method. The advanced sensor placement criterion basically combines the vibration energy based observability measurement as well as further information on evaluating sensor influence in terms of H2 norm to balance the low and high frequency modes. By eliminating several nests with worst fitness values in each generation and using self-adaptive feedback scaling factor, the proposed elimination mechanism Cuckoo search (EMCS) algorithm is almost three times faster than the standard one. Subsequently, an enhanced active disturbance rejection control (ADRC) method is proposed for the first time in the active vibration control and wind load alleviation of flexible UAV. It is demonstrated that the enhanced ADRC/PID approach with obtained sensor locations can result in a 45.83% reduction in generalized vibrations energy and about 52.16% reduction in wind load alleviation when compared with designs where the sensor locations are not optimum. Finally, the simulation results show that the optimization algorithm can effectively find the optimal location of sensors. Meanwhile, the suppression of aeroservoelastic can be realized with the utilization of the active control.
       
  • A new continuous adaptive finite time guidance law against highly
           maneuvering targets
    • Abstract: Publication date: February 2019Source: Aerospace Science and Technology, Volume 85Author(s): Jianguo Guo, Yifei Li, Jun Zhou A novel continuous adaptive finite time guidance (CAFTG) law is proposed for homing missiles. Firstly, three-dimensional nonlinear dynamics describing the pursuit situation of the missile and the target are introduced to obtain the mathematical model of engagement. Secondly, in order to improve the accuracy of interception, a nonlinear disturbance observer with finite time convergence is employed to estimate the acceleration of a target and compensate the guidance law. A continuous guidance scheme with robustness is constructed via sliding mode control theory, which guarantees finite time convergence by Lyapunov stability theory. Finally, simulations are conducted on the nonlinear dynamic models and results demonstrate the effectiveness of proposed guidance method.
       
  • Flow simulation and drag decomposition study of N3-X hybrid wing-body
           configuration
    • Abstract: Publication date: February 2019Source: Aerospace Science and Technology, Volume 85Author(s): Hyoungjin Kim, May-Fun Liou Flow simulation and drag decomposition study was conducted for the flow field around the clean airframe of the N3-X hybrid wing body configuration at a cruise flight condition. Adopted flow solver was GO-flow; an unstructured hybrid mesh cell-vertex finite volume RANS solver. The OVERFLOW code was also used to cross check the results. We have found that the major flow features such as surface pressure distributions and shock waves from the two flow solvers were almost identical to each other. Evaluation of drag force using near- and mid-field approaches was also conducted. Comparisons of various drag components between two tested CFD codes are examined with various grid resolutions for angle of attack sweep at the cruise condition. From the results, it is recommended for the computational mesh of the unstructured CFD code to have anisotropic (stretched) meshes near the leading and trailing edges and enough resolution in wake regions for accurate estimation of the drag coefficient. Nacelle installation effects were also tested using inviscid simulations by flow-through single-passage nacelle and mail slot nacelle, and it was found that fan suction effects should be properly considered for the accurate modeling of the nacelle installation effects. Lessons learned in this study enhance confidence in using CFD based integrative approach to encompass design of future hybrid wing body aircraft.
       
  • Control of corner separation via dimpled surface for a highly loaded
           compressor cascade under different inlet Mach number
    • Abstract: Publication date: February 2019Source: Aerospace Science and Technology, Volume 85Author(s): Huawei Lu, Yi Yang, Shuang Guo, Wenxuan Pang, Fan Yang, Jingjun Zhong In this paper, flow characteristics of dimpled compressor cascades are numerically discerned by solving the compressible Reynolds-averaged Navier–Stokes equations on structured grids over the inlet Mach number ranging from 0.3 to 0.8. Four rows of parallel, spherical dimples with a depth of 0.2 mm and a depth-to-diameter ratio of 0.25 are embedded along 10%–32% and 38%–60% chord of suction surface respectively. Results showed that two dimple configurations can both reduce the total pressure loss in all researched Mach number except that of 0.8. The suppression of three-dimensional separation near the hub-corner region is the primary cause of loss reduction. What's more, the separation bubble on the suction surface can be also eliminated or depressed due to the disturbance and higher level of turbulence kinetic energy within boundary layer. According to the distribution of axial vorticity in flow passage, the dimpled vortex can also suppress the migration of secondary flow along spanwise direction under a relative higher cross pressure gradient.
       
  • Fuzzy logic based equivalent consumption optimization of a hybrid electric
           propulsion system for unmanned aerial vehicles
    • Abstract: Publication date: February 2019Source: Aerospace Science and Technology, Volume 85Author(s): Ye Xie, Al Savvaris, Antonios Tsourdos This paper presents an energy management strategy for a hybrid electric propulsion system designed for unmanned aerial vehicles. The proposed method combines the Equivalent Consumption Minimization Strategy (ECMS) and fuzzy logic control, thereby being named Fuzzy based ECMS (F-ECMS). F-ECMS can solve the issue that the conventional ECMS cannot sustain the battery state-of-charge for on-line applications. Furthermore, F-ECMS considers the aircraft safety and guarantees the aircraft landing using the remaining electrical energy if the engine fails. The main contribution of the paper is to solve the deficiencies of ECMS and take into consideration the aircraft safely landing, by implementing F-ECMS. Compared with the combustion propulsion system, the hybrid propulsion system with F-ECMS at least reduces 11% fuel consumption for designed flight missions. The advantages of F-ECMS are further investigated by comparison with the conventional ECMS, dynamic programming and adaptive ECMS. In contrast with ECMS and dynamic programming, F-ECMS can accomplish a balance between sustaining the battery state-of-charge and electric energy consumption. F-ECMS is also superior to the adaptive ECMS because there are less fuel consumption and lower computational cost.
       
  • Recurrence network analysis for uncovering dynamic transition of
           thermo-acoustic instability of supercritical hydrocarbon fuel flow
    • Abstract: Publication date: February 2019Source: Aerospace Science and Technology, Volume 85Author(s): Hao Zan, Weixing Zhou, Xuefeng Xiao, Long Lin, Junlong Zhang, Haowei Li Characterizing the transition of thermo-acoustic instability of supercritical hydrocarbon fuel flow is a fundamental problem eliciting a great deal of attention from different disciplines. We experimentally and theoretically investigate the transition process between thermo-acoustic stability and instability. The method of recurrence network is applied to analyze the pressure time series of supercritical hydrocarbon fuel flow. As a result, we can distinguish a thermo-acoustic transition process from normal signals by real-time detecting the complexity variation of pressure signals of fuel in cooling channels with the recurrence network method. Then, we construct the recurrence network from experimental data under the stable, transition and unstable states, and investigate the degree distribution and motifs distribution. We find that the degree distribution and motifs distribution allow quantitatively uncovering the complexity dynamic process. We investigate the during time of transition process, and find that the mass flow rate and the inlet pressure will make an influence on the transition time. These findings present a first step towards an improved understanding on the transition of thermo-acoustic instability from a complex network perspective. Moreover, the investigation on the transition of thermo-acoustic instability under supercritical pressure condition offers guidance on the control of scramjet fuel supply, which can secure stable fuel flowing in regenerative cooling system.
       
 
 
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