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Aerospace Science and Technology
Journal Prestige (SJR): 0.796
Citation Impact (citeScore): 3
Number of Followers: 397  
  Hybrid Journal Hybrid journal (It can contain Open Access articles)
ISSN (Print) 1270-9638 - ISSN (Online) 1270-9638
Published by Elsevier Homepage  [3206 journals]
  • Experimental investigation of inclining the upstream wall of a scramjet
    • Abstract: Publication date: April 2020Source: Aerospace Science and Technology, Volume 99Author(s): Matthew A. Trudgian, Will O. Landsberg, Ananthanarayanan VeeraragavanAbstractThe cavity flameholder has been widely used in scramjets to improve mixing and combustion rates. While many have investigated modifications to the cavity downstream-wall, few have investigated modifying the upstream-wall - a rear-facing step. The rapid expansion produces a low pressure region and base drag force, while only the larger of the induced twin vortices contributes substantially to mixing. In this paper, the upstream-wall of a scramjet combustor cavity was inclined, aiming to reduce cavity base drag and remove superfluous cavity vortex structures. Upstream-wall angles of 90°, 45° and 22.5° were examined, at cavity length-to-depth (L/D) ratios from 4 to 7. Cavity downstream-wall angles of 90° and 22.5° were examined for each configuration. Reflected shock tunnel experiments were conducted at Mach 7 enthalpy, scramjet combustor conditions. Experimental Schlieren imaging showed that reducing the upstream-wall angle did not significantly affect shear layer separation for L/D ratios of 4 and 5. At L/D ratios 6 and 7, however, reducing the upstream-wall angle saw the shear layer penetrating deeper into the cavity. Reynolds-averaged Navier-Stokes computations examined the internal flow structure and it was shown that reducing the upstream-wall angle to 45° retained the dominant mixing vortex in the cavity, with base drag reduced by 21% compared to the standard 90° upstream-wall case.
  • Optimal control of turbulent premixed combustion instability with annular
           micropore air jets
    • Abstract: Publication date: March 2020Source: Aerospace Science and Technology, Volume 98Author(s): Hao Zhou, Chengfei Tao, Zihua Liu, Sheng Meng, Kefa CenAbstractThis article experimentally investigates the suppressing of combustion instability in a model gas turbine combustor with annular air jets. Three parameters were optimized to find the control effectiveness—variations of the air injection flow rate, number, and diameter of jet nozzles. Results indicate that the suppression ratio of the self-excited oscillations can reach 95% with optimal control, the transition and saturation region of control effectiveness are identified, and the four jet nozzles design case can lead to the best effect of damping. Besides, mode shifting of flame heat release rate is observed, the amplitude of heat release rate increases to 2∼4 times than that of the uncontrolled flame, the main resonance frequency at 264.5 Hz was eliminated and triggered to a new oscillation frequency near 110 Hz. Moreover, after air injection control, the flame length and temperature changes inversely proportional to the air injection flow rate, as the air injection transforms the flow field and improves the turbulent combustion velocity. This study not only explored the mechanism behind combustion instability control under annular micropore air jets but also proposed the application of air injection as a practical tool for damping of combustion instability, which could contribute to the prevention of potential thermoacoustic instabilities in gas turbines.
  • A globally fixed-time solution of distributed formation control for
           multiple hypersonic gliding vehicles
    • Abstract: Publication date: March 2020Source: Aerospace Science and Technology, Volume 98Author(s): Yao Zhang, Xiao Wang, Shengjing TangAbstractThis paper investigates the distributed formation control problem for multiple hypersonic gliding vehicles (HGVs) under direct communication topology. Based on a formation flight framework of multiple HGVs, a globally fixed-time formation control scheme is proposed by using the fixed-time stability and the hierarchical control theory. Firstly, a fixed-time consensus law is addressed in the three-dimensional formation flight dynamics to generate the virtual control inputs. In order to design appropriate actual control inputs to track the virtual control commands rapidly and accurately, a composite fixed-time feedback control law is then put forward, where a new fixed-time tracking differentiator is designed to obtain the information on the time derivatives of virtual control inputs. Following, the problem of “explosion of complexity” can be resolved effectively. Finally, the globally fixed-time convergence is achieved in spite of the initial conditions of HGV's states, where the upper bound of the settling time can be estimated in advance by only using the control parameters and the direct communication topology. The desired formation configuration can be established within a prescribed convergence time. Extensive numerical simulations are performed to demonstrate the effectiveness and superiority of the proposed formation control strategy.
  • Characteristics of reaction zone in a dual-mode scramjet combustor during
           mode transitions
    • Abstract: Publication date: Available online 14 February 2020Source: Aerospace Science and TechnologyAuthor(s): Wubingyi Shena, Yue Huang, Yancheng You, Lizhe YiAbstractThe mechanism of controllable combustion modes (scramjet and ramjet) is one of the most significant challenges for dual-mode ramjet/scramjet engines. The detailed characteristics of the reaction zone during the transitions of combustion modes within the Hyshot II combustor are numerically investigated with 7 species-7 steps hydrogen/air reaction mechanism and the thickened flame model (TFM). In the case of continuous increasing or decreasing equivalence ratios (ERs), the critical ERs for the mode transitions are 0.474 or 0.539, respectively. As the ER increases or decreases, the location of the leading shock of the combustor shock-train moves upstream or downstream, respectively. When the ER continuously increases, the pressure ratio at the leading shock to the outlet of the isolator is greater than when the ER continuously decreases. Inversely, the mass-weighted average Mach number inside the combustor for a continuously increasing ER is less than that for the continuously decreasing ER. In the scramjet mode, the primary combustion reaction zone is located at the rear of the combustor. In the ramjet mode, combustion occurs in the separation zone at the middle of the combustion chamber, while the reaction zone is closer to equilibrium at the critical ER of 0.474 as compared to the scramjet mode.
  • Optimal guidance laws with prescribed degree of stability
    • Abstract: Publication date: Available online 14 February 2020Source: Aerospace Science and TechnologyAuthor(s): Ilan Rusnak, Haim Weiss, Gyorgy HexnerAbstractThe prescribed degree of stability criterion is used. This quadratic criterion involves an increasing exponential time dependent term in the integral part of the criterion. This criterion is used for derivation of guidance laws. The derived guidance law has the classical structure of guidance gain times the zero-effort miss. The important issue is the fact that initially the guidance gain and thus the commanded acceleration are larger than in the conventional Proportional Navigation guidance law, but near the end, the commanded acceleration is smaller. The new guidance law attempts to close the zero effort miss earlier in the scenario than the conventional guidance law.
  • Attitude control of space solar power satellite with large range of mutual
           motion among subsystems
    • Abstract: Publication date: Available online 14 February 2020Source: Aerospace Science and TechnologyAuthor(s): Xiangfei Ji, Yiqun Zhang, Guanheng Fan, Meng Li, Xianli LiAbstractThe transmitting antenna needs to be oriented to the Earth, while the PV cell array of the SSPS-OMEGA system rotates in a large range inside the concentrator for the orientation to the Sun, so that there is a large range of unfixed mutual movement among the subsystems. For the study of dynamic characteristics and attitude control problem of the space solar power satellite, the multi-body dynamics model of the system is built, and the gravity gradient torque equation applied on the PV cell array is derived in the process of mutual motion. A simple attitude control system and the corresponding electric thrusters configuration scheme are designed. The simulation results show that the attitude of 2-GW SSPS-OMEGA in roll and yaw direction is well controlled. However, due to the large range rotation of the PV cell array, the attitude in the pitch direction is controlled to converge to the microwave beam pointing accuracy of mechanical adjustment in longer time with the residual gravity gradient torque. In the long period, the attitude deviation angle in the pitch direction would stabilize periodically within ±0.018 degree. This provides a reference for further study of beam-control of the transmitting antenna.
  • Dynamic analysis of a hyper-redundant space manipulator with a complex
           rope network
    • Abstract: Publication date: Available online 13 February 2020Source: Aerospace Science and TechnologyAuthor(s): Shuguang Ma, Bin Liang, Tianshu WangAbstractThis paper studies the dynamics of a space manipulator. The space manipulator is designed for precise on-orbit servicing missions in a highly constrained environment. The concerned manipulator has hyper-redundant degrees of freedom and moves with a piecewise constant curvature, which enhances the flexibility and controllability. Such manipulator consists of a large number of links and a complex rope network. When the manipulator is driven, the interacting forces between the links and ropes introduce complexity into the dynamic behavior. In terms of dynamic modeling, the manipulator is a very complex system. This paper proposes a dynamic model of the manipulator based on methods of multibody dynamics. The ropes are assumed to be massless and linear elastic. The equations of motion are derived using space operator algebra. The vibration of the manipulator is investigated. The governing equations of the vibration are derived by applying the perturbation method to the proposed dynamic model. The values of the natural frequencies are investigated for the elasticities of the ropes. The proposed dynamic model is also applied in numerical simulation. The explicit fourth-order Runge-Kutta method is utilized to solve the equations of motion numerically. In numerical simulation, an upper bound of the time step is encountered. The value of upper bound is found to be related to the elasticities of the ropes. Such phenomena are studied by analyzing the stability region of the Runge-Kutta methods. Besides, the computational efficiency of the numerical simulation is also limited by the value of upper bound. Two modifications of the dynamic model are introduced to relax the upper bound of the time step. The effects of the modifications are demonstrated by numerical results.
  • Anti-disturbance attitude control of flexible spacecraft with quantized
    • Abstract: Publication date: April 2020Source: Aerospace Science and Technology, Volume 99Author(s): Linlin Hou, Haibin SunAbstractIn this paper, the problem of anti-disturbance control for flexible spacecraft with quantized states is investigated. By using quantized states, an extended state observer is constructed to estimate the flexible modal and external disturbances. Then based on the output of extended state observer and quantized states, an anti-disturbance controller is designed via recursive design method. Afterwards, the Lyapunov stability analysis is established to exhibit the boundedness of the closed-loop signals of the spacecraft systems. The merit and effectiveness of the obtained results are verified via simulation comparison.
  • Modeling of a 7-elements GOX/GCH4 combustion chamber using RANS with
           Eddy-Dissipation Concept model
    • Abstract: Publication date: April 2020Source: Aerospace Science and Technology, Volume 99Author(s): JianFei Wei, Mai Ye, Silong Zhang, Jiang Qin, Oskar J. HaidnAbstractA 7-element rocket combustion chamber using GOX/GCH4 as propellant has been modeled and simulated to get a more comprehensive knowledge for combustion and heat transfer process inside a combustion chamber. All the computational cases in our investigation use the Eddy-Dissipation Concept (EDC) combustion model for detailed chemistry and two equations RANS model for turbulence closure. The preliminary results of the base case without considerations on the reactions within the boundary layer show a good agreement with the experimental data in terms of pressure distribution, but the wall heat load is overestimated about 30%. Further investigation found that changing the turbulence model will alter the heat transfer characteristics significantly by delayed combustion process; changing the Prandtl number will tune the wall heat load slightly without changing the combustion field too much when using RANS model. It has been also found that the different reduced kinetic mechanism will slightly change the heat transfer characteristics through reducing the overall pressure level in the chamber by altering the thermodynamic properties of the combustion products. After carefully estimation, the boundary layer reactions for methane rocket chamber with multi-elements account for the most overestimation of the wall heat load.
  • Experimental study of dielectric barrier discharge plasma-assisted
           combustion in an aero-engine combustor
    • Abstract: Publication date: April 2020Source: Aerospace Science and Technology, Volume 99Author(s): Yi Chen, Li-Ming He, Li Fei, Jun Deng, Jian-Ping Lei, Han YuAbstractA plasma-assisted combustion test platform was established to validate the feasibility of enhancement on the performances of aero-engine combustion chamber by plasma-assisted combustion technology. The plasma-assisted combustion actuator using dielectric barrier discharge to produce plasma was designed to assemble in front of the combustor diffuser. Comparative experiments between conventional and plasma-assisted combustion were carried out to analyze the combustion enhancement effect on the combustor performances, such as average outlet temperature, combustion efficiency, pattern factor and lean blowout performance. The experimental results show that the combustion efficiency is improved by plasma, especially in fuel-rich condition (α=0.8), the increment reaches 2.31%, when α are 1 and 2, the increments are 1.21% and 0.94% respectively when voltage is 40 kV, and the increment is only 0.39% in too lean conditions (α=4). On the other hand, the higher the voltage, the more obvious increase of the combustion efficiency. The uniformity of the outlet temperature field and lean blowout performance is also improved by plasma-assisted combustion. Under the fuel-rich condition, patten factor is reduced by more than 5%. And the expansion of lean blowout limit is about 4.2%, 9.1%, 12.2% and 23.7% when the inlet velocity is range from 60 m/s to 120 m/s.
  • Multivariable finite-time composite control strategy based on immersion
           and invariance for quadrotor under mismatched disturbances
    • Abstract: Publication date: April 2020Source: Aerospace Science and Technology, Volume 99Author(s): Xinkai Li, Hongli Zhang, Wenhui Fan, Jiawei Zhao, Cong WangAbstractA novel multivariable finite-time composite integral sliding mode control algorithm with disturbance observer is investigated for a quadrotor unmanned aerial vehicle in the presence of mismatched and matched disturbances. A finite-time disturbance observer based on a multivariable supertwisting algorithm is designed as a stepping stone to estimate and cancel directly the mismatched and matched disturbances simultaneously. A composite adaptive law based on immersion and invariance theory, which introduces supervision factors and dynamic scaling factors, is proposed to compensate for the inaccuracy of modeling and the uncertainty of system parameters. Moreover, a novel integral sliding mode controller with a supertwisting term is designed by combining the composite adaptive law and the disturbance observer. The finite-time tracking of the time-varying trajectory is proven via Lyapunov-based stability analysis. Comparative simulation results illustrate the effectiveness and superiority of the proposed strategy.
  • Shear buckling analysis of functionally graded (FG) carbon nanotube
           reinforced skew plates with different boundary conditions
    • Abstract: Publication date: April 2020Source: Aerospace Science and Technology, Volume 99Author(s): O. Civalek, M.H. JalaeiAbstractIn this paper, a four-nodded straight-sided geometric element is proposed for shear buckling problem of functionally graded material (FGM) composite and carbon nanotube (CNT) reinforced composite skew plate. By using the transformation rule, quadrilateral plate field is mapped into a square domain in the computational space using discrete singular convolution (DSC) method. Related governing equations of skew plate buckling and boundary conditions of the problem are transformed from the physical domain into a square computational domain by using the geometric transformation based singular convolution. The discretization process is achieved via the DSC method together with numerical differential and two-different regularized kernel such as regularized Shannon's delta (RSD) and Lagrange delta sequence (LDS) kernels. The accuracy of the present DSC results is first verified via exiting results in literature. Then, some parametric studies have been presented to show the effects of CNT volume fraction, CNT distribution pattern, geometry of skew plate and skew angle on the shear buckling responses of FG-CNTR composite skew plates with different boundary conditions. Some new results related to critical buckling of FGM and CNT reinforced composite skew plate have been presented which can serve as benchmark solutions for future investigations.
  • Modeling and analysis of nonlinear spacecraft relative motion via harmonic
           balance and Lyapunov function
    • Abstract: Publication date: April 2020Source: Aerospace Science and Technology, Volume 99Author(s): Ayansola D. OgundeleAbstractThe harmonic balance method, an approximate method for the study of nonlinear oscillating systems described by ordinary nonlinear differential equations, is applied to the cubic polynomial approximation of the relative motion to develop approximate periodic solutions of the motion. Firstly, the cubic nonlinear terms neglected in the formulation of linear time invariant Clohessy-Wilshire (CW) equations are obtained from the original nonlinear equation of motion. Secondly, using the CW solutions as the generating solutions, harmonic balance technique is applied to the cubic nonlinear terms and correction terms are obtained. The new and improved harmonically linearized model contains CW equations and the correction terms. Using total energy as a Lyapunov function, the stability of the new model is determined. The additional terms in the new harmonically linearized spacecraft relative motion enabled us to obtain excellent approximate periodic solutions of the nonlinear equations as evidenced from the numerical simulation results.
  • Numerical research on the trailing-edge sweep of supersonic/transonic
    • Abstract: Publication date: April 2020Source: Aerospace Science and Technology, Volume 99Author(s): Xin Li, Lucheng Ji, Ling ZhouAbstractAlthough the sweep design used in fans and compressors has been the subject of considerable research and development, its application in turbines is relatively rare. This paper first studies the aerodynamic effects and mechanics of a general backward sweep in supersonic plane cascades. In response to the specific challenge of severe changes in working conditions, a new structure, combined with the characteristics of an unchanged throat and sweep around the trailing-edge, is proposed and studied. After two targeted modifications, the trailing-edge sweep is applied to a turbine stage. The results show that, similar to fans and compressors, the aerodynamic influence of the sweep in the turbine changes the working conditions and spanwise loading in spanwise tubes. Proper trailing-edge sweep design can weaken the strength of the 3D shock wave under similar flow conditions, providing a new approach for shock-weakening treatments. By applying the trailing-edge sweep, the turbine stage achieves an efficiency improvement of nearly 0.4%.
  • Analysis of aerodynamic/propulsive couplings during mode transition of
           over-under turbine-based-combined-cycle engines
    • Abstract: Publication date: April 2020Source: Aerospace Science and Technology, Volume 99Author(s): Jialin Zheng, Juntao Chang, Jicheng Ma, Daren YuAbstractMode transition control is a critical issue of a Turbine-Based-Combined Cycle (TBCC) engine when the main thrust provider changes from a gas turbine engine to a scramjet engine. Compared with a scramjet-powered integrated aircraft, a TBCC-powered aircraft involves more aerodynamic/propulsive couplings during the TBCC mode transition because of additional surfaces, such as the inlet splitter and the nozzle splitter, and the complexity of the combined engine. Those additional couplings which do not exist in a scramjet-powered vehicle would make it harder to decouple the integrated system and then to design control systems separately. This paper focuses on those additional couplings and analyzes the effects of those couplings on the control law design to realize a TBCC mode transition under a required flight condition by solving the corresponding output-tracking problem. Feedback linearization is adopted to design the control law based on the control-oriented model, which is proved to be still effective for the curve-fitted model. Simulation results demonstrate that the control strategy can meet the tracking requirements, under which the integrated aircraft is capable of maintaining a steady flight during the TBCC mode transition. The effects of the three additional couplings mentioned above on the control laws are simulated and analyzed.
  • Simulations of airborne collisions between drones and an aircraft
    • Abstract: Publication date: March 2020Source: Aerospace Science and Technology, Volume 98Author(s): Xiaohua Lu, Xinchao Liu, Yulong Li, Yingchun Zhang, Hongfu ZuoAbstractDrones are a considerable thread to the safe operation of transport aircraft with the wide application and development trend of drones. The airborne collision safety between the windshield of transport aircraft and the five light unmanned aircraft vehicles (UAVs) under various possible flight conditions is mainly studied in this paper. The collision simulation model of drone and aircraft windshield is established under PAM-CRASH software environment, which is used to simulate the dynamic response and damage of the windshield after the impact of drone and aircraft windshield under various conditions. The simulation results with that of collision test between whole-level drones and a full-scale aircraft with windshield is compared and analyzed to verify the validity of the simulation. The drone with similar bird mass is also selected to verify the appropriateness of the airworthiness rules of the windshield structure against bird impact to light UAV. The influencing parameters are important to the research results and are taken into a little consideration in the study. The research shows that the configuration, material, weight, speed and attitude of drones all have significant effects on the impact damage to the aircraft windshield, and a “parallel” attitude of a 1360 g drone which was proven to be a very severe gesture resulted in an unairworthy consequence at an impact velocity of 154.8 m/s. The windshield damage from a 1360 g drone strike is much heavier than that from a bird of a little big mass. It is expected that the high fidelity simulation model can be used to simulate all kinds of working conditions, reduce the test times, and save a lot of economic costs and provide credible technical support and basis for the airworthiness authorities to formulate precise and convincing regulatory regulations in future.
  • A Modeling Method for Aero-engine by combining Stochastic Gradient Descent
           with Support Vector Regression
    • Abstract: Publication date: Available online 13 February 2020Source: Aerospace Science and TechnologyAuthor(s): Li-Hua Ren, Zhi-Feng Ye, Yong-Ping ZhaoAbstractAero-engine aerodynamic model is widely applied to identify the aerodynamic parameters of components like compressor pressure, turbine temperature and so on. A data-driven modeling method for the aero-engine aerodynamic model by combining stochastic gradient descent with support vector regression (SGDSVR) is proposed. A novel support vector regression (SVR) training mechanism that combines batch learning with online learning is presented according to the demand and characteristic of the aero-engine aerodynamic model. In the training mechanism, batch learning is to build the initial model and online learning is to modify the online model based on the initial model. An improved sequential minimal optimization (SMO) algorithm is introduced during building the initial model phase and the SGDSVR algorithm is proposed during modifying online model phase. The simulation data of an aero-engine component-level model and the flight data of a certain aircraft are used to test the modeling method and the proposed method shows better performance compared with traditional methods.
  • Wind tunnel measurements of the surface pressure fluctuations on the new
           VEGA-C space launcher
    • Abstract: Publication date: Available online 7 February 2020Source: Aerospace Science and TechnologyAuthor(s): R. Camussi, A. Di Marco, C. Stoica, M. Bernardini, F. Stella, F. De Gregorio, F. Paglia, L. Romano, D. BarbagalloAbstractAn extensive wind tunnel test campaign devoted to the characterization of the aerodynamic and aeroacoustic behaviour of the new Space Launcher VEGA-C has been carried out in the trisonic wind tunnel available at the National Institute for Aerospace Research (INCAS) in Bucharest. The present paper summarizes the main results of the aeroacoustic investigation with a specific focus on the buffeting analysis. For this reason, the results presented herein are limited to the transonic conditions which are the most critical in terms of occurrence of flow instabilities. The scaled instrumented model of the VEGA-C launcher has been equipped with flush mounted Kulite sensors that provided the distribution of the wall pressure fluctuations. Pressure measurements at the wall of the wind tunnel have been carried out as well, with the scope of measuring reference signals to be used for cleaning the model pressure data from the back-ground noise. Numerical simulations have been also performed to facilitate the physical interpretation of the experimental outcomes by qualitatively identifying the position of the shockwaves and tracking their evolution along the launcher for increasing Mach number. Data processing provides temporal statistics of the wall pressure signals as well as spectral quantities that give clear indications about the absence of buffeting.
  • Observer-based control for spacecraft electromagnetic docking
    • Abstract: Publication date: April 2020Source: Aerospace Science and Technology, Volume 99Author(s): Keke Shi, Chuang Liu, James D. Biggs, Zhaowei Sun, Xiaokui YueAbstractElectromagnetic docking could enable autonomous spacecraft docking with no need for propellant consumption and without plume contamination. This paper addresses the robust electromagnetic docking problem for spacecraft in the presence of external disturbances, fault signals, unknown mass, elliptical eccentricity, measurement errors and input constraints. In this scenario, an intermediate observer is developed to estimate the relative motion information and the lumped disturbance resulting from these uncertainties. Based on this, an anti-disturbance controller is proposed, where the compensation of the lumped disturbance is considered. It is proved via Lyapunov analysis that the intermediate observer-based controller can achieve the objective of spacecraft electromagnetic docking with input constraints and in the presence of uncertainties. Finally, the observer-based controller is illustrated, in simulation, to demonstrate the effectiveness and improved performance compared with a disturbance observer-based controller.
  • Composite wing box deformed-shape reconstruction based on measured
           strains: Optimization and comparison of existing approaches
    • Abstract: Publication date: April 2020Source: Aerospace Science and Technology, Volume 99Author(s): Marco Esposito, Marco GherloneAbstractThe reconstruction of the displacement field of a structure (shape sensing) has become crucial for the Structural Health Monitoring of aerospace structures and for the progress of the recently developing morphing structures. As a consequence, shape sensing techniques based on discrete surface strains measurements have seen a consistent expansion in the last few years. In this paper, the three main shape sensing methods, the Modal Method, the Ko's displacements theory and the inverse Finite Element Method, are presented. The most recent and also novel improvements are discussed and added to the methods' formulations. Then, the three methods are numerically applied to a complex aerospace structure such as that of a composite wing box experiencing bending and twisting deformations. For the first time, a detailed investigation on the optimal strain sensors configuration is performed for all the three techniques simultaneously. Finally, the methods' performances, in terms of accuracy of the reconstruction and of number of required sensors, are compared. The three methods show different characteristics that make them suitable for different applications, depending on the level of accuracy and the number of strain information required. The iFEM is proven to be the more accurate but the more demanding in terms of required sensors; the Ko's displacement theory is capable of giving a rough estimation of the displacement field, but requires a small amount of sensors; the Modal Method represents a trade-off between the other two in terms of accuracy and number of sensors required.
  • Highly maneuvering target interception via robust generalized dynamic
           inversion homing guidance and control
    • Abstract: Publication date: April 2020Source: Aerospace Science and Technology, Volume 99Author(s): Belkacem Kada, Uzair Ansari, Abdulrahman H. BajodahAbstractThe Robust Generalized Dynamic Inversion (RGDI) homing guidance (HG) and control design methodology is proposed for intercepting highly maneuvering targets. The two-loops RGDI based system is designed by prescribing dynamic constraints that encapsulate the interception kinematic and dynamic objectives. The constraints are generalized-inverted using the dynamically scaled Moore-Penrose Generalized Inverse to solve for the HG and control variables. Sliding mode control (SMC) elements are augmented in the two loops, and are aimed to equip the closed loop RGDI system with a guaranteed semi-global practically stable tracking performance and robustness against tracking performance degradation due to modeling and parametric uncertainties, exogenous disturbances, and generalized inversion scaling. Numerical simulations of the RGDI system are conducted using nonlinear planar engagement kinematics and interceptor dynamics. The proposed system's performance is compared with a classical SMC based HG & Control system and with a conventional Augmented Proportional Navigation HG system that is made by replacing the outer guidance loop in the RGDI system for different interception scenarios. The simulations results reveal that the RGDI system supersedes its counterparts in accomplishing direct collision and hit-to-kill HG interception requirements, demonstrating its high-level performance abilities and robustness attributes against agile target maneuvers in the presence of uncertainties, disturbances, and measurement noises.
  • Experimental study of the critical incidence phenomena in low speed
           compressor stators with both conventional and 3D blading designs
    • Abstract: Publication date: Available online 7 February 2020Source: Aerospace Science and TechnologyAuthor(s): Guangfeng An, Shuai Zhang, Xianjun Yu, Baojie Liu, Guoqiang YiAbstractThree-dimensional (3D) corner separation, which can arise significant flow blockage and loss production, is an inherent and detrimental flow feature inside the stator of axial compressors, hence, numerous active and passive methods have been investigated to prevent the corner separation turning into corner stall in recent years. This paper investigates experimentally the 3D separating flows and the Critical Incidence Phenomenon inside the stator passage of a serious of low-speed large-scale axial compressors with or without 3D blading designs, and it is expected that some rules for design of highly loaded stator could be found. Firstly, the experimentally observed Critical Incidence Phenomenon in the previous study was reviewed, and this phenomenon was validated in the other compressors with different design parameters. Then, the effects of blade loading, blade airfoils, and 3D blading designs on the critical incidence were discussed. And the results revealed that blade loading and 3D blading designs seem to have no impact on the critical incidence, while blade airfoils should be considered as an impact factor on it. Besides, though 3D blading designs could not increase the critical incidence, it can reduce the actual inlet incidence instead, hence, after the stator is redesigned using the 3D stacking, the corner stall is delayed. Finally, based on the above-mentioned analysis, a critical-incidence-based metric named Incidence Reserve was proposed to guide the design of the highly loaded stator. It is expected that at the initial stage of a stator the distribution of the Incidence Reserve along the span is as consistent as possible, so that the detrimental hub corner stall will be eliminated or deferred.
  • Reinforcement learning for angle-only intercept guidance of maneuvering
    • Abstract: Publication date: April 2020Source: Aerospace Science and Technology, Volume 99Author(s): Brian Gaudet, Roberto Furfaro, Richard LinaresAbstractWe present a novel guidance law that uses observations consisting solely of seeker line-of-sight angle measurements and their rate of change. The policy is optimized using reinforcement meta-learning and demonstrated in a simulated terminal phase of a mid-course exo-atmospheric interception. Importantly, the guidance law does not require range estimation, making it particularly suitable for passive seekers. The optimized policy maps stabilized seeker line-of-sight angles and their rate of change directly to commanded thrust for the missile's divert thrusters. Optimization with reinforcement meta-learning allows the optimized policy to adapt to target acceleration, and we demonstrate that the policy performs better than augmented zero-effort miss guidance with perfect target acceleration knowledge. The optimized policy is computationally efficient and requires minimal memory, and should be compatible with today's flight processors.
  • Convex relaxation for optimal rendezvous of unmanned aerial and ground
    • Abstract: Publication date: Available online 31 January 2020Source: Aerospace Science and TechnologyAuthor(s): Zhenbo Wang, Spencer T. McDonaldAbstractIn this paper, we present an optimal rendezvous approach that facilitates the coordination of unmanned aerial vehicles (UAVs) and unmanned ground vehicles (UGVs). Of particular interest is the problem of UAV-UGV rendezvous for autonomous aerial refueling, which is an important capability for the future use of UAVs. The main contribution of this work is the development of a promising method for the generation of optimal rendezvous trajectories in real time using convex optimization. First, the UAV-UGV rendezvous problem is formulated as a nonconvex optimal control problem using an error dynamic model, where the state and control variables are highly coupled. Then, great effort is devoted to reducing the nonconvexity and coupling of the dynamics through change of variables. Based on a lossless convexification technique, a sequential convex programming algorithm is designed and the solution is obtained by solving a sequence of convex optimization problems. Theoretical proof is provided to demonstrate the exactness of convexification. Furthermore, a simple line search technique is introduced to handle the model error resulting from the linear approximations and to improve the convergence of the designed successive process. Simulation results of two refueling rendezvous situations validate that the proposed method is capable of converging to the solutions within a second, which shows great potential for real-time applications.
  • Active fault-tolerant control for a quadrotor helicopter against actuator
           faults and model uncertainties
    • Abstract: Publication date: Available online 31 January 2020Source: Aerospace Science and TechnologyAuthor(s): Ban Wang, Yanyan Shen, Youmin ZhangAbstractThis paper proposes an active fault-tolerant control strategy for a quadrotor helicopter against actuator faults and model uncertainties while explicitly considering fault estimation errors based on adaptive sliding mode control and recurrent neural networks. Firstly, a novel adaptive sliding mode control is proposed. In virtue of the proposed adaptive schemes, the system tracking performance can be guaranteed in the presence of model uncertainties without stimulating control chattering. Then, due to the fact that model-based fault estimation schemes may fail to correctly estimate fault magnitudes in the presence of model uncertainties, a fault estimation scheme is proposed by designing a parallel bank of recurrent neural networks. With the trained networks, the severity of actuator faults can be precisely estimated. Finally, by synthesizing the proposed fault estimation scheme with the developed adaptive sliding mode control, an active fault-tolerant control mechanism is established. Moreover, the issue of actuator fault estimation error is explicitly considered and compensated by the proposed adaptive sliding mode control. The effectiveness of the proposed active fault-tolerant control strategy is validated through real experiments based on a quadrotor helicopter subject to actuator faults and model uncertainties. Its advantages are demonstrated in comparison with a model-based fault estimator and a conventional adaptive sliding mode control.
  • The pressure wave induced by an asymmetrical Dielectric Barrier Discharge
           plasma actuator under the influence of residual charge
    • Abstract: Publication date: Available online 30 January 2020Source: Aerospace Science and TechnologyAuthor(s): Xin Zhang, Y.D. Cui, Chien-Ming Jonathan Tay, B.C. KhooAbstractIn order to improve the control effect of the plasma actuator at high Mach and Reynolds number, the pressure wave induced by an asymmetrical Alternating Current (AC) Dielectric Barrier Discharge (DBD) plasma actuator which is under the influence of residual charge on the dielectric surface is investigated by using a high accuracy phase-lock image Schlieren technique and a pressure-field microphone in quiescent air. It is of great significance that a single pressure wave created by the AC-DBD plasma actuator with the residual charge which co-exists with the starting vortex in the induced flow field is first observed. Based on the experimental results, a hypothesis is proposed on the formation mechanism of the induced pressure wave. This finding might inspire a new discussion on the control of the temporal and spatial distributions of the surface charge which might be helpful to flow separation at high Reynolds and Mach numbers.
  • A dual adaptive fault-tolerant control for a quadrotor helicopter against
           actuator faults and model uncertainties without overestimation
    • Abstract: Publication date: Available online 29 January 2020Source: Aerospace Science and TechnologyAuthor(s): Ban Wang, Xiang Yu, Lingxia Mu, Youmin ZhangAbstractThis paper presents a dual adaptive fault-tolerant control strategy for a quadrotor helicopter based on adaptive sliding mode control and adaptive boundary layer. Within the proposed adaptive control strategy, both model uncertainties and actuator faults can be compensated without the knowledge of the uncertainty bounds and fault information. By virtue of the proposed adaptive control scheme, the minimum discontinuous control gain is adopted, which significantly reduces the control chattering effect. As compared to the existing adaptive sliding mode control schemes in the literature, larger actuator faults can be tolerated by employing the proposed control scheme while suppressing control chattering. Moreover, boundary layer is used to smoothen control discontinuity and further eliminate control chattering. Nevertheless, the choice of boundary layer thickness is a trade-off between system stability and tracking accuracy. By explicitly considering this fact, an adaptive boundary layer is developed and synthesized with the proposed adaptive control framework to ensure stability and tracking accuracy of the considered system. When the control parameter tends to be overestimated, the thickness of boundary layer can be appropriately adjusted to avoid control parameter overestimation. Simulation and experimental tests of a quadrotor helicopter are both conducted to validate the effectiveness of the proposed control scheme. Its advantages are demonstrated in comparison with a conventional adaptive sliding mode control scheme.
  • Analysis of near stall condition of high bypass fan rotor based on
           airworthiness certification
    • Abstract: Publication date: April 2020Source: Aerospace Science and Technology, Volume 99Author(s): Kai Zhang, A.J. WangAbstractIn order to ensure flight safety, the near stall condition test is one of the most important steps in the airworthiness certification phase of civil aircraft. The twisted and swept fan is one of the most important components of the high bypass ratio engine. The unsteady flow field of the fan rotor was calculated by CFD simulation. The unsteady flow at the tip of blades is an important cause of the development of the fan rotor from near stall condition to stall condition. Dynamic mode decomposition method (DMD) was applied to analysis unsteady flow field of the blade tip area. And compressed sensing dynamic mode decomposition (CPDMD) method was tried to deal with massive grid based on dynamic mode decomposition method. The dynamic mode decomposition method successfully separates the flow structure of the unsteady flow field. The compressed sensing dynamic mode decomposition method compresses the initial flow field snapshot to obtain low-dimensional flow field data. Then the DMD modality is got by analyzing the low-dimensional flow field. Finally, the DMD mode with the initial flow field snapshot dimension is obtained. So the CPDMD method effectively reduces algorithm runtime and reduces the computational resource requirements. This provides the possibility to process a large amount of flow field data that would otherwise be impossible to process.
  • Deformation behaviors and energy absorption of auxetic lattice cylindrical
           structures under axial crushing load
    • Abstract: Publication date: March 2020Source: Aerospace Science and Technology, Volume 98Author(s): Yongguang Guo, Jian Zhang, Liming Chen, Bing Du, Houchang Liu, Liliang Chen, Weiguo Li, Yizhi LiuAbstractAuxetic lattice cylindrical shell consisting of re-entrant lattice cell has superior mechanics performance due to negative Poisson's ratio (NPR). In this work, finite element models were established and finite element method (FEM) was adopted to explore the deformation behaviors and energy absorption of re-entrant auxetic lattice cylindrical shells and the SILICOMB lattice cylindrical shell in different crushing velocity. Result indicated that the deformation behaviors of auxetic lattice cylindrical shells are expressed as NPR mode, and different deformation behaviors (‘Z’ mode, ring mode, diamond mode and mixed mode) of lattice sandwich cylindrical shells are mainly determined by the ratio of the core wall thickness to the skin. The SILICOMB lattice cylindrical shell has better performance than the conventional hexagon honeycomb lattice cylindrical shell when crushing under high velocity. Meanwhile, the SILICOMB sandwich cylindrical shell has the best performance on the specific energy absorption among these four configurations and the maximum increment is 20.77%. The design of the auxetic and the SILICOMB cylindrical shells may provide a new idea which can be applied in practical application as impact resistance structures on aerospace.
  • Design and aerodynamic performance analysis of a variable-sweep-wing
           morphing waverider
    • Abstract: Publication date: March 2020Source: Aerospace Science and Technology, Volume 98Author(s): Pei Dai, Binbin Yan, Wei Huang, Yifei Zhen, Mingang Wang, Shuangxi LiuAbstractThe wide speed range and large flight envelope of the hypersonic vehicle require that its aerodynamic configuration still has good aerodynamic performance at low Mach number. Therefore, the variable Mach number waverider is proposed to achieve good flight performance in a wide speed range. In this paper, based on the delta-winged variable Mach number waverider, a variable-sweep-wing morphing waverider is proposed and studied, including four specific sweep-wing configurations, namely loiter, standard, dash and wing-retracted configurations. In the current study, the aerodynamic performances of this variable-sweep-wing morphing waverider with four configurations are investigated under subsonic/supersonic/hypersonic flight conditions. At the same time, the numerical approaches employed are validated against the available experimental data in the open literature. The obtained results show that compared with the wing-retracted configuration, the best flight performance of this variable-sweep-wing morphing waverider can be achieved using different configuration for different flight condition. However, within the hypersonic speed range, the aerodynamic performance is improved through morphing but its advantages are not as large as that in the subsonic speed. Besides, effects of wing downwash and shockwave are also analyzed. In conclusion, the variable-sweep-wing morphing waverider improves both low-speed and high-speed aerodynamic performances, and it expands the flight speed range.
  • Drag bookkeeping on an aircraft with riblets and NLF control
    • Abstract: Publication date: March 2020Source: Aerospace Science and Technology, Volume 98Author(s): Benedetto Mele, Lorenzo Russo, Renato TognacciniAbstractIn the frame of aerodynamic drag reduction for low-emission target, Natural Laminar Flow (NLF) control and riblets are the most interesting passive techniques. The drag breakdown in the case of riblets installed is a topical matter for understanding on which form of drag riblets act. Indeed, linear theories and flat-plate experiments show that riblets act on friction drag whereas other experiments in pressure-gradient flow revealed an increased performance of riblets that was not interpreted. A contribution to clarify this effect has been recently provided analyzing the effect of riblets on form drag in two-dimensional pressure-gradient flows. In the present paper the effect of riblets on pressure drag is discussed also in three-dimensional flows analyzing CFD solutions of the flow around an innovative regional turbo-prop aircraft (wing-body configuration) with NLF and riblets installed. For the first time, the contribution of each of two drag reduction systems is identified and a deep analysis on which form of drag riblets act in three-dimensional flow is proposed thanks to a far-field aerodynamic drag breakdown.
  • Automatic stability control using tip air injection in a multi-stage axial
           flow compressor
    • Abstract: Publication date: March 2020Source: Aerospace Science and Technology, Volume 98Author(s): Jichao Li, Yang Liu, Juan Du, Hongwu ZhangAbstractAutomatic stability control is experimentally investigated using tip air injection in a multi-stage axial flow compressor. The warning signal is obtained according to the auto- and cross-correlation algorithm integrated in the digital signal processing controller for online monitoring wall pressure signal at different stages. It is used as a feedback signal to drive the discrete proportional jet valves. The control law depends on the stall margin improvement (SMI) versus injected momentum ratio relationship. The real-time detection algorithms based on the correlation theory are experimentally studied using a collection of pressure transducers installed symmetrically on the casing at different stages. The results reveal that because of the initial appearance of stall inception at the first stage, which induced a stall in all stages of the compressor, only the correlation coefficient detected at the first stage exhibits a decreasing trend in the throttling process. Moreover, only the tip air injection at the first stage can obtain a large SMI. Therefore, the correlation coefficient detected at the first stage is used as the feedback signal in the controller after verifying the influence of injected air on the correlation coefficient. Compared with steady injection, the proposed automatic stability control can save energy when the compressor is stable and provide protection to extend the stall margin when close to the stall point. With a nearly equal SMI as the steady injection (the maximum SMI is 20%), the energy of the active injected air is approximately one-fifth of the steady injection. This active control can fully consider the failure control such that when the control system is inoperative, the aero-engine continues safe operate under normal circumstances.
  • Nonsingular terminal sliding mode control for a quadrotor UAV with a total
           rotor failure
    • Abstract: Publication date: March 2020Source: Aerospace Science and Technology, Volume 98Author(s): Zhiwei Hou, Peng Lu, Zhangjie TuAbstractBased on nonsingular terminal sliding mode control (NTSMC), a flight controller is proposed in this paper for a quadrotor with a total rotor failure. The proposed method is a finite-time position and attitude tracking approach with strong robustness. At first, the fault-tolerant controller for the quadrotor with a total rotor failure is derived, and the model uncertainties and wind disturbances are considered. The dynamic model of the quadrotor is introduced and divided into two control loops: the inner control loop and the outer control loop. Based on the division of the control system, the NTSMC based inner controller is designed which makes the attitude dynamics converge to the desired attitude in finite-time. And the NTSMC based outer controller is derived which generates the desired attitude for the inner controller and makes the dynamics converge to the desired position in finite-time. The stability of the closed-loop system is analyzed by Lyapunov theory and the stability conditions are obtained. Then, in order to improve the practicability of the control algorithm, a flight controller for a fault-free quadrotor is proposed which has a similar structure compared with the fault-tolerant one. A fault detection and isolation method is applied to detect the fault and reconfigure the flight controllers. Moreover, two estimation methods for external disturbance and model uncertainties are applied to enhance the robustness of the proposed flight controller. The estimated wind disturbances results are introduced into the outer controller to compensate for the effect of disturbance while the model uncertainties estimator is applied in the inner control loop. Finally, numerical simulation results show the great performance of the proposed flight control method.
  • Effect of blended blade tip and winglet on aerodynamic and aeroacoustic
           performances of a diagonal fan
    • Abstract: Publication date: March 2020Source: Aerospace Science and Technology, Volume 98Author(s): Chun Zhang, Lucheng Ji, Ling Zhou, Shijun SunAbstractTip clearance flow significantly influences the aerodynamic and aeroacoustic performances of a fan. Its mechanism and control have remained a hotspot in the field of turbomachinery. In this paper, blended blade tip and winglet (BBTW) technology is proposed based on the combination of blended blade and end wall and winglet technologies, while the flow field and sound field of an original blade and the BBTW-blade of a diagonal fan are numerically simulated. The calculation results show that the BBTW design reduces the pressure difference between the suction surface and the pressure surface in the blade tip area and enhances the necking effect of tip clearance, thus inhibiting tip clearance leakage flow and improving the aerodynamic performance of the diagonal fan. In addition, as BBTW weakens the tip leakage vortex intensity, thereby reducing the sound pressure level of the near-field and the far-field noise.
  • Low speed lateral-directional aerodynamic and static stability analysis of
           a hypersonic waverider
    • Abstract: Publication date: March 2020Source: Aerospace Science and Technology, Volume 98Author(s): Tamas Bykerk, Dries Verstraete, Johan SteelantAbstractHypersonic waveriders have the potential to significantly reduce travel times on long haul civilian transport routes. The design of hypersonic aircraft is heavily influenced by the aerodynamic efficiency at the cruise Mach number, resulting in less than ideal geometries for subsonic flight. Waverider aerodynamics and stability in the low speed regime is rarely investigated and not well understood, but is crucial for horizontal take-offs and landings. This paper presents a combination of numerical simulation results and experimental data for the low speed propelled variant of the Mach 8 HEXAFLY-INT waverider. Aerodynamic, control and stability testing for lateral-directional cases was conducted in the University of Sydney 4 foot by 3 foot low speed facility. Computational fluid dynamics simulations are compared with wind tunnel tests for angles of attack between -5 and 15 degrees and angles of sideslip between -8 and 8 degrees. Throughout these ranges, aileron and rudder deflections up to 10 degrees are investigated. Results show that the vehicle aerodynamics are dominated by asymmetric wing and fin vortices, resulting in non-linear aerodynamic forces. At a centre of gravity location of 44.4% of the vehicle length the aircraft is stable directionally, but has lateral instability at angles of attack below -2 degrees. This is attributed to the low mounted wings with anhedral. The instability is minor and is not expected to result in an uncontrollable condition. Lowering the centre of gravity by approximately 2 centimetres, or 17% of the local fuselage height, can correct the instability. Lateral-directional dynamic stability was predicted using static derivatives and was found to be stable through the entire AoA range tested, with no dependence on mass moments of inertia. Both aileron and rudder controls are found to provide sufficient control authority, but the aircraft may benefit from increased rudder size. The results from this study, along with previous work on longitudinal stability and performance show the feasibility of moving to the flight test program, but aircraft dynamic stability must first be investigated.
  • Practical finite time adaptive robust flight control system for
           quad-copter UAVs
    • Abstract: Publication date: March 2020Source: Aerospace Science and Technology, Volume 98Author(s): Karam Eliker, Said Grouni, Mohamed Tadjine, Weidong ZhangAbstractThis paper investigates the practical finite time control problem with uncertain parameters, external disturbances, and input saturation for quad-copter unmanned aerial vehicle systems. Firstly, an adaptive robust controller based on backstepping with fast terminal sliding mode control is designed for the major control loops. Secondly, only four adaptation laws are used to estimate the quad-copter uncertain parameters while a projector algorithm is used to guarantee the estimation within a prescribed range. Thirdly, an adaptive switching gain is developed to compensate the lumped disturbances. Finally, a compensator term is introduced in control design to reduce the adverse effect caused by the input saturation. The proposed control scheme can attenuate chattering phenomenon and guarantee that all states of the closed-loop system are practical finite time stable. The validity of the proposed flight control system is confirmed by using several flight scenarios under various conditions and a comparison study with other works is made showing the effectiveness, robustness, adaptiveness, and energy efficiency of the proposed approach.
  • Information-reusing alignment technology for rotating inertial navigation
    • Abstract: Publication date: Available online 28 January 2020Source: Aerospace Science and TechnologyAuthor(s): Qiangwen Fu, Sihai Li, Yang Liu, Feng WuAbstractThis paper proposes a real-time information-reusing alignment strategy for rotating inertial navigation system (INS), which is devoted to improving alignment performance in a limited time by making full use of all sensor data. In the forward coarse alignment stage, all sensor data was recorded for subsequent fine alignment while the rough attitude was determined, and the well-designed data extraction strategy significantly reduced the memory space and the computation load without losing solution accuracy. In the fine alignment stage, the stored data was used for backtracking navigation update and Kalman filtering in the local level geographic frame rather than the inertial frame, simplifying the fine alignment algorithm. The lever arm between the center of the inertial measurement unit (IMU) and the rotating center of the turntable was automatically estimated by the fine alignment Kalman filter, which helped to suppress motion interference and improve alignment performance. Finally, the simulation tests and the practical experiments justified the proposed information-reusing alignment method.
  • Aerodynamic shape optimization by continually moving ROM
    • Abstract: Publication date: Available online 27 January 2020Source: Aerospace Science and TechnologyAuthor(s): Li Zhou Li, Jin Jiu Li, Jun Zhang, Kuan Lu, Mei Ni YuanAbstractThis study presents a multi-round shape optimization method based on the aerodynamic reduced order model. In this method, each optimization round employs the first-order Taylor reduced order model to provide aerodynamic characteristics in the local design space around the baseline point. At the end of each optimization round, the baseline point moves toward the optima and a new reduced order model is built at the new baseline point for the next optimization round. In this way, the provided method explores the whole design space. The NACA0012 airfoils in subsonic and transonic flows are used to verify the method with the lift-drag ratio as optimization objective. The design parameters are the shape control parameters of the airfoil. The optimization algorithm is a genetic algorithm. After a total of 20 simulations by two rounds of optimizations, the lift-drag ratios increase from 0 to 25 and from 0 to 8 in the subsonic case and transonic case, respectively. An additional trick was suggested for addressing shockwave effect in the transonic case, in which the influence of the airfoil shape on the aerodynamic forces is divided into the following two parts: the smoothed part to address the influence of the shape on the smooth aerodynamic distribution, and the shockwave part to address the influence of the shape on shockwave strength, shockwave position, and their effect on the aerodynamic distribution. The optimizations of the airfoils in subsonic and transonic flows yield different trends. The results of the subsonic case tend to increase the airfoil curvature to increase the lift-drag ratio. The results of the transonic case tend to move the upper surface shockwave to the trailing edge, while moving the lower surface shockwave to the leading edge. Independent of the case, the provided optimization method can properly address the resulting trends and increase the efficiency of the optimization.
  • Corner separation dynamics in a high-speed compressor cascade based on
           detached-eddy simulation
    • Abstract: Publication date: Available online 27 January 2020Source: Aerospace Science and TechnologyAuthor(s): Ruiyu Li, Limin Gao, Chi Ma, Shiyan Lin, Lei ZhaoAbstractCorner separation is a vital unsteady flow phenomenon in a compressor and plays an essential role in flow field instability. Moreover, corner separation under high subsonic Mach number conditions is a scientific problem with practical engineering significance. The prime motivation of this work is using the most cost-effective method to more accurately resolve the corner separation phenomenon in a high-speed compressor cascade and to determine its inherent unsteady behavior under high-incidence conditions. Two configurations of the cascade are investigated, an incidence of 5.0° (stable) and 7.5° (unstable, near-stall), using a detached eddy simulation (DES) set to Re=7.5×105 and Ma=0.6 at the inlet boundary. The flow structures are analyzed after verifying the accuracy of the simulation. The dynamic mode decomposition (DMD) method is applied to analyze the main sources of instability in the flow field. By comparing the stable and unstable conditions, the suction surface separation vortex (SSV) found in the recirculation region is the key factor for flow instability. The mixture of the SSV and corner vortex (CV) is the main reason for the flow field becoming unstable and eventually causing a high-incidence stall. The formation and unsteady characteristics of the SSV at the near-stall condition are discussed. Suggestions for flow control and stall warning are also given. Thus, the results can provide a theoretical basis for the flow control of a high-speed compressor blade.
  • Effect of Reynolds number and slot guidance on passive infrared
           suppression device
    • Abstract: Publication date: Available online 27 January 2020Source: Aerospace Science and TechnologyAuthor(s): Lakhvinder Singh, S.N. Singh, S.S. SinhaAbstractInfrared suppression system (IRSS) is a critical component for the survivability of a combat vehicle. An ejector diffuser acts as a passive IRSS device with the primary role in suppressing heat signatures. The current work on passive IRSS device deals with the design modifications in the diffuser part of the ejector diffuser. The modifications are in terms of the slots provided inline with the diffuser wall and a guidance at the slot which has not been explored earlier. We performed numerical simulations to examine the effect of Reynolds number and slot guidance shape on the performance of guided-slot ejector-diffuser. Further, experimental results have been used to validate our numerical approach. We find that the performance of the curved-guided-slot ejector diffuser is independent of the Reynolds number in the range of 4.3×104≤Renz≤4.2×105. Further, three types of guidance at the slots for an ejector diffuser have been investigated along with a baseline case of no-slot-guidance. Two guidance have curved shape while the third is straight plate guidance. The difference in the performance of the two shapes of the curved guidance at the slot is marginal whereas a noticeable difference is observed for the straight plate guidance. Curved guidance in comparison to straight plate guidance has higher mass entrainment (≈6%), marginally better wall temperature distribution and significantly higher static pressure recovery (Cp=0.41 in comparison to Cp=0.15). Compared to the no guidance case, the curved plate guidance has better performance in terms of cumulative mass entrainment (≈8%) and lower wall temperatures.
  • Investigation of flow characteristics inside a dual bell nozzle with and
           without film cooling
    • Abstract: Publication date: Available online 27 January 2020Source: Aerospace Science and TechnologyAuthor(s): Mayank Verma, Nitish Arya, Ashoke DeAbstractIn this study, we perform a two-dimensional axisymmetric simulation to assess the flow characteristics and understand the film cooling process in a dual bell nozzle. The secondary stream with low temperature is injected at three different axial locations on the nozzle wall, and the simulations are carried out to emphasize the impact of injection location (secondary flow) on film cooling of the dual bell nozzle. The cooling effect is demonstrated through the temperature and pressure distributions on the nozzle wall or, in-turn, the separation point movement. Downstream of the injection point, the Mach number and temperature profiles document the mixing of the main flow and secondary flow. The inflection region is observed to be the most promising location for the injection of the secondary flow. We have further investigated the effect of Mach number of the secondary stream. The current study demonstrates that one can control the separation point in a dual bell nozzle with the help of secondary injection (Mach number) so that an optimum amount of thrust can be achieved.
  • Effects of installation location of fluidic oscillators on aerodynamic
           performance of an airfoil
    • Abstract: Publication date: Available online 27 January 2020Source: Aerospace Science and TechnologyAuthor(s): Sang-Hyuk Kim, Kwang-Yong KimAbstractThis study numerically investigates the effects of the installation location of fluidic oscillators on the aerodynamic performance of a NACA 0015 airfoil. The oscillators are embedded in the suction surface of the airfoil. An unsteady flow analysis was performed using three-dimensional Reynolds-averaged Navier-Stokes equations with a shear stress transport turbulence model. The angle of attack was varied from 0 to 24 degrees, and four installation locations between the leading and trailing edges of the airfoil were used to analyze the lift and drag performance. The installation location with the highest lift coefficient depended on the angle of attack (and thus the location of flow separation), but greater drag reduction was obtained when the fluidic oscillators were installed closer to the leading edge regardless of the angle of attack. This location of the oscillators also produced the highest lift-to-drag ratio.
  • The three-dimensional acoustic field in cylindrical and annular ducts with
           an axially varying mean temperature
    • Abstract: Publication date: Available online 27 January 2020Source: Aerospace Science and TechnologyAuthor(s): Jingxuan Li, Aimee S. Morgans, Lijun YangAbstractThis paper presents analytical solutions for the three dimensional acoustic field in cylindrical and annular ducts with dependence of mean temperature on axial position. A wave equation for the pressure perturbation is constructed in cylindrical coordinates, applying a zero mean flow condition. Separation of variables is used to express the pressure perturbation as a product of functions which vary only axially, radially and circumferentially. The axial dependence of the mean temperature means that a general analytical solution for the axial second order ordinary differential equation (ODE) cannot be obtained. Variable transformation is applied, yielding a standard second order ODE with known solutions for linear and quadratic axial mean temperature dependence. The acoustic field and resonant frequencies for an annular duct with linear/quadratic axial mean temperature variation predicted using these solutions match perfectly with those calculated using the linearised Euler equations. The analytical solution for the linear mean temperature profile is applied to more complicated profiles in a piecewise linear manner, axially segmenting the temperature profile into regions that can be approximated as linear. The acoustic field and resonant frequency are predicted very accurately even when very few axial segments are used.
  • Satellite proximate pursuit-evasion game with different thrust
    • Abstract: Publication date: Available online 23 January 2020Source: Aerospace Science and TechnologyAuthor(s): Dong Ye, Mingming Shi, Zhaowei SunAbstractThis paper investigates the proximity satellite pursuit-evasion game where the pursuer carries three orthogonal thrusters, each of which can exert a magnitude-limited force along one axis, while the evader has a single thruster which can produce a bounded acceleration in any direction. The pursuer or the evader tries their best to capture or escape by performing orbital maneuvers. By establishing a local moving coordinate frame on the originally revolving orbit of the evader, we reduce the dynamics of each player to the linear Clohessy-Wiltshire equations. We then transform the problem of finding the optimal controls associated with the saddle point solution of the game into the two-point boundary value problem, which is solved by combining the heuristic searching and Newton method. At last, by numerical simulations, we discuss the effectiveness of the proposed algorithm in finding the open-loop solution to the game. We show that, in contrast with pursuit-evasion games where each player has one single thruster, the problem considered in this paper may not be solved efficiently by the indirect method, since there exist some initial states of the players such that the proposed algorithm fails to solve the open-loop control.
  • Free vibrational characteristics of grid-stiffened truncated composite
           conical shells
    • Abstract: Publication date: Available online 22 January 2020Source: Aerospace Science and TechnologyAuthor(s): M. Zarei, G.H. Rahimi, M. HemmatnezhadAbstractIn this paper, the free vibrational behavior of composite conical shells stiffened by bevel stiffeners is investigated using experimental, analytical and numerical techniques. The smeared method is employed to superimpose the stiffness contribution of the stiffeners with those of shell in order to obtain the equivalent stiffness parameters of the whole structure. Due to the specific geometry of the conical shell, the whole structure is converted to a conical shell with variable stiffness and thickness. The stiffeners are considered to be of beam-type which support shear load and bending moments in addition to the axial loads. The geodesic path is applied to the stiffeners. The governing equations have been derived based on the first-order shear deformation theory and using the Ritz method. In order to validate the analytical achievements, the experimental modal test is conducted on a stiffened cone. The specimen has been fabricated by a specially-designed filament winding setup. A 3-D finite element model was also built using ABAQUS software to further validate the analytical results and help with parametric study. Comparison of the results obtained from the three approaches revealed good agreements. The effects of the shell geometrical parameters and variations in the cross stiffeners angle on the natural frequencies have been discussed and investigated. The present achievements are novel and can be used as a benchmark for further studies.
  • Effect of turbine rotor disc vibration on hot gas ingestion and
           rotor-stator cavity flow
    • Abstract: Publication date: Available online 22 January 2020Source: Aerospace Science and TechnologyAuthor(s): Xingyun Jia, Lidong He, Hai ZhangAbstractGas ingestion numerical model and rotor-stator (RS) cavity model with rotor vibration are proposed to investigate the effect of rotor disc vibration on the flow characteristics in the RS cavity and gas ingestion. Results show that the sealing effectiveness at the rim seal clearance is increased due to the rotor disc vibration, and the amplitudes of the pressure fluctuation and the radial velocity of the gas ingestion flow at rim seal clearance are reduced. The speed of the Kelvin-Helmholtz vortex is approximately 82.28% of that without rotor vibration. Numerical results reveal that the effect of the rotor vibration on the RS cavity flow covers the whole cavity, the unsteady flow at the low radius region in RS cavity is caused by the vibration rotor disc. The frequency of the pressure oscillation at the low radius region is approximately equal to the vibration frequency of the rotor disc as the periodic variation of the tip clearance shape in the horizontal intake section of sealing air and the induced periodic mass flow rate of inlet sealing air. The flow characteristics in RS cavity are together determined by the rotor vibration parameters and the Kelvin-Helmholtz vortices in rim seal region.
  • A new approach of endwall recirculation in axial compressors
    • Abstract: Publication date: Available online 22 January 2020Source: Aerospace Science and TechnologyAuthor(s): Hossein KhaleghiAbstractThis study reports on a novel design of endwall recirculation in a transonic axial flow compressor rotor. The bled air from downstream of the rotor is injected over the tip clearance of the blades in the circumferential direction opposite to the blade rotation. Results reveal that the treated casing effectively increases the operating range of the rotor at the expense of some efficiency loss. It is shown that the endwall recirculation applied in this work does not reduce the incidence angle upstream of the blade but reduces the pressure difference between the pressure and suction surface and pushes the leakage vortex and the passage shock downstream. The endwall flow was found to be responsible for stall initiation in the treated casing (similar to the smooth casing).
  • Flowing residence characteristics in a dual-mode scramjet combustor
           equipped with strut flame holder
    • Abstract: Publication date: Available online 21 January 2020Source: Aerospace Science and TechnologyAuthor(s): Hongchao Qiu, Junlong Zhang, Xiaoxue Sun, Juntao Chang, Wen Bao, Silong ZhangAbstractIn this paper, the numerical and experimental investigations have been conducted to study the flowing residence characteristics in the supersonic combustor. A thin strut equipped in the center of the combustor was adopted as the flame holder. The airflow residence time was calculated by a new method, and results proved that the airflow residence time was longest in the center of the recirculation zone. Then, the interaction between the strut geometrical parameters and the airflow residence time was investigated. The results revealed that the thickness of the strut played a decisive role on the residence time and with the strut thickness increasing, the residence time became longer. The sweep angle and the length of the strut had few influences on the residence time. A series of experiments have been conducted at the flight condition of Ma=6.5. Experimental results indicated that, with the help of the long residence time in the recirculation zone, an initial flame was formed in the recirculation zone at the tailing of the strut, and grew to form a steady flame in the center of the main flow. These results are valuable for the future optimization of the combustion stabilization in the supersonic combustor using strut as the flame holder.
  • Numerical analysis on the disintegration of gas-liquid interface in
           two-phase shear-layer flows
    • Abstract: Publication date: Available online 21 January 2020Source: Aerospace Science and TechnologyAuthor(s): Gaoming Xiang, Zhaoxin Ren, Seungcheol Kim, Bing WangAbstractThe gas-liquid two-phase interfacial flow widely occurs in atomization or liquid film cooling in the field of aerospace engineering. In this paper, the gas-liquid two-phase shear layer is simulated using volume of fraction method by numerical solver Gerris. Both the two-phase interface morphology and growth of the shear layer show a good agreement with previous experimental observation. The disintegration mechanism of the gas-liquid interface is revealed from the simulation results, which is mainly due to the initial perturbation developing in the gas stream. Different disintegration patterns are discovered, which are wavy disintegration, roller type disintegration and breakdown of ligaments under different velocity ratios and density ratios. The shear layer thickness grows faster under higher pressure conditions, which indicates that the interface disintegration is more violent under higher pressure conditions. Increasing the density ratio will lead to higher frequency of the fundamental wave, which determines the first deformation and further breaking-down of gas-liquid interface close to the entrance of the flow.
  • Quaternion based linear time-varying model predictive attitude control for
           satellites with two reaction wheels
    • Abstract: Publication date: Available online 14 January 2020Source: Aerospace Science and TechnologyAuthor(s): Ali Golzari, Hossein Nejat Pishkenari, Hassan Salarieh, Taleb AbdollahiAbstractAttitude control of a satellite having only two reaction wheels is a challenging issue. To address this problem, previously published researches considered some simplifying assumptions on the satellites such as diagonality of the moment of the inertia matrix. On the other hand, in some works, the total angular momentum of the satellite is assumed to be zero. In this paper, a linear time-variant model predictive control (LTV MPC) is designed to control a satellite with two reaction wheels. This control method can be applied to a satellite with a non-diagonal inertial matrix in the presence of external torques, to rotate the satellite toward the desired directions in the space and orbit. The simulations results show that this method has a small amount of computational cost and enables an under-actuated satellite to perform large real-time maneuvers with acceptable accuracy.
  • Dynamic mode decomposition analysis of flow characteristics of an airfoil
           with leading edge protuberances
    • Abstract: Publication date: Available online 8 January 2020Source: Aerospace Science and TechnologyAuthor(s): Ming Zhao, Yijia Zhao, Zhengxian LiuAbstractA numerical investigation of the flow mechanisms and fluctuating aerodynamic performances of an airfoil with leading-edge protuberances is presented within post-stall regime at a Reynolds number of 1.2×105. In detail, a large eddy simulation (LES) has been conducted and then validated through quantitative comparisons with experimental and numerical results. Furthermore, dynamic mode decomposition (DMD) analysis has been carried out. Superior to the proper orthogonal decomposition (POD) method, DMD could extract mode with single-frequency characteristics. Therefore, the physical backgrounds of corresponding modes could be identified, and the flow control mechanisms could be uncovered, together with the influence on spanwise coherent structures. From the analysis at a streamwise section near laminar separation bubbles (LSBs), it has been found that the improved aerodynamic characteristics at peaks stem from the strong streamwise vortices induced by leading-edge protuberances, which lead to momentum transfer process from troughs to neighboring peaks. Meanwhile, DMD modes corresponding to the shear layers of LSBs have also been found, and the frequency characteristics are quantitatively depicted. On the other hand, from the analysis at selected lateral slices, the first DMD mode of smooth airfoil case is obviously related with Karman vortex shedding process according to the spatial distribution and frequency characteristics. Although wavy leading-edge causes the breakdown of spanwise coherent structures and the degradation of flow energy, Karman vortex shedding pattern could also be identified at particular troughs in modified airfoil case.
  • Breakup criterion for droplets exposed to the unsteady flow generated by
           an incoming aerodynamic surface
    • Abstract: Publication date: Available online 8 January 2020Source: Aerospace Science and TechnologyAuthor(s): P. Lopez-Gavilan, A. Velazquez, A. García-Magariño, S. SorAbstractAn experimental and theoretical study is presented on the problem of droplet breakup exposed to a continuously accelerating flow generated by an incoming aerodynamics surface. Droplet breakup experiments were carried out in a rotating arm facility. Droplet diameters were of the order of 1 mm. The maximum velocity of the airfoils located at the end of the rotating arm was 90 m/s. Droplet deformation was computed using a phenomenological model developed previously by the authors. The dynamics of this deformation was coupled to an instability model based on the growth of Rayleigh-Taylor waves at the droplet surface. It was found that, within the experimental uncertainty, breakup occurs when the instability wavelength approaches the droplet hydraulic diameter assuming that it flattens and deforms as an oblate spheroid. This fact allowed for the generation of a theoretical closed-form droplet deformation and breakup model that predicts the onset of breakup with discrepancies of about ±10% when compared to the experimental results. Finally, as an application case, this closed-form model is used to simulate an actual situation in which the objective is to investigate whether a series of droplets that are approached by an airfoil either impact on its surface, or break prior to collision, or break without colliding, or pass through undamaged.
  • Influence of the inlet distortion on fan stall margin at different
           rotational speeds
    • Abstract: Publication date: March 2020Source: Aerospace Science and Technology, Volume 98Author(s): Wenqiang Zhang, Sina Stapelfeldt, Mehdi VahdatiAbstractThe aim of this paper is to develop a reliable and accurate numerical strategy that can be used to study the effects of inlet distortions on the aerodynamic stability of fan blades in aero-engines. As an initial step towards achieving this goal, three-dimensional unsteady Reynolds-Averaged-Navier-Stokes (URANS) simulations were carried out to predict the influence of total pressure distortion on the loss of stall margin of a fan blade. NASA rotor 67, for which a significant amount of measured steady and unsteady data is available, was used for this study. It was observed that the size of the time step has a significant effect on the solution near stall and hence the stall margin of the blade.In the second part of this work, unsteady simulations were carried out to study the effects of rotational speed on the stall margin and stalled operating point of the blade. The results showed that for the same level and pattern of inlet distortion, the stall margin of the blade decreases as the corrected speed decreases. However, the in-stall total pressure losses decrease as the speed decreases. Finally, in our research it has become apparent that there is a big lack of public domain measured data for the cases with inlet distortion, and therefore, validated CFD results can be very helpful to other researchers in the field.
  • Probabilistic stability analysis of functionally graded graphene
           reinforced porous beams
    • Abstract: Publication date: Available online 24 January 2020Source: Aerospace Science and TechnologyAuthor(s): Kang Gao, Duy Minh Do, Ruilong Li, Sritawat Kitipornchai, Jie YangAbstractThis paper presents the first attempt to study the probabilistic stability characteristics of functionally graded (FG) graphene platelets (GPLs) reinforced beams by taking into account the multidimensional probability distributions, such as stochastic porosity and GPL distribution patterns as well as random material properties. For this purpose, a non-inclusive Chebyshev metamodel (CMM), which is implemented on deterministic analysis using discrete singular convolution (DSC) method with excellent computational efficiency and accuracy, is proposed and used to obtain both deterministic and probabilistic results including probability density functions (PDFs), cumulative density functions (CDFs), means and standard deviations of the critical buckling load. The present analysis is rigorously validated through direct comparisons against the results obtained by a direct quasi-Monte Carlo simulation (QMCS) method and those available in open literature. The influences of material properties, porosity distribution, GPL dispersion pattern and boundary condition on probabilistic buckling behaviour of the FG-GPL beam are comprehensively investigated. The global sensitivity analysis is also conducted. The results suggest that the critical buckling load of the FG-GPL beam is most sensitive to porosity distribution, followed by porosity coefficient and GPL weight fraction.
  • Safety analysis for the posfust reliability model under possibilistic
           input and fuzzy state
    • Abstract: Publication date: Available online 24 January 2020Source: Aerospace Science and TechnologyAuthor(s): Chunyan Ling, Zhenzhou Lu, Xiaobo Zhang, Suting ZhouAbstractThe structural input can be simply classified as probabilistic one and possibilistic one, and the structural state can be divided into binary state and fuzzy state. Combining different kinds of structural input and state, the reliability analysis model can be sorted into four types, i.e., the one based on probabilistic input and binary state (probist), the one based on probabilistic input and fuzzy state (profust), the one based on possibilistic input and binary state (posbist) and the one based on possibilistic input and fuzzy state (posfust). Various researches about the former three reliability models have been developed, whereas the study on the posfust reliability model is rare. Due to this fact, the fuzzy failure credibility is established to measure the safety degree of the posfust model, and it is a fuzzy safety index defined by the definite integral of the failure membership function. Then, to estimate the defined fuzzy failure credibility efficiently, two adaptive Kriging based methods are proposed. The first one is based on the definition of the fuzzy failure credibility, in which the lower and upper bounds of performance function are mainly concerned by the strategy of constructing the Kriging model. The second one combines the adaptive Kriging and fuzzy simulation algorithm, and it is based on the relationship between the defined fuzzy failure credibility and the failure credibility in the posbist reliability model. Several examples are provided to verify the rationality of the established fuzzy failure credibility, and the efficiency and accuracy of the proposed methods.
  • Global reliability sensitivity analysis based on state dependent parameter
           method and efficient sampling techniques
    • Abstract: Publication date: Available online 24 January 2020Source: Aerospace Science and TechnologyAuthor(s): Yushan Liu, Luyi LiAbstractIn order to efficiently assess the influence of input variables on the failure of the structural systems, an improved global reliability sensitivity analysis (SA) method is proposed in this paper. The new method is based on the state dependent parameter (SDP) method and the efficient sampling techniques. In the new method, the efficient sampling techniques are first used to generate samples that are more efficient for reliability analysis, and then the SDP method is further employed to estimate the global reliability sensitivity index by using the same set of samples as the reliability analysis. Two efficient sampling methods, e.g., importance sampling (IS) and truncated importance sampling (TIS), are employed in this paper, and the strategies of combining these methods with the SDP method for global reliability SA are discussed. Compared with the existing SDP method, the new method is more efficient for global reliability SA of structural systems. Three examples are used in the paper to demonstrate the efficiency and precision of the new methods.
  • Accurate control of Aerial Manipulators outdoors. A reliable and
           self-coordinated nonlinear approach
    • Abstract: Publication date: Available online 24 January 2020Source: Aerospace Science and TechnologyAuthor(s): J.Á. Acosta, C.R. de Cos, A. OlleroAbstractIn recent years, standard Unmanned Aerial Vehicles have been enhanced in order to execute manipulation tasks, named Aerial Manipulators and composed of an Unmanned Aerial Vehicle and a Robot Manipulator. These aerial vehicles are extremely demanding on the flight control system, which needs to keep flight stability while autonomously executing complex tasks. This work continues a recent line of research where a nonlinear control strategy is proposed to comply with outdoor high-level demands to achieve enhanced accuracy and safety. This solution combines nonlinear control subsystems with decentralized priorities but coordinating their relative movements to allow the aerial vehicle to accommodate itself while reaching the manipulator target. The complete strategy is composed of: a passive nonlinear dynamic controller for the UAV, an integral kinematic multi-task controller for the manipulator, and an optimizer to coordinate their relative movements. Theoretical stability results are reported, along with implementation-oriented add-ons and an extensive analysis in realistic simulations. These include aerodynamic effects (e.g. unsteady wind disturbances and rotor propulsion models), collision avoidance, grasping and a comparison with a standard strategy on a benchmark mission complex enough for validation.
  • Experimental study on ejector-to-ramjet mode transition in a divergent
           kerosene-fueled RBCC combustor with low total temperature inflow
    • Abstract: Publication date: Available online 24 January 2020Source: Aerospace Science and TechnologyAuthor(s): Lei Shi, Guojun Zhao, Yiyan Yang, Fei Qin, Xianggeng Wei, Guoqiang HeAbstractRBCC engine commonly fulfills the ejector-to-ramjet mode transition at a Mach number far below 4, and thus the divergent kerosene-fueled RBCC combustor will encounter tough technical challenges of completing the mode transition and maintaining sustained subsonic combustion at low total temperature inflow conditions. These important issues were studied by direct-connect tests in a divergent JP-10-fueled RBCC combustor at M∞=3 and T0=630 K, and two different rocket control methods, in terms of “stepped (switching off)” and “progressive (maintaining a significant reduction in the mass flow rate)” rocket control method, were validated. As a result: (1) Credible ignition could be obtained when the fuel was directly injected into the high temperature zone comprehensively affected by the rocket plume and its secondary combustion zone. (2) Self-sustained ramjet combustion in the stepped rocket control case required an excellent collaboration among the different combustion enhancement facilities (fuel struts and cavities) in the combustor. (3) Rocket-aided mode transition and ramjet combustion in the progressive rocket control case required less limitation on the combustor configuration, and permitted an improved reliability for the engine operation, although it sacrificed a small part of engine specific impulse.
  • On the mode transition of a double bypass variable cycle compression
    • Abstract: Publication date: Available online 24 January 2020Source: Aerospace Science and TechnologyAuthor(s): Baojie Liu, Ruoyu Wang, Xianjun YuAbstractBy establishing a three dimensional model of a double bypass variable cycle compression system, the flow patterns and matching characteristics of each working component during mode transition are investigated using numerical simulation. Results show that transition from the single bypass mode to double bypass mode by opening the mode selector valve (MSV) alone would increase the fan operation point while decreasing that of the core driven fan stage (CDFS) and the high pressure compressor (HPC). It would also incur outer bypass flow recirculation and bring radial inflow distortions to the CDFS. Deviations of the fan aerodynamic performance lie mainly in its aft stage, while the HPC first stage undertakes most of the inlet distortion. Reducing the bypass backpressure during mode transition is an effective way to alleviate the outer bypass flow recirculation but would further choke the CDFS. Closing the forward variable area bypass injector (FVABI) could raise the CDFS matching point so as to improve the stator performance without influencing the outer bypass ratio. It is recommended to decrease the bypass backpressure and close FVABI simultaneously in the real transition process.
  • Predicting distributed roughness induced transition with a four-equation
           laminar kinetic energy transition model
    • Abstract: Publication date: Available online 24 January 2020Source: Aerospace Science and TechnologyAuthor(s): Zaijie Liu, Yatian Zhao, Shusheng Chen, Chao Yan, Fangjie CaiAbstractA four-equation laminar kinetic energy transition model is developed to predict the distributed roughness induced transition using local variables. Based on the laminar kinetic energy transition model (kT-kL-ω model), three key improvements are devised in the work. Firstly, an additional transport equation for roughness amplification factor is combined with the laminar kinetic energy transition model. Secondly, the effective length scale is modified through the roughness amplification factor to consider the enhancement of the first and second unstable mode characteristic timescale. Additionally, for the sake of modeling the roughness effects in the full turbulent zone, the wall boundary condition for the specific turbulence dissipation rate is amended. Numerical results, including flat plate with distributed roughness, sharp biconic configuration with large roughness and hemisphere with different roughness heights, demonstrate that the proposed four-equation transition model is competent for accurate transition prediction at different roughness heights and Reynolds numbers. Besides, with the modification for the wall boundary condition of the specific turbulence dissipation rate, the present model outperforms the original model in simulating turbulent augmentations of skin friction and turbulent heating over rough surfaces. Thus, the present model has attraction and feasibility for simulating distributed roughness induced transition. While more physical mechanisms of roughness induced transition should be considered to further refine this four-equation laminar kinetic energy transition model.
  • Alternative approaches for UAV dead reckoning based on the immunity
    • Abstract: Publication date: Available online 24 January 2020Source: Aerospace Science and TechnologyAuthor(s): Mohanad Alnuaimi, Mario G. PerhinschiAbstractThe immunity paradigm has been recently investigated as a potential solution to the problem of unmanned aerial vehicles (UAV) navigation when external sensor systems and information of opportunity are not available. The effectiveness of specialized immunity cells in memorizing the characteristics of invading antigens, as well as the characteristics of the response of the host organism, or the antibodies, represent the source of inspiration for the development of an on-board intelligent system, an artificial immune system (AIS). The AIS is expected to provide corrections to a simple dead reckoning algorithm, such that autonomous trajectory tracking control of a UAV can be maintained over extended periods of time. Variables that can characterize the dynamic state of the vehicle and can be measured exclusively with on-board sensors are assimilated to the antigens, while the antibodies consist of the adjustments applied to the integration algorithm, such that accurate estimations of vehicle position and velocity are obtained. The AIS is built as a collection of instantaneous values of antigens and corresponding antibodies, referred to as artificial memory cells. The AIS is constructed using data collected under nominal conditions, when all sensors and systems function correctly and vehicle position and velocity are available. Two alternative approaches for building and operating the AIS are implemented and investigated in this paper. One relies on providing corrections to the output and the other to the input of the integration scheme. The AIS for both scenarios have been generated and tested using the West Virginia University unmanned aerial systems simulation environment. Both alternative approaches have been demonstrated to achieve similar performance in terms of significantly improving the vehicle position and velocity estimation algorithm and leading to desirable trajectory tracking and mission completion.
  • Numerical investigation of the error caused by the aero-optical
           environment around an in-flight wing in optically measuring the wing
    • Abstract: Publication date: March 2020Source: Aerospace Science and Technology, Volume 98Author(s): Yan Liu, Yingtao Yuan, Xiang Guo, Tao Suo, Yulong Li, Qifeng YuAbstractA major concern for in-flight wing deformation measurement using optical methods is the influence of the aero-optical environment around the wing because the quality of the images degrades due to variations of the index of refraction and then the deformation ameasurement errors will be introduced. By taking a typical 2D OA309 airfoil as the research object, the error caused by the aero-optical environment is analyzed and numerically evaluated to investigate the effects of Mach number, angle of attack, and the placement of the camera on optical wing deformation measurement. At first, the error caused by the aero-optical environment in stereoscopic camera setup is analyzed, and virtual displacement is proposed as an evaluation parameter. Then, based on CFD computation, the fourth-order Runge-Kutta ray-tracing method in combination with triangular mesh interpolation is presented to determine the virtual displacement under different conditions. The results indicate that the virtual displacement in the streamwise direction has the largest level around the leading edge and the Root-Mean-Square value over the wing surface is dependent linearly on the square of Mach number. The virtual displacement reaches the minimum value at −2° angle of attack; within the range of stall angle of attack, the further away from −2° angle of attack, the larger virtual displacement the corresponding aero-optical environment will produce. The virtual displacement increases as the camera-wing distance increases, but the growth rate is slowed down. The overall virtual displacement reaches the minimum when the camera viewing angle is close to 90°. The proposed computational method and results can be helpful to appropriately position the camera in the actual in-flight wing measurement and minimize errors caused by the aero-optical environment.
  • Hypersonic boost–glide vehicle strapdown inertial navigation system /
           global positioning system algorithm in a launch-centered earth-fixed frame
    • Abstract: Publication date: Available online 16 January 2020Source: Aerospace Science and TechnologyAuthor(s): Kai Chen, Jun Zhou, Fu-Qiang Shen, Han-Yan Sun, Hao FanAbstractTo suit the features of hypersonic vehicles and meet the requirement of their flight control system, we propose an integrated navigation algorithm in the launch-centered earth-fixed (LCEF) frame. First, we introduce a system mechanization for the strapdown inertial navigation algorithm in the LCEF frame and derive discrete update algorithms of strapdown inertial navigation attitude, velocity, and position in the LCEF frame. A coning effect compensation algorithm is deduced in the update algorithm of attitude, and a sculling effect compensation algorithm is deduced in the update algorithm of velocity. Then, we derive the attitude, velocity, and position error equations in the LCEF frame; we further derive a strapdown inertial navigation system (SINS) / global positioning system (GPS) integrated navigation filter state equation and measurement equation of the LCEF frame as well as introduce a simultaneous method of SINS/GPS integrated navigation. Finally, considering the typical hypersonic boost–glide vehicle as the object, the SINS/GPS algorithm is applied in a semi-physical simulation environment; verification results yield a position error less than 10 m, a velocity error less than 0.2 m/s, and an attitude error less than 0.05°.
  • Contact dynamics and relative motion estimation of non-cooperative target
           with unilateral contact constraint
    • Abstract: Publication date: Available online 16 January 2020Source: Aerospace Science and TechnologyAuthor(s): Yuchen She, Shuang Li, Jiaqian HuAbstractThis paper investigates the contact dynamics and the relative motion estimation problems between the target and the end-effector during the post-capturing operation. The unilateral contact constraint is taken into consideration, and several innovative methods have been proposed to properly estimate the relative displacement and velocity between the target and the manipulator's end-effector. Firstly, the contacting dynamics is profoundly analyzed by introducing several friction models into the system. Secondly, a new dynamical model is proposed by adopting the extra degree-of-freedom method, where a mapping is established between the position of the target and three new geometrical parameters. The mapping is then generalized in this paper to properly describe the relative motion for non-prismatic target. Thirdly, by considering the target to be completely non-cooperative, a relative motion measuring and estimation algorithm is proposed based on the computer vision and feature matching principle. Finally, the proposed methods are validated by computer simulation. The simulation results show that the new dynamical model can correctly describe the relative sliding motion between the target and end-effector. It has also been proved that the relative motion measuring algorithm can efficiently extract all the motion information, including the generalized position, velocity and acceleration of all extra degrees of freedom.
  • Nonlinear resonance of FG multilayer beam-type nanocomposites: Effects of
           graphene nanoplatelet-reinforcement and geometric imperfection
    • Abstract: Publication date: Available online 15 January 2020Source: Aerospace Science and TechnologyAuthor(s): Hu Liu, Han Wu, Zheng LyuAbstractIn this paper, the nonlinear dynamic response of a FG multilayer beam-type nanocomposite reinforced with graphene nanoplatelet (GNP) by considering the initial geometric imperfection is investigated on the basis of nonlocal strain gradient Euler-Bernoulli beam theory. Four patterns of GNP distribution incorporating the uniform distribution (UD) and O-, X-, and A- FG pattern distributions are taken into account and the effective elastic properties of the beam-type nanocomposite are evaluated in the framework of Halpin-Tsai scheme. The first-order vibrational mode is employed to represent the initial geometric imperfection of the nonlinear FG beam-type nanocomposite. Correspondingly, the nonlinear amplitude-frequency response of the imperfect FG multilayer beam-type nanostructures subjected to the excitation resonance is analyzed with the aid of multiple scale method. Firstly, the present model is validated with a comparison of two previous works. Then, a comprehensive investigation is conducted to evaluate the effects of GNP distributed pattern, weight fraction of GNPs, geometric imperfection amplitude, boundary condition, excitation amplitude, nonlocal and strain gradient size scale parameters on the nonlinear frequency-response of FG multilayer beam-type nanostructures. The current work is beneficial for the application of GNP as reinforcement to enhance mechanical performances of nanostructures.
  • Active fault-tolerant attitude tracking control with adaptive gain for
    • Abstract: Publication date: Available online 15 January 2020Source: Aerospace Science and TechnologyAuthor(s): Hui Hu, Lei Liu, Yongji Wang, Zhongtao Cheng, Qinqin LuoAbstractIn this paper, an active fault-tolerant control (AFTC) method with adaptive gain is proposed for spacecrafts which subject to model uncertainty, external disturbance and actuator fault. A novel adaptive-gain finite-time observer is designed to solely estimate fault in the presence of disturbance for fault diagnosis. By incorporating the dual-layer gain adaption scheme into the nonlinear integral sliding mode control (ISMC), a reconfigurable controller is designed to accommodate fault. Due to the reaching phase elimination of ISMC, the gain tracking to the unknown bounded disturbance can be accelerated while it is maintained that the control gain will not be overestimated and the bound knowledge of fault/disturbance is not required. Simulation results verify the effectiveness and improvement of the proposed FTC method.
  • On the low-velocity impact responses of auxetic double arrowed honeycomb
    • Abstract: Publication date: Available online 15 January 2020Source: Aerospace Science and TechnologyAuthor(s): Qiang Gao, Wei-Hsin Liao, Liangmo WangAbstractAuxetic structures with negative Poisson's ratio have been widely studied due to their unique mechanical properties, especially the dynamic properties under low-velocity impact. In this paper, the dynamic properties of an excellent candidate for engineering application, double arrowed honeycomb (DAH), are analyzed thoroughly. The deformation patterns of the DAH under high and low velocity impact are compared, which shows there is a great difference between them. Analytical models for predicting the dynamic strength at the crushing/supporting end of DAH are developed based on cellular collapse mechanism. It shows that the crushing strength rises with the increase of the relative density and impact velocity. Numerical simulations are also performed to verify the accuracy of the analytical models. It is found that there is good agreement between the analytical and numerical predictions for the dynamic characteristics. Also, the energy absorptions of the DAH are also thoroughly studied based on the analytical models.
  • Tip leakage flow and aeroacoustics analysis of a low-speed axial fan
    • Abstract: Publication date: Available online 14 January 2020Source: Aerospace Science and TechnologyAuthor(s): Bo Luo, Wuli Chu, Haoguang ZhangAbstractAxial fans are widely used in the aerospace field, and new regulations and environmental concerns are prompting manufacturers to design efficient low-noise axial fans. The aerodynamic performance and acoustic emissions of axial fans are substantially affected by the unavoidable tip clearance. Herein, a systematic analysis is performed on axial fans with different tip-clearance sizes to gain a clear understanding of the characteristics of tip leakage flow and investigate the generation mechanism of aeroacoustics. The reliability and accuracy of the numerical predictions are successfully validated through a comparison with experimental data. The unsteady pressure information on the blade surface is examined to clarify the main noise sources. The results show that the enhanced intensity of the blade tip vortex and thereby the enhanced interaction with the blade surface are the main drivers for the extra broadband noise when the tip clearance is increased. For the low-speed axial fans, the low-frequency broadband noise below 600 Hz is mainly related to the blade tip vortex and the interaction of the boundary layer instabilities with the leading edge and trailing edge, whereas the broadband noise above 1200 Hz is mainly due to the turbulent boundary layer fluctuations.
  • Numerical analysis on combustion instabilities in end-burning-grain solid
           rocket motors utilizing pressure-coupled response functions
    • Abstract: Publication date: Available online 14 January 2020Source: Aerospace Science and TechnologyAuthor(s): Shixiang Ji, Bing Wang, Dan ZhaoAbstractThe combustion instability has been one of severe problems suffered by solid rocket motors (SRMs) for a long term. This paper proposes a burning rate model utilizing the pressure-coupled response function for description of the spatiotemporal burning of AP-HTPB composite propellant grains in SRMs. The model is successfully incorporated to the axisymmetric internal ballistic simulation by the source terms of gas-phase governing equations. The prediction of propellant combustion connected with acoustic pressures can be realized by means of an in-house high-order numerical solver, in which the numerical fluxes are reconstructed by the fifth-order WENO scheme and the viscous terms are discretized by a sixth-order compact scheme. The thermoacoustic combustion instabilities in the longitudinal modes are triggered by a burning rate pulse imposed to the steady flow. The analysis indicates that the pressure oscillations grow as a primary symptom of combustion instabilities. The influences of several factors on production of the instability symptom are discussed. It is shown that the pressure-coupled response function, the pressure index, and the reaction heat of the propellant, and the specific heat ratio of the burnt gas promote the pressure oscillation growth process evidently, but the magnitude and the imposed region of burning rate pulses have less effects.
  • Transitional boundary layer study over an airfoil in combined pitch-plunge
    • Abstract: Publication date: Available online 13 January 2020Source: Aerospace Science and TechnologyAuthor(s): Hassan Akhlaghi, Mohammad-Reza Soltani, Mohammad-Javad MaghrebiAbstractTransitional boundary layer over an airfoil in the combined pitch-plunge oscillating motions at low Reynolds number is experimentally investigated. The study is based on the data obtained from both surface hot-film and surface static pressure. Different oscillation zones prior to, within, and beyond the static stall angle of attack of the airfoil are considered. The hysteresis loops for the laminar separation bubble, transition, relaminarization points, as well as the airfoil lift coefficient are investigated and are compared for different types of dynamic motions. Two different types of combined pitch-plunge motions, constructive and destructive, are considered. In the constructive motion, the amplitude of the equivalent angle of attack increases, while in the destructive one the reverse happens. From the present it is concluded that clockwise or counterclockwise direction of the hysteresis loops is determined by the apparent mass and by the wake effects when oscillating below and within the static stall angles of attack for this airfoil. It is further presumed that the apparent mass effect is more pronounced for the combined motions for the lower pitch amplitude cases. Moreover, massive flow separation and dynamic stall vortices when oscillating within and beyond the static stall angles of attack are believed to play a significant role in the hysteresis loops. It is found that the destructive combined motion results in a large hysteresis of the flow phenomena over the surface of the model. The trends of the hysteresis loops for laminar separation points may differ compared to the transition points in combined pitch-plunge motion.
  • Wind estimation using quadcopter motion: A machine learning approach
    • Abstract: Publication date: Available online 13 January 2020Source: Aerospace Science and TechnologyAuthor(s): Sam Allison, He Bai, Balaji JayaramanAbstractIn this article, we study the well known problem of wind estimation in atmospheric turbulence using small unmanned aerial systems (sUAS). We present a machine learning approach to wind velocity estimation based on quadcopter state measurements without a wind sensor. We accomplish this by training a long short-term memory (LSTM) neural network (NN) on roll and pitch angles and quadcopter position inputs with forcing wind velocities as the targets. The datasets are generated using a simulated quadcopter in turbulent wind fields. The trained neural network is deployed to estimate the turbulent winds as generated by the Dryden gust model as well as a realistic large eddy simulation (LES) of a near-neutral atmospheric boundary layer (ABL) over flat terrain. The resulting NN predictions are compared to a wind triangle approach that uses tilt angle as an approximation of airspeed. Results from this study indicate that the LSTM-NN based approach predicts lower errors in both the mean and variance of the local wind field as compared to the wind triangle approach. The work reported in this article demonstrates the potential of machine learning for sensor-less wind estimation and has strong implications to large-scale low-altitude atmospheric sensing using sUAS for environmental and autonomous navigation applications.
  • Fuzzy controller design of micro-unmanned helicopter relying on improved
           genetic optimization algorithm
    • Abstract: Publication date: Available online 13 January 2020Source: Aerospace Science and TechnologyAuthor(s): Yanpeng Hu, Yanping Yang, Shu Li, Yaoming ZhouAbstractIn this paper, the dynamic model of unmanned helicopter is modeled by frequency domain identification method. An adaptive fuzzy Proportion Integration Differentiation (PID) controller is established, with its design carried out from the aspects of the input and output fuzzy subset, the membership function selection, the fuzzy rule generation, as well as the fuzzy reasoning. An improved genetic algorithm is employed to optimize the initial expert empirical fuzzy rules, which avoids the traditional method from falling into the local optimal solution in the process of optimization. Specifically, adaptive crossover and mutation probability are adopted to accelerate the convergence speed of the genetic algorithm. We then apply the proposed algorithm to optimize the membership function and fuzzy control rules of the fuzzy controller. Simulation results indicate that our adaptive fuzzy PID has better control effect and anti-interference ability than the traditional PID control law. Additionally, our adaptive fuzzy PID controller is verification by flight test thought AF25B unmanned helicopter platform.
  • Thermal protection of a hypersonic vehicle by modulating stagnation-point
           heat flux
    • Abstract: Publication date: Available online 11 January 2020Source: Aerospace Science and TechnologyAuthor(s): Quan Han, Chengdong Sun, Yi Tao, Zhongwu Li, Yan Zhang, Yunfei ChenAbstractThe heat flux at the stagnation-point is usually used to evaluate the heat transfer efficiency between high-temperature gas and a hypersonic flying vehicle. Controlling the stagnation-point heat flux is of critical important for the design of a thermal protection system for a hypersonic flying vehicle. In this work, we demonstrate that the nonequilibrium molecular dynamic (NEMD) method can well describe the shock wave and the stagnation-point heat flux. In particular, our results show that the heat flux is more sensitive to the change of the peak amplitude of the hypersonic flow in the bimodal velocity probability density function (PDF) at the stagnation point. Decreasing the peak amplitude of the hypersonic flow in the PDF can effectively reduce the stagnation-point heat flux. Three types of the outer shapes for the flying objects are employed to compare the heat flux at the stagnation-point. The NEMD simulation results confirmed that the heat flux can be decreased 16.8% for the spherical cavity outer shape compared with the flat shape in the leading edge.
  • Integrated guidance and control system design for laser beam riding
           missiles with relative position constraints
    • Abstract: Publication date: Available online 10 January 2020Source: Aerospace Science and TechnologyAuthor(s): Qian Peng, Jianguo Guo, Jun ZhouAbstractA finite time disturbance observer-based barrier Lyapunov function integrated guidance and control law (FTDO-BLF) is proposed for laser beam riding missiles with relative position constraints in this paper. Firstly, the laser beam riding missile's model is formulated, in which the mismatched uncertainties including disturbances and system perturbations are considered. Secondly, the relative position constraints are introduced to design the barrier Lyapunov function integrated guidance and control law (BLF), and the finite time disturbance observer (FTDO) is utilized for estimating and compensating the mismatched uncertainties to the BLF. Thirdly, the stability of FTDO-BLF is proved, and its superiority is demonstrated by the simulation completed under different cases. Consequently, the proposed FTDO-BLF attenuates the disturbances effectively and is robust against white noise without violating the relative position constraints.
  • Evaluation of modelling parameters for computing flow-induced noise in a
           small high-speed centrifugal compressor
    • Abstract: Publication date: Available online 10 January 2020Source: Aerospace Science and TechnologyAuthor(s): S. Sharma, J. García-Tíscar, J.M. Allport, S. Barrans, A.K. NicksonAbstractDevelopments in computing infrastructure and methods over the last decade have enhanced the potential of numerical methods to reasonably predict the aerodynamic noise. The generation and propagation of the flow induced noise are aerodynamic phenomena. Although the fluid flow dynamics and the resultant acoustics are both governed by mass and momentum conservation equations, former is of convective and/or diffusive nature while the latter is propagative showing insignificant attenuation due to viscosity except for small viscothermal losses. Aeroacoustic modelling of systems with intricate geometries and complex flow is still not mature due to challenges in the accurate tractable representation of turbulent viscous flows. Therefore, state-of-the-art for computing flow-induced noise in small centrifugal compressors is reviewed and critical evaluation of various parameters in the numerical model is undertaken in this work. The impact of various turbulence formulations along with corresponding spatial and temporal resolutions on performance and acoustic predictions are quantified. The performance predictions are observed to be within 1.5% of the measured values irrespective of turbulence and timestep parameters. The noise generated by the impeller is observed to be reasonably correlated with the measurements and the absolute values of the sound pressure levels along with decay rates predicted by LES and SBES formulations are better than the similar predictions from DES and URANS formulations. The impact of timestep size is observed and is determinant of the frequency up to which spectra can be appropriately resolved. Furthermore, results emphasise the importance of high spatial resolution for scale resolving turbulence formulations to yield better results and the information can be used to select appropriate numerical configuration considering time and accuracy trade-offs.
  • The influence of the wedge shock generator on the vortex structure within
           the trapezoidal cavity at supersonic flow
    • Abstract: Publication date: Available online 10 January 2020Source: Aerospace Science and TechnologyAuthor(s): Zhixiong Li, Tran Dinh Manh, M. Barzegar Gerdroodbary, Nguyen Dang Nam, R. Moradi, Houman BabazadehAbstractThe cavity flame holder is a conventional technique for efficient fuel mixing and auto-ignition inside the Scramjet engine. In this study, a computational method is applied to study the impact of the wedge shock generator on the fuel mixing performance inside the combustor at the supersonic flow of M=4. The main focus of the current work is to compare the vortex structure formed inside the cavity in various conditions. To perform simulations, a three-dimensional model with a structure grid is produced to fine reliable solution for our study. To capture the real flow structure, the SST turbulence model is applied for obtaining viscosity. The obtained results demonstrate that the presence of the wedge shock generator increase the mainstream entrains into the cavity. This declines the temperature inside the cavity and strengthens the fuel mixing rate which is highly significant for the efficient combustion process within the scramjet engine. According to our findings, secondary circulation weakens by the amplification of the oblique shock.
  • Robust inertial navigation system/ultra wide band integrated indoor
           quadrotor localization employing adaptive interacting multiple
           model-unbiased finite impulse response/Kalman filter estimator
    • Abstract: Publication date: Available online 9 January 2020Source: Aerospace Science and TechnologyAuthor(s): Yuan Xu, Yuriy S. Shmaliy, Xiyuan Chen, Yueyang Li, Wanfeng MaAbstractA new adaptive interacting multiple model (IMM)-based unbiased finite impulse response (UFIR)/Kalman filter (KF) algorithm is proposed to provide accurate and robust mini quadrotor position information in indoor environments. The IMM-UFIR/KF algorithm is designed based on an integrated localization system, which combines recent data obtained from the inertial navigation system (INS) and ultra wide band (UWB) unit. In this scheme, adaptive UFIR filter and Kalman filters operate in parallel. The outputs of both filters are fused with probabilistic weights to correct the INS position, thereby reaching the best accuracy available from INS and UWB. It is shown experimentally that the output of the designed adaptive IMM-UFIR/KF-based localization system, which takes advantages of the KF accuracy and the UFIR filter robustness, ranges close to the most accurate estimator at each time instant.
  • Numerical simulation of turbulent thermal boundary layer and generation
           mechanism of hairpin vortex
    • Abstract: Publication date: Available online 9 January 2020Source: Aerospace Science and TechnologyAuthor(s): Heng Li, Duo Wang, Hongyi XuAbstractThree-dimensional flow fields are computed along a ribbed flat-plate by direct numerical simulation. Inlet Reynolds number defined by the rib height is at Re=1200 and Mach number is at Ma=0.2. The detailed bypass transition process and its generation mechanisms are discussed and analyzed. Results present the relatively larger skin-friction coefficients in the bypass transition region. Moreover, the Reynolds stresses and turbulent kinetic energy are stronger in the bypass transition region, giving rise to a larger Nusselt number distribution. Various large-scale vortices are found downstream, where the momentum thickness Reynolds number is in the range of 250
  • Development of k-R turbulence model for wall-bounded
    • Abstract: Publication date: Available online 9 January 2020Source: Aerospace Science and TechnologyAuthor(s): Yang Zhang, Md Mizanur Rahman, Gang ChenAbstractA new low-Reynolds-number turbulence model is formulated based on the turbulent kinetic energy k and eddy-viscosity parameter R=k2/ϵ. In the derivation, most of the diffusion terms emerging from the transformation have been preserved in the proposed model, maintaining the closest relationship with its parent k-ϵ model; the near-wall viscous balance is guaranteed by including a far-wall damping function. The turbulent structure parameter (i.e., Bradshaw-relation) that may enhance the predicted accuracy for non-equilibrium flows, is incorporated into the current model. The coefficients and functions are constructed such as to preserve the anisotropic characteristics of turbulence, encountered in rotational and irrotational flows. Fully-developed turbulent channel and flat-plate flows are computed to validate the model ability in replicating the near-wall turbulence. Two-dimensional asymmetric plane diffuser and airfoil flows are simulated to verify the model accuracy in capturing non-equilibrium flows with separation and reattachment. Computation of a three-dimensional wing with shock-wave is presented, rectifying the model competency in predicting large adverse-pressure-gradient flows. Furthermore, model predictions are compared with other two frequently-used turbulence models; a good correlation is obtained between the current model and experimental data.
  • Fuzzy adaptive nonlinear information fusion model predictive attitude
           control of unmanned rotorcrafts
    • Abstract: Publication date: Available online 9 January 2020Source: Aerospace Science and TechnologyAuthor(s): Qingzheng Xu, Zhisheng Wang, Yue LiAbstractThis study proposes a novel nonlinear information fusion model predictive control (NIFMPC) approach for the attitude control of unmanned rotorcraft subject to external wind disturbance. In the proposed control structure, the fusion estimation of control input is obtained by fusing the soft constraint information and the hard one via the nonlinear information fusion estimation method. The output error convergence of the closed-loop system is analyzed using the Lyapunov stability theory. In order to avoid input saturation, the fuzzy logic system (FLS) is adopted to design the adaptive fuzzy gain-scheduling (AFGS)-NIFMPC, in which the soft constraint information related to control action are adaptively scheduled according to fuzzy rules, with control input and its rate of change as FLS inputs and soft constraint information as FLS outputs. Finally, simulation results are presented to demonstrate the effectiveness and robustness of the proposed control strategies.
  • Experimental study on flame/flow dynamics in a multi-nozzle gas turbine
           model combustor under thermo-acoustically unstable condition with
           different swirler configurations
    • Abstract: Publication date: Available online 9 January 2020Source: Aerospace Science and TechnologyAuthor(s): Can Ruan, Feier Chen, Tao Yu, Weiwei Cai, Xinling Li, Xingcai LuAbstractThis paper presents an experimental study on the combustion stability characteristics and flame/flow dynamics in a multi-nozzle, lean premixed prevaporized (LPP), swirl-stabilized gas turbine model combustor with different swirler configurations in the presence of self-excited combustion instability. The flame structure was characterized using high-speed OH⁎ chemiluminescence imaging and the flow field across the centerline of three interacting flames was measured by high-speed planar Particle Image Velocimetry (PIV). Two sets of the swirler configurations were considered in this paper, featuring different combinations of swirl rotational directions. The first one consisted of three co-rotating swirlers, whereas the central swirler in the second configuration was replaced by a swirler with counter-rotation direction. These two configurations were termed as COS and CNS combustors respectively in the rest of the paper. It was found that these two combustors exhibited similar stable and unstable operating domains in terms of different equivalence ratios and inlet air velocities. At the same test condition, the amplitude of dominant instability of COS combustor was 130 dB, which was stronger than that of CNS combustor (120 dB). Phase-averaged PIV measurements showed that both COS and CNS combustor featured three recirculation zones downstream the swirlers, and high axial velocity was present after the merging of the adjacent flames. These two flow structures varied periodically, but in different manners for COS and CNS combustors. Phase-averaged OH⁎ chemiluminescence images indicated that the flame was primarily anchored in recirculation zones close to the swirler and most of the heat release was found to occur in flame interaction regions, where large-scale reaction intensity variations occurred. Furthermore, a greater phase delay between heat release rate and acoustic pressure was observed in CNS combustor, which contributed to a weaker instability comparing with that in the COS combustor.
  • Performance analysis of rotatable energy system of high-altitude airships
           in real wind field
    • Abstract: Publication date: Available online 9 January 2020Source: Aerospace Science and TechnologyAuthor(s): Weiyu Zhu, Yuanming Xu, Jun Li, Lanchuan ZhangAbstractEnergy output and consumption play a decisive role in the endurance of high-altitude airships. Although rotatable energy system (RES) based on the sun tracker is beneficial to the improvement of output performance of high-altitude airships, the effects of wind velocity and airship attitude angles on energy consumption are often ignored. The aims of this paper are to analyze the performance of RES of high-altitude airships and to obtain a favorable operating strategy in real wind field. The energy output models of high-altitude airships using RES and fixed energy system (FES) are presented to study the effect of yaw angle on the output performance. Based on the energy consumption model, the energy consumption of high-altitude airships with the change of wind velocity is calculated. The optimization model and operation strategy to obtain the maximum difference between energy output and consumption are also proposed, and a case study of real wind field is conducted to verify the strategy. The results show that the effects of yaw angle on power output and rotation angle of RES of high-altitude airships are remarkable especially at high latitudes. Moreover, considering the energy consumption from wind resistance, the optimal operating yaw angle of high-altitude airships using FES is to keep upwind angle 0°, and for high-altitude airships using RES, it is feasible to maximize the difference between energy output and consumption by increasing upwind angle of airships, especially when the wind speed is less than 6 m/s.
  • Infinite horizon model predictive control tracking application to
    • Abstract: Publication date: Available online 8 January 2020Source: Aerospace Science and TechnologyAuthor(s): William B. Greer, Cornel SultanAbstractA novel formulation for infinite horizon model predictive control that is general and can accurately approximate and minimize the control cost is proposed. This method allows output tracking over a finite horizon when constraints are considered and equilibrium tracking in the linearized dynamics after that for the infinite horizon when constraints are not considered. The method allows prescribing weights simultaneously on the output tracking error, control rates, and control energy. Methods of weight selection are discussed to ensure strong cost convexity for reliable optimization. The control method is used to bring a helicopter that is far from a ship to a stable flight condition that is much closer to the ship. This is very important for helicopter ship landing automation.
  • Experimental study on the flow field distribution characteristics of an
           open-end swirl injector under ambient pressure
    • Abstract: Publication date: Available online 8 January 2020Source: Aerospace Science and TechnologyAuthor(s): Xiujie He, Chen Chen, Yang Yang, Zhihui YanAbstractTo investigate the influence of ambient pressure on the atomization characteristics of an open-end swirl injector, the spray morphology, droplet diameter distribution and velocity distribution were measured under ambient pressures of 0.1 MPa, 0.5 MPa, 1.0 MPa, 1.5 MPa and 2.0 MPa using a high-speed camera and Phase Doppler Interferometer (PDI). As the ambient pressure increased, the spray region and intact liquid film area decreased, and the spray angle, spray width and breakup length shortened. Increased ambient pressure made droplet SMD larger and the SMD curves flatter. The droplet Sauter Mean Diameter (SMD) at each measuring point along the radial direction increased gradually and then decreased slightly, but there was no significant difference between different measuring sections. Ambient pressure had a more significant influence on axial velocity than the tangential velocity. As the ambient pressure increased, the axial velocity decreased, and the region where droplets had negative axial velocity was compressed backward to the injector outlet. Additionally, the tangential velocity near the outlet fluctuated greatly along the radial distribution. As the axial distance increased, the velocity curves became much flatter.
  • Investigation on effects of shock wave on vortical wake flow in a turbine
           nozzle cascade
    • Abstract: Publication date: Available online 8 January 2020Source: Aerospace Science and TechnologyAuthor(s): Ben Zhao, Mingxu Qi, Hong Zhang, Xin ShiAbstractVortical wake flow and its interaction with shock waves in a transonic turbine nozzle cascade were investigated by experimental and numerical methods. Measurements including aerodynamic performance and Schlieren photography were conducted on a validating turbine nozzle cascade. Numerical simulations were performed on both the validating turbine cascade and a single-passage turbine nozzle using two numerical schemes, the RANS and the detached eddy simulation (DES) simulations. The comparison of the wake flow patterns indicated that the DES calculation was more adaptable in simulating the vortical wake flow than the RANS method. Based on the DES results, it was found that, after crossing the shock wave, the vortical wake flow had three changes in its flow characteristic: 1) the vorticity increase in the vortex core; 2) the reduction of the distance between two adjacent vortices with the same direction; 3) the deflection of the entire vortex structure including the vortex core and its preparation stage. Those changes were explained by aerodynamic principles. To further validate the findings, the investigation was carried out at a subsonic condition, and the disappearance of those changes was extra evidence to support that the interaction of the shock wave contributes to the vortex changes.
  • Online trajectory optimization for power system fault of launch vehicles
           via convex programming
    • Abstract: Publication date: Available online 8 January 2020Source: Aerospace Science and TechnologyAuthor(s): Yuan Li, Baojun Pang, Changzhu Wei, Naigang Cui, Yongbei LiuAbstractThis paper presents an online trajectory optimization algorithm for launch vehicles based on convex programming to ensure flight safety in case of power system fault. Due to high complexity of the power system, the engine may break down during the flight, causing significant decrease of thrust or energy. In this case, the nominal trajectory will be infeasible as the dynamical model and energy state is different from the normal status, thus the online trajectory optimization and re-planning are considered. For different kinds of engine failures, different terminal orbital constraints are proposed. When the mass flow rate of fuel decreases, the energy loss is little but the dynamical model changes obviously, so the location of the injection point cannot be guaranteed. In this case, the terminal orbital elements are constrained except the true anomaly, so that the payload of launch vehicles can still settle into the nominal orbit, and the true anomaly is optimized for minimum fuel consumption. As for the energy-loss failure, the strategy to change the target orbit is proposed considering the requirement of launch mission and subsequent orbit transfer insertion. The terminal constraints are proposed analytically in this paper. In order to solve the nonconvex trajectory optimization problem accurately and rapidly, the optimization problem is transformed into convex optimization problems by various convexification techniques, including the lossless convexification and successive convexification. Finally, the high efficiency and accuracy of the proposed algorithm is verified by numerical experiments. The algorithm proposed in this paper has potential applications in onboard trajectory optimization and re-planning of launch vehicles in case of power system fault to ensure the accomplishment of the launch mission.
  • Over-expanded separation transitions of single expansion ramp nozzle in
           the accelerating and decelerating processes
    • Abstract: Publication date: Available online 8 January 2020Source: Aerospace Science and TechnologyAuthor(s): Yang YuAbstractFlow separation is a basic fluid-dynamics phenomenon that occurs in supersonic nozzles at low pressure ratios. In the over-expanded flowfield of nozzles, the subsonic area is formed, such as separation bubble, recirculation zone, and flow downstream Mach reflection. In static ambient, the recirculation zone is opened with the environment. When the external flow is not still, it will interact with the recirculation zone. As a result, the external flow will make significant influences on the over-expanded separation flowfield. This paper aims to present a detailed description of separation patterns transitions caused by external Mach number variation, analyze the mechanism of separation transitions. Accelerating and decelerating processes of a long flap single expansion ramp nozzle (SERN) is investigated by numerical simulation. Separation transitions and the performance of SERN influenced by the separation transitions have been discussed. In the accelerating process, free shock separation (FSS) is a stable separation pattern and lasts for a long time. The decelerating process is not just the inverse process because a new over-expanded flowfield pattern appears in the decelerating process. Meanwhile, there is an apparent hysteresis between the accelerating and decelerating processes.
  • Scaling laws and similarity models for the preliminary design of
           multirotor drones
    • Abstract: Publication date: Available online 8 January 2020Source: Aerospace Science and TechnologyAuthor(s): M. Budinger, A. Reysset, A. Ochotorena, S. DelbecqAbstractMultirotor drones modelling and parameter estimation have gained great interest because of their vast application for civil, industrial, military and agricultural purposes. At the preliminary design level the challenge is to develop lightweight models which remain representative of the physical laws and the system interdependencies. Based on the dimensional analysis, this paper presents a variety of modelling approaches for the estimation of the functional parameters and characteristics of the key components of the system. Through this work a solid framework is presented for helping bridge the gaps between optimizing idealized models and selecting existing components from a database. Special interest is given to the models in terms of reliability and error. The results are compared for various existing drone platforms with different requirements and their differences discussed.
  • Glide guidance for reusable launch vehicles using analytical dynamics
    • Abstract: Publication date: Available online 7 January 2020Source: Aerospace Science and TechnologyAuthor(s): Hongyu Zhou, Xiaogang Wang, Naigang CuiAbstractThis paper proposes a novel glide guidance law for the reusable launch vehicle (RLV), in which longitudinal and lateral paths are separately designed to quickly derive guidance commands. First, an altitude profile is designed in the longitudinal plane. Using this profile, the terminal altitude, flight path angle, and position are naturally met and the quasi-equilibrium glide condition (QEGC) is easily satisfied. Second, analytical solutions of altitude, velocity, and flight path angle are deduced in the glide phase; so the values of path constraints and performance index can also be analytical calculated. Third, a new concept of virtual target is proposed to design the lateral motion. By analytically determining the virtual target with the solutions in the glide phase, the cross-range is online adjusted to control the dissipation in velocity. Finally, the guidance law is proposed based on these analytical solutions. In the guidance law, longitudinal and lateral motions are controlled by online updating the profile and revising the virtual target, respectively. Integration-based predictors are unnecessary in comparison with conventional guidance laws, so a short guidance period can be used to improve the robustness. The effectiveness and the robustness of the proposed method are demonstrated with various scenarios and Monte Carlo simulation, respectively.
  • Combustion characterization of a CH4/O2 rapid mixed swirl torch igniter
           for hybrid rocket motors
    • Abstract: Publication date: Available online 3 January 2020Source: Aerospace Science and TechnologyAuthor(s): Yi Wu, Zixiang Zhang, Fuwen Liang, Ningfei WangAbstractCombustion characteristics of a rapid mixed swirl torch igniter of CH4/O2 for hybrid rocket motors has been experimentally investigated. The igniter torch consists of four tangential slits circumferentially equispaced on the internal micro-combustion chamber allows tangential injection of O2 gas, meanwhile the CH4 flows into the main channel and interacts with four tangential O2 injections flow tgo generate a swirl mixing of CH4 and O2. The ignitability of this swirl torch igniter was verified by large number of experiments in variation of equivalence ratio and total flowrate of CH4/O2. The flame structure of the igniter and effect of main oxidizer injection acting as a cross flow on the flame has been investigated by using OH* and CH* chemiluminescences techniques. It is found that the torch igniter can reliably ignite instantaneously in the range of equivalence ratio 0.2∼1.4 and the effect of main oxidizer injection on the flame is tiny and can be neglected. In addition, the performance of this rapid mixed swirl flame has been investigated by using a lab scaled hybrid rocket motor and compared with the same hybrid rocket motor ignited by catalytic bed. The performance of this igniter and its comparison with catalytic bed ignition will be discussed and analyzed in detail.
  • Reinforcement learning in dual-arm trajectory planning for a free-floating
           space robot
    • Abstract: Publication date: Available online 3 January 2020Source: Aerospace Science and TechnologyAuthor(s): Yun-Hua Wu, Zhi-Cheng Yu, Chao-Yong Li, Meng-Jie He, Bing Hua, Zhi-Ming ChenAbstractA free-floating space robot exhibits strong dynamic coupling between the arm and the base, and the resulting position of the end of the arm depends not only on the joint angles but also on the state of the base. Dynamic modeling is complicated for multiple degree of freedom (DOF) manipulators, especially for a space robot with two arms. Therefore, the trajectories are typically planned offline and tracked online. However, this approach is not suitable if the target has relative motion with respect to the servicing space robot. To handle this issue, a model-free reinforcement learning strategy is proposed for training a policy for online trajectory planning without establishing the dynamic and kinematic models of the space robot. The model-free learning algorithm learns a policy that maps states to actions via trial and error in a simulation environment. With the learned policy, which is represented by a feedforward neural network with 2 hidden layers, the space robot can schedule and perform actions quickly and can be implemented for real-time applications. The feasibility of the trained policy is demonstrated for both fixed and moving targets.
  • Automatic modeling of aircraft external geometries for preliminary design
    • Abstract: Publication date: Available online 3 January 2020Source: Aerospace Science and TechnologyAuthor(s): Agostino De Marco, Mario Di Stasio, Pierluigi Della Vecchia, Vittorio Trifari, Fabrizio NicolosiAbstractThis article introduces a high-fidelity geometry definition methodology enabling Multidisciplinary Design, Analysis and Optimization (MDAO) of aircraft configurations. All definitions and functional features have been implemented within the JPAD software, a Java-based computing library for aircraft designers, which provides a dedicated geometric modeling module called JPADCAD. The geometric module, that comes as an application programming interface (API) built on top of the Open CASCADE Technology solid modeling kernel, is conceived for the automatic production of parametric aircraft CAD geometries. The tool allows the definition of input geometries for low-fidelity as well as high-fidelity aerodynamic analyses, hence proves to be a key factor in the entire MDAO process, particularly in conceptual or preliminary design analysis workflows. The main goal of such a geometric library remains ease of use and support for automation to minimize unnecessary or repetitive human effort.The backbone of the presented methodology is the parametric definition of a generic commercial transport aircraft configuration that translates into software data structures and functionalities of CAD surface modelers. These aspects are discussed in the first part of the article. The second part presents a use case example of the geometric modeling API, where an automated aerodynamic analysis workflow is used to construct a prediction model for canard-wing configurations.
  • Dynamic buckling optimization of laminated aircraft conical shells with
           hybrid nanocomposite martial
    • Abstract: Publication date: Available online 3 January 2020Source: Aerospace Science and TechnologyAuthor(s): Shun-Peng Zhu, Reza Kolahchi, Behrooz Keshtegar, Nguyen-Thoi TrungAbstractDynamic buckling optimization in laminated truncated nanocomposite conical aircraft shell in moisture and temperature environments as well as magnetic fields is considered in this article. The structural layers are hybrid nanocomposite consist of polymer, carbon nanotubes (CNT) and carbon fibers based on Halpin-Tsai model. Utilizing theory of Mindlin, the final equations are solved and derived by method of Bolotin and differential quadrature method (DQM). In the optimization process utilizing improved meta-heuristic algorithm basis Grey Wolf optimization (GWO), the instability and frequency of the structure are utilized to define the subjective and objective functions. The GWO is improved using an adjusting randomly process with normal distribution. The main contribution of this study is maximizing the inequality and frequency constraint to control its instability. In the optimization procedure, the cone semi vertex angle, moisture and layers number change are optimized and the temperature influences, carbon fiber volume percentage, magnetic field and CNT radius are considered. The outcome shows that the proposed improved GWO may provide better abilities to search the global conditions compared to GWO because of rising flexibility to study optimum conditions of this complex problem. It is observed that the optimum frequency of system without retrofitting by CNTs is lower than the case of ωCNT≠0.
  • Mechanisms of lobed jet mixing: About circularly alternating-lobe mixers
    • Abstract: Publication date: Available online 3 January 2020Source: Aerospace Science and TechnologyAuthor(s): Zhi-qiang Sheng, Jing-yuan Liu, Yu Yao, Yi-hua XuAbstractTwo configurations of circularly arranged alternating-lobe nozzles were adopted to form lobed mixers with/without a mixing duct. The jet mixing of each mixer was numerically simulated with the unchanged initial conditions of the primary and secondary streams, except the altered initial velocity of the secondary stream. The jet-mixing mechanisms of the circularly alternating-lobe mixers were synthetically analysed by combining the evolution of the flow field structures and the process of heat and mass transfer in the mixing field. It is found that the transverse flow is usually caused by the lobed geometry, and the entrainment of the primary stream also plays a role in certain circumstances. There are two mechanisms for the deflection of the transverse flow at the lobe peaks and troughs. One of these mechanisms is to be suppressed to deflect the flow while the other is deflected by the reaction force and induction effect. The primary and secondary streams deflect to bring the transverse interval between them, and subsequently, the streamwise vortex core appears at the transverse interval. The deflected flow consistently “digging” in the radial and circumferential radiation increases the dimension of the streamwise vortices. The transverse flow velocity decreases and the direction becomes unstable leading to the breakdown of the streamwise vortices. The transverse flow brings the heat and mass transfer. Under the two mechanisms, the frontiers of the primary and secondary streams deflect. Initially, the primary and secondary streams flow around the streamwise vortex core. Subsequently, the mixed stream flows around the vortex core, and the mixing stream area gradually expands outward. The heat and mass transfer decrease in scale when the streamwise vortices break down. Because of the velocity gradient at the interface, shear instability occurs to generate the normal vortex ring. The heat and mass transfer pushes the interface, which leads to the stretch of the normal vortex ring. The mixing speed varies due to the heat and mass transfer. The velocity gradient decreases fast in the rapid mixing segment, where the normal vortex ring breaks first.
  • Theoretical study on energy performance of a stratospheric solar aircraft
           with optimum Λ-shaped rotatable wing
    • Abstract: Publication date: Available online 2 January 2020Source: Aerospace Science and TechnologyAuthor(s): Mingjian Wu, Zhiwei Shi, Haisong Ang, Tianhang XiaoAbstractThe sun-tracking rotatable wing can be used to improve energy performance of solar aircraft at low sun elevation angles. This work mainly focuses on studying the energy performance of a symmetric Λ-shaped rotatable wing solar aircraft through developing net energy optimization model, which considers the coupling of additional solar energy conversion and extra energy consumption. The derived energy conversion models reveal that, only when the morphing angle is larger than the sun elevation angle, the Λ-shaped solar aircraft is capable of achieving more energy conversion than that of planar wing. The optimum flight attitude control is adjusting its yawing angle to make wingspan axis coincide with the horizontal projection of sunlight. The numerical results of energy performance demonstrate that, with solar cell efficiency of 0.3, the optimum Λ-shaped solar aircraft can achieve 27 kWh more net energy nearby 55°N in winter, corresponding to more than 8 hours increase of flight endurance. Meanwhile, the yearly maximum perpetual flight date at 65°N increases by 34 days. In addition, the Λ-shaped solar aircraft is more effective to improve energy performance with higher conversion efficiency of solar cell. For the solar cell efficiency of 0.45, the net energy is more than 45 kWh and the increment of maximum flight latitude is 10.7° on winter solstice. The results indicate that the optimum Λ-shaped rotatable wing is an effective method to improve energy performance of solar aircrafts.
  • Bionic design for the aerodynamic shape of a stratospheric airship
    • Abstract: Publication date: Available online 2 January 2020Source: Aerospace Science and TechnologyAuthor(s): Yang Yueneng, Xu Xin, Zhang Bin, Zheng Wei, Wang YidiAbstractThe aerodynamic shape of a stratospheric airship is closely to its aerodynamic drag and propulsion energy consumption. The physalia physalis, which has unique characteristics of cystic shape and flow resistance, provides important bionic inspirations for aerodynamic shape design of the stratospheric airships. This paper proposed a bionic methodology for aerodynamic shape design inspired by the physalia physalis. First, the morphological characteristics of the physalia physalis is investigated via image processing, and its profile is obtained using edge detection. Second, the aerodynamic shape of the stratospheric airship is designed via “morphological imitation”, and the profile of the airship hull is described by function curves. Finally, the computational mesh model of the stratospheric airship is developed, and the drag coefficients, lift coefficients and lift-drag ratio are obtained by computation studies, respectively. Computation results demonstrate that the designed stratospheric airship has better aerodynamic performances than the conventional stratospheric airship.
  • Method for eliminating aerodynamic lift vibration of rigid rotor
           helicopters based on the novel sine-trim model
    • Abstract: Publication date: Available online 31 December 2019Source: Aerospace Science and TechnologyAuthor(s): Wei-Liang LyuAbstractA novel methodology called sine-trim is proposed from the phasor superposition principle for AC voltages in electrical engineering to eliminate the aerodynamic lift vibration of rigid rotor helicopters. An advancing blade concept rotor aerodynamic analysis model is presented on the basis of free-wake method associated with the inner- and outer-trim modules. The inner-trim module transforms the blade lift distribution curve of a rigid lift-offset rotor into a similar sinusoidal one by adjusting the blade pitch in each azimuth. An acceleration algorithm that could greatly improve the inner-trim calculation efficiency is provided. The outer-trim module selects the amplitude of the lower rotor as the preset trim value, and it is reliable and effective in trimming the overall forces and moments of the helicopter by modifying the lift sinusoid curve parameters of the upper and lower rotors. In the sine-trim, the upper and lower rotors have similar sinusoidal lift distribution parameters, and their phases and amplitudes exhibit a direct inverse relationship. Calculation results show that after being sine-trimmed, the total aerodynamic lift vibration of the rigid rotors could be eliminated regardless of the change in flight and operating conditions, such as flight speed, number of blades, and rotor rotational speed. The optimal blade pitch control distribution, which is a comprehensive superposition of many higher harmonic inputs, is obtained in all azimuths. For the same lift-offset, after sine-trimming, the effective lift-to-drag ratio may have a small degree of loss compared with the conventional advancing blade concept rotor.
  • Trajectory optimization for parafoil delivery system considering
           complicated dynamic constraints in high-order model
    • Abstract: Publication date: Available online 12 December 2019Source: Aerospace Science and TechnologyAuthor(s): Hao Sun, Shuzhen Luo, Qinglin Sun, Zengqiang Chen, Wannan Wu, Jin TaoAbstractTrajectory optimization is essential for all unmanned aerial vehicles, especially the parafoil delivery system, as the optimized trajectory must be practical and realizable. However, with the traditional 3-degree of freedom (DOF) model used in trajectory optimization, the realizability of the optimized trajectory is largely reduced without consideration of its dynamic constraints. Moreover, the landing and tracking precision of the system will also be reduced. To solve this problem, in this paper, a novel trajectory optimization method is explored. Different from existing research, trajectory optimization is achieved with a complete 6-DOF dynamic model of the parafoil delivery system. When this is combined with the Gauss pseudospectral method, complicated dynamic constraints can be taken into account in the trajectory optimization. In addition to achieving a global optimal control objective under dynamic constraints, the proposed method also can include multiple external environment constraints, such as terrain, wind disturbance and flared landing. Finally, our results show that the proposed method can achieve all the control objectives successfully. The optimized trajectory is smooth and realizable. Compared with other trajectory optimization methods that use a 3-DOF model, the advantage of the proposed method is clear. Considering the complicated dynamic constrains of the system, it has much better realizability and is more consistent with reality.
  • Analysis of the maximum flight Mach number of hydrocarbon-fueled scramjet
           engines under the flight cruising constraint and the combustor cooling
    • Abstract: Publication date: Available online 30 November 2019Source: Aerospace Science and TechnologyAuthor(s): Wang Youyin, Cheng Kunlin, Tang Jingfeng, Liu Xiaoyong, Bao WenAbstractScramjet is the optimal propulsion system for hypersonic vehicles, and it is important to study its maximum flight Mach number. Upper limit and its influence factors of the flight Mach number of dual-mode scramjet used in hypersonic vehicles are investigated in this paper. Scramjet performance analysis model with the constant-area heating process, cooling requirement analysis model of the combustor and cruise model of the scramjet-powered hypersonic vehicle are established. Specific impulse of dual-mode scramjet engine with hydrocarbon fuel is similar to that of liquid oxygen kerosene rocket engine when the freestream Mach number is above 10. The ratio of hypersonic vehicle reference area to scramjet inlet capture area is introduced to establish the size relationship between the scramjet engine and the hypersonic vehicle. Considering the balance of thrust and drag, smaller ratio of hypersonic vehicle reference area to scramjet inlet capture area leads to higher flight Mach number that hypersonic vehicle can cruise. When the angle of attack is 10 degrees, the ratio of hypersonic vehicle reference area to scramjet capture area should be less than 5.5 to ensure that the vehicle can work at Mach 8. The maximum length to diameter ratio of the combustor is used to indicate active cooling requirements, and it is less than 5 under the cooling requirement when flight Mach number is above 9. Reducing the heat flux at the combustor wall and increasing the fuel heat sink is helpful to increase the allowable combustion chamber length, so as to increase the maximum flight Mach number of the scramjet engine on the premise of meeting thermal protection requirement.
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