Journal Cover Aerospace Science and Technology
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   Hybrid Journal Hybrid journal (It can contain Open Access articles)
   ISSN (Print) 1270-9638
   Published by Elsevier Homepage  [3175 journals]
  • Vibro-acoustic response and sound transmission loss characteristics of
           truss core sandwich panel filled with foam
    • Authors: M.P. Arunkumar; Jeyaraj Pitchaimani; K.V. Gangadharan; M.C. Leninbabu
      Pages: 1 - 11
      Abstract: Publication date: July 2018
      Source:Aerospace Science and Technology, Volume 78
      Author(s): M.P. Arunkumar, Jeyaraj Pitchaimani, K.V. Gangadharan, M.C. Leninbabu
      This paper presents the studies carried out for improving the acoustic behavior of truss core sandwich panel, which is mostly used in aerospace structural applications. The empty space of the truss core is filled with polyurethane foam (PUF) to achieve better vibro-acoustic and sound transmission loss characteristics. Initially equivalent elastic properties of the foam filled truss core sandwich panel are calculated. Then, the vibration response of the panel under a harmonic excitation is obtained based on the equivalent 2D finite element model. Finally, the vibration response is given as an input to the Rayleigh integral code built in-house to obtain the acoustic and sound transmission loss characteristics. The results revealed that PUF filling of the empty space of the truss core, significantly reduces resonant amplitudes of both vibration and acoustic responses. It is also observed that foam filling reduces the overall sound power level significantly. Similarly, sound transmission loss studies revealed that, sudden dips at resonance frequencies are significantly reduced. Also an experiment is conducted on forced vibration response of honeycomb core sandwich panel to show that equivalent 2D model can be used for predicting sound power level and transmission loss behavior.

      PubDate: 2018-04-15T07:30:57Z
      DOI: 10.1016/j.ast.2018.03.029
      Issue No: Vol. 78 (2018)
       
  • Optimization of rough transonic axial compressor
    • Authors: Zhihui Li; Yanming Liu
      Pages: 12 - 25
      Abstract: Publication date: July 2018
      Source:Aerospace Science and Technology, Volume 78
      Author(s): Zhihui Li, Yanming Liu
      The influence of wall roughness on the performance of the axial transonic compressor stage was investigated with different values of roughness added to the blade, hub and shroud sections. The dimensionless sand-grain roughness model was used to capture the roughness effect and the results indicated that the increment of both end wall and blade surface roughness caused the deterioration of compressor stage performance. The sensitivity analysis method was used to distinguish which section mostly contributes to the whole performance degradation. Approximately a 95.31% degradation of the compressor peak efficiency came from the induced blade roughness, 3.58% from the hub surface roughness and only 1.08% from the casing surface. The present study also investigated how the optimized design of compressor blades was affected by considering a surface roughness effect representative of in-service use. Two optimization strategies were proposed to improve the compressor efficiency and total pressure ratio by changing the distributions of the blade angles along the chord. The first strategy considered the compressor surface to be hydraulically smooth and the consequent Pareto Front designs were degraded by increasing the level of surface roughness with the second approach considering the surface roughness from the outset of optimization. The optimization result showed that the degraded compressors from the first strategy was still among the best performing Pareto Front designs in terms of adiabatic efficiency and pressure ratio when compared to the second approach. This means that the roughness effect can be regarded as an additional factor and be considered in the end of the design process for single-stage compressors.

      PubDate: 2018-04-15T07:30:57Z
      DOI: 10.1016/j.ast.2018.03.031
      Issue No: Vol. 78 (2018)
       
  • Effect of dual-catalytic bed using two different catalyst sizes for
           hydrogen peroxide thruster
    • Authors: Seonuk Heo; Sungkwon Jo; Yongtae Yun; Sejin Kwon
      Pages: 26 - 32
      Abstract: Publication date: July 2018
      Source:Aerospace Science and Technology, Volume 78
      Author(s): Seonuk Heo, Sungkwon Jo, Yongtae Yun, Sejin Kwon
      For a catalytic bed in hydrogen peroxide based propulsion systems, a high pressure drop can cause significant problems. Hence, a dual-catalytic bed was suggested to reduce the pressure drop across the catalytic bed. Catalysts of two different sizes (1/8 inch, and 1.18–2.00 mm) were employed, which were fabricated using an impregnation method with MnO2 and PbO as the active materials. The upstream and downstream sides of the dual-catalytic bed were loaded with the catalyst with dimensions of 1.18–2.00 mm and 1/8 inch, respectively. The effectiveness of the dual-catalytic bed was verified by conducting hot-fire tests with hydrogen peroxide monopropellant mode. The trends in the pressure drop across the catalytic bed and the characteristic velocity efficiency were investigated with respect to the mass flux and mass ratio of the loaded catalysts. As the mass ratio of the smaller catalyst was reduced to 18.3%, the pressure drop constantly decreased with an identical mass flux, though most of the fed hydrogen peroxide was still fully decomposed.

      PubDate: 2018-04-15T07:30:57Z
      DOI: 10.1016/j.ast.2018.03.032
      Issue No: Vol. 78 (2018)
       
  • Global stabilization of the linearized three-axis axisymmetric spacecraft
           attitude control system by bounded linear feedback
    • Authors: Weiwei Luo; Bin Zhou; Guang-Ren Duan
      Pages: 33 - 42
      Abstract: Publication date: July 2018
      Source:Aerospace Science and Technology, Volume 78
      Author(s): Weiwei Luo, Bin Zhou, Guang-Ren Duan
      In this paper, the three-axis attitude stabilization of the axisymmetric spacecraft with bounded inputs is studied. By constructing some novel state transformations, saturated linear state feedback controllers are constructed for the considered attitude control system. By constructing suitable quadratic plus integral Lyapunov functions, globally asymptotic stability of the closed-loop systems is proved if the feedback gain parameters satisfy some explicit conditions. By solving some min–max optimization problems, a global optimal feedback gain for the underactuated attitude stabilization system is proposed such that the convergence rate of the linearized closed-loop system is maximized. Numerical simulations show the effectiveness of the proposed approaches.

      PubDate: 2018-04-15T07:30:57Z
      DOI: 10.1016/j.ast.2018.02.021
      Issue No: Vol. 78 (2018)
       
  • A dual-rate hybrid filtering method to eliminate high-order position
           errors of GPS in POS
    • Authors: Zhuangsheng Zhu; Chi Li; Wen Ye
      Pages: 43 - 53
      Abstract: Publication date: July 2018
      Source:Aerospace Science and Technology, Volume 78
      Author(s): Zhuangsheng Zhu, Chi Li, Wen Ye
      The Position and Orientation System (POS) serves as a key component for the airborne remote sensing system, which integrates Strapdown Inertial Navigation System (SINS) and Global Position System (GPS) to provide the reliable and continuous motion compensation using Kalman Filter (KF). However, the high-order position errors resulting from C/A (Coarse/Acquisition) Code GPS cannot be effectively compensated or estimated by the traditional KF, which severely weakens the imaging quality. In this paper, we propose a Dual-rate Hybrid Filter (DHF) to deal with the high-order position errors based on Least Squares Support Vector Machine (LSSVM) and Kalman Filter. DHF builds a low update rate filter by integrating high-precision SINS and online LSSVM to isolate the high-order position errors. Meanwhile, the high update rate filter of DHF maintains the advantages of traditional SINS/GPS integrated navigation system to restrain the accumulation errors of system. The experimental results show that the proposed method significantly reduces the high-order position errors by 84.6% at each sampling period comparing with the conventional single KF based SINS/GPS integrated navigation system.

      PubDate: 2018-04-15T07:30:57Z
      DOI: 10.1016/j.ast.2018.03.036
      Issue No: Vol. 78 (2018)
       
  • A new sliding mode control design for integrated missile guidance and
           control system
    • Authors: Jianguo Guo; Yu Xiong; Jun Zhou
      Pages: 54 - 61
      Abstract: Publication date: July 2018
      Source:Aerospace Science and Technology, Volume 78
      Author(s): Jianguo Guo, Yu Xiong, Jun Zhou
      A new sliding mode control algorithm for integrated guidance and control (IGC) system is proposed in this paper. Firstly, the IGC model is established and the nonlinearities, target maneuvers, perturbations caused by variations of aerodynamic parameters, etc. are viewed as disturbance, so that the IGC system becomes a mismatched uncertain linear system. Secondly, a second-order disturbance observer is used to estimate the disturbances and their derivatives. Thirdly, an integral sliding mode surface is designed to obtain the rudder deflection command directly instead of the back-stepping control (BC) algorithm used in conventional IGC system, which achieves the real sense of IGC, and the stability of the system is proven strictly by Lyapunov stability theory. Finally, the superiority of the proposed IGC method is verified by comparing the simulation results of different methods under different cases.

      PubDate: 2018-04-15T07:30:57Z
      DOI: 10.1016/j.ast.2018.03.042
      Issue No: Vol. 78 (2018)
       
  • Pendulum maneuvering strategy for hypersonic glide vehicles
    • Authors: Jianwen Zhu; Ruizhi He; Guojian Tang; Weimin Bao
      Pages: 62 - 70
      Abstract: Publication date: July 2018
      Source:Aerospace Science and Technology, Volume 78
      Author(s): Jianwen Zhu, Ruizhi He, Guojian Tang, Weimin Bao
      In order to improve the penetration performance of hypersonic glide vehicles, a lateral pendulum maneuvering strategy is proposed. A single radar trajectory tracking model is established and EKF is used to estimate the entire trajectory parameters. Based on the analysis of the composition of the defense system and the intercepting mechanism, the pendulum maneuvering trajectory is designed, and the influencing factors of the gliding penetration performance are analyzed. Then, an integrated index of penetration performance consists of the hit point prediction error, intercepting velocity, overload and the energy consumption caused by maneuver is constructed. Furthermore, a maneuvering strategy is proposed that the first maneuver is performed to enlarge prediction error of hitting point when the vehicle entrances the radar coverage, and the second one is carried out in the intercept zone to increase the maneuvering overload. The two maneuvers are the combat of the glider to the early warning system and the intercept system respectively, which can effectively enhance the penetration performance with less energy consumption.

      PubDate: 2018-04-15T07:30:57Z
      DOI: 10.1016/j.ast.2018.03.038
      Issue No: Vol. 78 (2018)
       
  • Assessment of low-fidelity fluid–structure interaction model for
           flexible propeller blades
    • Authors: Jurij Sodja; Roeland De Breuker; Dejan Nozak; Radovan Drazumeric; Pier Marzocca
      Pages: 71 - 88
      Abstract: Publication date: July 2018
      Source:Aerospace Science and Technology, Volume 78
      Author(s): Jurij Sodja, Roeland De Breuker, Dejan Nozak, Radovan Drazumeric, Pier Marzocca
      Low-fidelity fluid–structure interaction model of flexible propeller blades is assessed by means of comparison with high-fidelity aeroelastic results. The low-fidelity model is based on a coupled extended blade-element momentum model and non-linear beam theory which were both implemented in Matlab. High-fidelity fluid–structure interaction analysis is based on coupling commercial computational fluid dynamics and computational structural dynamics codes. For this purpose, Ansys CFX® and Ansys Mechanical® were used. Three different flexible propeller blade geometries are considered in this study: straight, backward swept, and forward swept. The specific backward and forward swept blades are chosen due to their aeroelastic response and its influence on the propulsive performance of the blade while a straight blade was selected in order to serve as a reference. First, the high-fidelity method is validated against experimental data available for the selected blade geometries. Then the high- and low-fidelity methods are compared in terms of integral thrust and breaking power as well as their respective distributions along the blades are compared for different advancing ratios. In a structural sense, the comparison is performed by analyzing the blade bending and torsional deformation. Based on the obtained results, given the simplicity of the low-fidelity method, it can be concluded that the agreement between the two methods is reasonably good. Moreover, an important result of the comparison study is an observation that the advance ratio is no longer a valid measure of similarity in the case of flexible propeller blades and the behavior of such blades can change significantly with changing operating conditions while keeping the advance ratio constant. This observation is supported by both high- and low-fidelity methods.

      PubDate: 2018-04-15T07:30:57Z
      DOI: 10.1016/j.ast.2018.03.044
      Issue No: Vol. 78 (2018)
       
  • Influence of Mach number and angle of attack on the two-dimensional
           transonic buffet phenomenon
    • Authors: Nicholas F. Giannelis; Oleg Levinski; Gareth A. Vio
      Pages: 89 - 101
      Abstract: Publication date: July 2018
      Source:Aerospace Science and Technology, Volume 78
      Author(s): Nicholas F. Giannelis, Oleg Levinski, Gareth A. Vio
      Within a narrow band of flight conditions in the transonic regime, self-sustained shock oscillations that involve the interaction between shock-waves and intermittently separated shear layers may develop. This phenomenon, known as transonic shock buffet, limits the flight envelope and is detrimental to both aircraft handling quality and structural integrity. In this investigation, numerical simulation of transonic shock buffet over the OAT15A aerofoil is performed to explore the buffet envelope. Unsteady Reynolds-Averaged Navier–Stokes simulations are validated against available experimental data to ascertain the most effective combination of simulation parameters to reproduce autonomous shock oscillations. From the baseline test case, the influence of Mach number and angle of attack on the nature of the buffet response is investigated. Radial Basis Function surrogate models are developed to represent the variation of buffet amplitude and frequency with flight condition. While the frequency is found to increase monotonically with both parameters, variation in buffet amplitude through the region of shock unsteadiness is more complex, particularly at high angles of attack. This is related to a bifurcation in the behaviour of the shock. As incidence increases from onset, the shock dynamics transition from periodic oscillations over the suction surface to quasi-periodic motions, whereby the shock is propelled forward into the oncoming flow during its upstream excursion.

      PubDate: 2018-04-15T07:30:57Z
      DOI: 10.1016/j.ast.2018.03.045
      Issue No: Vol. 78 (2018)
       
  • Steepest descent quaternion attitude estimator
    • Authors: Pawel Zagorski; Tomasz Dziwinski; Andrzej Tutaj
      Pages: 1 - 10
      Abstract: Publication date: June 2018
      Source:Aerospace Science and Technology, Volume 77
      Author(s): Pawel Zagorski, Tomasz Dziwinski, Andrzej Tutaj
      A new computationally inexpensive attitude determination algorithm based on the minimization of Wahba's loss function is presented in the paper. The estimation problem is converted into quaternion representation and solved with iterative prediction–correction scheme. Unlike Kalman filter approach, an iterative gradient optimization is used to estimate the attitude quaternion and gyroscope bias. Algorithm derivation is shown and its performance is tested. The presented case study assumes configuration with three types of sensors: Sun sensors with full angular coverage, a magnetometer and a MEMS rate gyroscope. Sensor model parameters are selected to mimic a pico or nano class satellite. Orbital environment is simulated with the Bouvier–Lyddane orbit model, the IGRF magnetic field model and geometric properties of the Earth–Sun system. Periodical loss of Sun sensor data due to eclipses is taken into account. Based on the presented case study a proposition of tuning procedure and a brief comment on algorithm stability are given. The tuning approach trades off estimate convergence versus noise rejection property. In a Monte Carlo test the proposed algorithm compares well against an EKF with an attitude error within 0.1 deg in sunlight and 0.4 deg in the eclipse. Finally, a simulation showing a possibility of operating the SDQAE algorithm while sampling each of the sensors at different rate is presented.

      PubDate: 2018-04-15T07:30:57Z
      DOI: 10.1016/j.ast.2018.01.030
      Issue No: Vol. 77 (2018)
       
  • Uncertainty propagation in aerodynamic forces and heating analysis for
           hypersonic vehicles with uncertain-but-bounded geometric parameters
    • Authors: Yuning Zheng; Zhiping Qiu
      Pages: 11 - 24
      Abstract: Publication date: June 2018
      Source:Aerospace Science and Technology, Volume 77
      Author(s): Yuning Zheng, Zhiping Qiu
      In this study, uncertainties in aerodynamic forces and heating properties of hypersonic vehicles are calculated and analyzed with consideration of uncertain-but-bounded geometric parameters. The aerodynamic shape of a hypersonic vehicle is created with a few geometric parameters containing physical meanings after applying the class and shape transformation (CST) method. Considering uncertainties in geometric parameters caused by manufacturing errors, interval variables are introduced to quantify geometric parameters and a novel interval-based CST method (ICST) is proposed to represent the uncertain aerodynamic shape. By means of hypersonic engineering methods, aerodynamic forces and heating properties of hypersonic vehicles can be predicted. The interval analysis method and novel Bernstein-polynomial-based method for calculating the lower and upper bounds of aerodynamic forces and heating properties are developed. The results of analyzing two numerical examples demonstrate the effectiveness and feasibility of the proposed method and further confirm the necessity of accounting for the uncertainties in geometric parameters when investigating aerodynamic forces and heating properties of hypersonic vehicles.

      PubDate: 2018-04-15T07:30:57Z
      DOI: 10.1016/j.ast.2018.02.028
      Issue No: Vol. 77 (2018)
       
  • Numerical study on solid-fuel scramjet combustor with fuel-rich hot gas
    • Authors: Xiang Zhao; Zhi-xun Xia; Bing Liu; Zhong Lv; Li-kun Ma
      Pages: 25 - 33
      Abstract: Publication date: June 2018
      Source:Aerospace Science and Technology, Volume 77
      Author(s): Xiang Zhao, Zhi-xun Xia, Bing Liu, Zhong Lv, Li-kun Ma
      Solid-fuel scramjet combustor with cavity faces challenges of ignition and flame holding. In the current study, a novel concept, solid-fuel scramjet combustor with fuel-rich hot gas, is proposed. The Reynolds-average Navier–Stokes (RANS) equations coupled with the SST k − ω turbulence model and the second-order spatially accurate upwind scheme are employed to calculate its flow field. The feasibility of the solid-fuel scramjet combustor with fuel-rich hot gas is studied based on the validation of numerical method. The comparison of the performance is made between the combustor with fuel-rich gas and the combustor with cavity. Various parameters, namely excess air coefficient of gas generator and mass flow rate of fuel-rich gas, are studied, and their effects on the combustion efficiency, total pressure recovery and fuel regression rate are analyzed. This is the basic study for experiments of a solid-fuel scramjet combustor with fuel-rich hot gas.

      PubDate: 2018-04-15T07:30:57Z
      DOI: 10.1016/j.ast.2017.12.024
      Issue No: Vol. 77 (2018)
       
  • An application of Deep Neural Networks to the in-flight parameter
           identification for detection and characterization of aircraft icing
    • Authors: Yiqun Dong
      Pages: 34 - 49
      Abstract: Publication date: June 2018
      Source:Aerospace Science and Technology, Volume 77
      Author(s): Yiqun Dong
      This paper applies the Deep Neural Networks to the in-flight parameter identification for detection and characterization of the aircraft icing. General dynamics of the aircraft are firstly presented, ice effects on the dynamics are characterized. Deep Neural Networks (DNNs) including Convolutional Neural Network (CNN) and Recurrent Neural Network (RNN) are briefly introduced. We propose a “state-image” approach for the pre-processing of the input flight state, then we design a DNN structure which models both local connectivity (using CNN) and temporal characteristics (using RNN) of the flight state. The identified parameters are exported from the DNN output layer directly. To fully evaluate the performance of the DNN-based approach, we conduct simulation tests for different cases which correspond to clean and aircraft icing at different locations (wing, tail, wing and tail) with different severities (moderate, severe). A comparison of the DNN-based approach with a baseline H ∞ -based identification algorithm (state-of-the-art for aircraft icing) is also delivered. Based on the test and comparison results, the DNN-based approach yields more accurate identification performance for more parameters, which shows promising applicability to the in-flight parameter identification problem.

      PubDate: 2018-04-15T07:30:57Z
      DOI: 10.1016/j.ast.2018.02.026
      Issue No: Vol. 77 (2018)
       
  • Extending slide-slip mesh update method to finite volume method
    • Authors: Kun Qu; Feng Xie; Jinsheng Cai
      Pages: 50 - 57
      Abstract: Publication date: June 2018
      Source:Aerospace Science and Technology, Volume 77
      Author(s): Kun Qu, Feng Xie, Jinsheng Cai
      For simulations of flows around rotating bodies, usually sliding mesh method or overset grid method is used. But both of them have to perform inter-mesh interpolation which introduces much numerical error and usually violates conservation. Shear-slip mesh update method (SSMUM) is another method for such flows. In each time step of SSMUM, a mesh slipping step follows a mesh deforming step to undo the deformation and results in a new mesh of good quality. Each vertex on the slipping interface can only move from one node to the next node in the circumference, which makes the interface always conformal and no need for inter-mesh interpolation. However, SSMUM didn't get enough attention in the community of finite volume method. In this paper, SSMUM is extended to cell center finite volume method. To guarantee conservation and obtain high order accuracy on a slipping interface, a remapping procedure is needed to transfer flow field from an old mesh to a new mesh. This was achieved by solving a linear convective PDE with one or two explicit steps, thus only resulting in little extra computing cost. Oscillating NACA0012 airfoil was simulated with the improved SSMUM. The results showed excellent agreement with the data by rigid rotating mesh. And the flow field was always smooth. It suggests that this improved SSMUM has advantages in getting conservative, smooth and high accuracy solutions for rotating problems.

      PubDate: 2018-04-15T07:30:57Z
      DOI: 10.1016/j.ast.2018.02.025
      Issue No: Vol. 77 (2018)
       
  • Studies on effusion cooling: Impact of geometric parameters on cooling
           effectiveness and coolant consumption
    • Authors: Vishal Venkatesh; Sriraam J.; Bala Vignesh D.; Subash K.; Ratna Kishore Velamati; Srikrishnan A.R.; Balajee Ramakrishnananda; Suresh Batchu
      Pages: 58 - 66
      Abstract: Publication date: June 2018
      Source:Aerospace Science and Technology, Volume 77
      Author(s): Vishal Venkatesh, Sriraam J., Bala Vignesh D., Subash K., Ratna Kishore Velamati, Srikrishnan A.R., Balajee Ramakrishnananda, Suresh Batchu
      This study is focused on the impact of certain important geometric parameters on cooling effectiveness and coolant consumption for effusion cooling of aircraft combustor liner. The three dimensional turbulent flow field in a domain representing the combustor with several rows of effusion coolant injection is considered for the analysis. The geometric parameters considered are: angle of injection of the coolant, axial and transverse pitch of the injection holes, hole spacing and hole diameter. Also, based on the analysis of the temperature field within the chamber, a novel concept of ‘variable hole diameter’ has been introduced to reduce coolant consumption. A symmetric 3D computational model including the combustion chamber, coolant chamber and the effusion plate was used for the study. Conjugate heat transfer was modeled between the effusion-cooled wall and the two chambers. A detailed mass flow rate analysis has been performed for the various cases in order to study the impact of parameters on coolant consumption. The proposed approach of using an effusion plate with variable hole diameters is found to be effective in reducing the net coolant consumption significantly while maintaining a given level of cooling effectiveness.

      PubDate: 2018-04-15T07:30:57Z
      DOI: 10.1016/j.ast.2017.12.044
      Issue No: Vol. 77 (2018)
       
  • Performance assessment of electrically driven pump-fed LOX/kerosene cycle
           rocket engine: Comparison with gas generator cycle
    • Authors: Hyun-Duck Kwak; Sejin Kwon; Chang-Ho Choi
      Pages: 67 - 82
      Abstract: Publication date: June 2018
      Source:Aerospace Science and Technology, Volume 77
      Author(s): Hyun-Duck Kwak, Sejin Kwon, Chang-Ho Choi
      An electrically driven pump-fed cycle for rocket engine is proposed and a viability of the proposed cycle is assessed compared to a gas generator cycle. The maximum possible thrust level is determined considering the technological maturity of the electric motor. Four types of battery cells were assessed in a screening test for the proposed cycle and the necessity of regenerative cooling for the battery pack is shown. The mass expressions of the proposed cycle and gas generator cycle are derived in terms of pump power and burning time. The basic features are demonstrated with respect to combustion chamber pressure, burning time, and thrust level. The results show that it is favorable to maintain a lower combustion chamber pressure, a longer burning time, and a higher thrust level to remedy the payload penalty incurred when the gas generator cycle is not used. In addition to focusing on the battery pack, the regenerative cooling effect on the battery pack mass is discussed. Further, the impact of optimal battery cell discharge time on the payload is explained. To estimate the payload for the proposed cycle quantitatively, hypothetical low earth orbit (LEO) and KSLV-II sun synchronous orbit (SSO) mission cases are used. In the analysis of the hypothetical LEO mission, it is found that the proposed cycle payloads are only 2.1% to 3.5% lower than those of the gas generator cycle when the combustion chamber pressure is 3.0 MPaA. For the KSLV-II SSO mission, the cargo payload is increased by 3.7% compared to that of gas generator cycle if the proposed cycle is employed for the third-stage engine.

      PubDate: 2018-04-15T07:30:57Z
      DOI: 10.1016/j.ast.2018.02.033
      Issue No: Vol. 77 (2018)
       
  • Combustion stabilizations in a liquid kerosene fueled supersonic combustor
           equipped with an integrated pilot strut
    • Authors: Junlong Zhang; Juntao Chang; Wen Shi; Wenxin Hou; Wen Bao
      Pages: 83 - 91
      Abstract: Publication date: June 2018
      Source:Aerospace Science and Technology, Volume 77
      Author(s): Junlong Zhang, Juntao Chang, Wen Shi, Wenxin Hou, Wen Bao
      The numerical and experimental investigations have been conducted to test a newly designed integrated pilot strut. The integrated pilot strut consists of two neighboring small strut with shallow cavities, with the help of which, the fuel injection and flame holding are achieved in the supersonic combustor. The flowing characteristics in the internal flow duct of the pilot strut are evaluated with the numerical simulation method, results proving that a lower speed zone generates in the internal flow duct in the supersonic combustor and the local equivalence ratio in the low speed zone is suitable for combustion. Then, a series of experiments have been conducted in the flight condition of Ma = 5 , with stagnation state T t = 1270 K, P t = 1.20 MPa. Experimental results show that a pilot flame generates in the internal flow duct of the pilot strut, based on which, the main fuel injected from the sidewall of the strut is ignited, and the global flame is established in the whole combustor. The combustion of the main fuel leads to a thermal chocking at the exit of the strut. Further, the thermal chocking is beneficial to the self-stabilization of the pilot flame. With the combustion organization strategy of the pilot strut flame holding, the global flame is stabilized in a wide range of equivalence ratio changing from 0.15 to 0.75 in the supersonic combustor, and the combustion characteristics in different equivalence ratios are analyzed in this paper. The integrated combustion organization approach by the pilot strut with internal cavities is demonstrated feasible and a high combustion performance is obtained.

      PubDate: 2018-04-15T07:30:57Z
      DOI: 10.1016/j.ast.2018.02.035
      Issue No: Vol. 77 (2018)
       
  • Disturbance observer based reliable H∞ fuzzy attitude tracking control
           for Mars entry vehicles with actuator failures
    • Authors: Huai-Ning Wu; Zi-Peng Wang; Lei Guo
      Pages: 92 - 104
      Abstract: Publication date: June 2018
      Source:Aerospace Science and Technology, Volume 77
      Author(s): Huai-Ning Wu, Zi-Peng Wang, Lei Guo
      This paper introduces a disturbance observer (DO) based reliable H ∞ fuzzy attitude tracking control design for Mars entry vehicles with actuator failures. Initially, to reduce the complexity of Takagi–Sugeno (T–S) fuzzy modeling, the two time-scale decomposition technique is used to divide the original nonlinear attitude tracking error model of Mars entry vehicles into a slow subsystem describing the attitude kinematics and a fast subsystem describing the attitude dynamics. The dynamic inversion control (DIC) method is subsequently applied to the slow subsystem to generate the angular velocity command. Then, the T–S fuzzy modeling method is employed to exactly represent the fast subsystem and a disturbance observer (DO) is constructed to estimate the modeled disturbance based on the derived tracking error fuzzy system of angular velocity. By the technique of linear matrix inequalities (LMIs), a DO based reliable H ∞ fuzzy controller of attitude tracking is developed to stabilize exponentially the angular velocity tracking error and the modeled-disturbance state estimation error with an H ∞ tracking performance both in nominal and actuator failure cases. Furthermore, it is shown that the original nonlinear tracking error system is also exponentially stable and satisfies an H ∞ tracking performance both in nominal and actuator failure cases under the proposed fuzzy control law together with the DIC law, provided that the timescale separation between the fast and slow subsystems is valid. Finally, simulation results illustrate the effectiveness of the proposed design method.

      PubDate: 2018-04-15T07:30:57Z
      DOI: 10.1016/j.ast.2018.02.032
      Issue No: Vol. 77 (2018)
       
  • Improved polar inertial navigation algorithm based on pseudo INS
           mechanization
    • Authors: Meng Liu; Guangchun Li; Yanbin Gao; Shutong Li; Qingwen Meng; Shitong Du
      Pages: 105 - 116
      Abstract: Publication date: June 2018
      Source:Aerospace Science and Technology, Volume 77
      Author(s): Meng Liu, Guangchun Li, Yanbin Gao, Shutong Li, Qingwen Meng, Shitong Du
      From the perspective of global inertial navigation system (INS), an improved inertial navigation algorithm is proposed to solve the navigation problem in polar regions. The implementation of navigation system is achieved with two executive units, navigation calculation (NC) unit and parameter transformation (PT) unit. The NC unit is executed autonomously with pseudo INS mechanization under the spherical Earth model. The pseudo INS mechanization is established with the pseudo-Earth frame, which is a generalized Earth frame and is reconstructed based on the switching position selected properly with the motion range of vehicle. Then, the smooth switching of navigation frame and the intrinsic unity of navigation algorithm can be achieved, thereby further unifying the polar navigation algorithm for both strapdown INS and platform INS in a brief form in global regions. On the other hand, the PT units are employed to correct and transform navigation parameters to a corresponding coordinate system with ellipsoidal Earth model and can be conducted simultaneously in parallel, thereby ensuring the navigation accuracy and communicating with other local navigation systems conveniently. Theoretical analysis and numerical results indicate the validity of the proposed algorithm.

      PubDate: 2018-04-15T07:30:57Z
      DOI: 10.1016/j.ast.2018.02.029
      Issue No: Vol. 77 (2018)
       
  • An integrated load sensing valve-controlled actuator based on
           power-by-wire for aircraft structural test
    • Authors: Yaoxing Shang; Xiaochao Liu; Zongxia Jiao; Shuai Wu
      Pages: 117 - 128
      Abstract: Publication date: June 2018
      Source:Aerospace Science and Technology, Volume 77
      Author(s): Yaoxing Shang, Xiaochao Liu, Zongxia Jiao, Shuai Wu
      The traditional loading system, for the full-sized aircraft structure test, requires the centralized hydraulic power supply and the distributed valve-controlled cylinder, which leads to complex and large pipeline systems. The layout reconfiguration process of the test platform is quite laborious. Moreover, the loading system efficiency is quiet low because of both the huge overflow loss and the huge throttling loss of the traditional valve-controlled system. In order to conduct the structure test more efficiently, this paper proposes a novel integrated Load Sensing Valve-Controlled Actuator (LSVCA) with high efficiency and low energy consumption, which can reduce the overflow loss by the intermittent operation of the motor and reduce the throttling loss by the variation of the supply pressure. It also simplifies the test platform reconfiguration due to its high-level integration and Power-By-Wire (PBW) feature. In this paper, the hydraulic working principle and the energy-saving analysis of the LSVCA is proposed. In order to verify the feasibility of the new principle and the effectiveness of the high efficiency, the mathematical model of the LSVCA is established. Furthermore, an experimental prototype of the LSVCA is tested. Test results indicate that the LSVCA is adequate to satisfy the criteria for full-sized aircraft structure test. The throttling efficiency of the LSVCA loading system is 1.75 times of the efficiency of the traditional loading system under the low-level load pressure condition.

      PubDate: 2018-04-15T07:30:57Z
      DOI: 10.1016/j.ast.2018.02.030
      Issue No: Vol. 77 (2018)
       
  • Receding horizon guidance of a small unmanned aerial vehicle for planar
           reference path following
    • Authors: Yoshiro Hamada; Taro Tsukamoto; Shinji Ishimoto
      Pages: 129 - 137
      Abstract: Publication date: June 2018
      Source:Aerospace Science and Technology, Volume 77
      Author(s): Yoshiro Hamada, Taro Tsukamoto, Shinji Ishimoto
      This paper describes a novel lateral guidance law for an unmanned aerial vehicle using nonlinear receding horizon optimization and shows its flight test results. The guidance law uses an extended Kalman filter which estimates steady wind velocities in order to follow a pre-specified reference path defined in a ground-fixed coordinate system. The guidance law can be applied to arbitrary reference path as long as the path is represented as a differentiable function of x and y in a ground-fixed coordinate system. A small-scale research vehicle developed by the Japan Aerospace Exploration Agency is used for flight tests, and the results demonstrate the high guidance performance of the proposed method.

      PubDate: 2018-04-15T07:30:57Z
      DOI: 10.1016/j.ast.2018.02.039
      Issue No: Vol. 77 (2018)
       
  • Hybridized attitude determination techniques to improve ballistic
           projectile navigation, guidance and control
    • Authors: Raúl de Celis; Luis Cadarso
      Pages: 138 - 148
      Abstract: Publication date: June 2018
      Source:Aerospace Science and Technology, Volume 77
      Author(s): Raúl de Celis, Luis Cadarso
      Ballistic projectiles accuracy depends on measurement systems for position and attitude determination. Precise rotation determination is an expensive task in aircraft, as it is usually determined by strap-down sensors such as fiber optic gyros or MEMS. Particularly in ballistic projectiles, these gyro determination devices increase their price as they need to bear enormous accelerations during the initial stages but not during the ballistic flight. A new approach to improve ballistic projectile navigation, guidance and control, which integrates hybrid attitude determination methods and gravity vector estimation method, is presented in this paper. Measurements of accelerometers, GNSS-sensors and Semi-Active photo-detectors are hybridized to get such a result. The attitude determination method, avoiding the use of gyroscopes, measures pairs of vectors, i.e., gravity, velocity and line of sight vectors, in a pair of reference systems, i.e., body fixed and north–east–down reference frames. Gravity vector estimation is based on flight mechanics and aerodynamics of a ballistic projectile, which involves a deeply nonlinear behavior, but it may be extrapolated to any aircraft, and later employed in an attitude determination algorithm. Modified proportional navigation techniques and previously developed control methods are employed during flight. The presented approach is tested on a realistic nonlinear model flight simulations to prove accuracy of proposed algorithms.

      PubDate: 2018-04-15T07:30:57Z
      DOI: 10.1016/j.ast.2018.02.040
      Issue No: Vol. 77 (2018)
       
  • Nonorthogonality analysis of acoustics and vorticity modes: Should
           thermoacoustic energy norm be time-invariant'
    • Authors: Lei Li; Dan Zhao; Xiaofeng Sun
      Pages: 149 - 155
      Abstract: Publication date: June 2018
      Source:Aerospace Science and Technology, Volume 77
      Author(s): Lei Li, Dan Zhao, Xiaofeng Sun
      Detrimental thermoacoustic instability occurs frequently in many propulsion systems, such as rocket motors, and aeroengines. To predict, quantify and analyze thermoacoustic instability, it is critical to choose a proper energy norm of acoustic disturbances in a thermoacoustic combustor. It has been shown that Chu's energy norm is not associated with spurious transient energy growth, when acoustic and vorticity perturbations are assumed to share the same wave numbers. The present work performs a more generalized investigation. It is shown that vorticity and acoustic modes are non-orthogonal, even if their wave numbers are different. Further study reveals that although the vorticity and acoustic modes are not orthogonal, the cross terms of these two flow disturbances do not lead to spurious energy growth.

      PubDate: 2018-04-15T07:30:57Z
      DOI: 10.1016/j.ast.2018.02.036
      Issue No: Vol. 77 (2018)
       
  • Correction of low-Reynolds number turbulence model to hydrocarbon fuel at
           supercritical pressure
    • Authors: Zhi Tao; Zeyuan Cheng; Jianqin Zhu; Xizhuo Hu; Longyun Wang
      Pages: 156 - 167
      Abstract: Publication date: June 2018
      Source:Aerospace Science and Technology, Volume 77
      Author(s): Zhi Tao, Zeyuan Cheng, Jianqin Zhu, Xizhuo Hu, Longyun Wang
      At supercritical pressure, due to the drastic change of thermophysical property near the pseudo-critical temperature, the density fluctuation and density variation, along with buoyancy, play an important role in supercritical turbulence modelling. Based on the original LS (Launder–Sharma) low-Reynolds number turbulence model, the buoyancy modification, the density fluctuation modification, the density variation modification and the empirical coefficients modification are considered and the corresponding correction terms are derived and applied in the governing equations. Numerical simulation of heat transfer to hydrocarbon fuel flowing through the uniformly heated round pipe at supercritical pressure has been performed by the modified LS turbulence model incorporated into the in-house numerical code. Inlet temperature varied from 373 K to 473 K, with heat flux varying from 241 kW/m2 to 470 kW/m2. Inlet mass flux was 736 kg/(m2⋅s) and operating pressure was 4 MPa. The flow directions included upflow and downflow. Compared with the original LS turbulence model, the modified LS turbulence model leads to the better agreement with the experimental results, with 41.16% improvement in computation accuracy in the current study. The consideration of density fluctuation and density variation effects makes the turbulence model more suitable for thermophysical property variation at supercritical pressure.

      PubDate: 2018-04-15T07:30:57Z
      DOI: 10.1016/j.ast.2018.02.038
      Issue No: Vol. 77 (2018)
       
  • Grey wolf optimization based sense and avoid algorithm in a Bayesian
           framework for multiple UAV path planning in an uncertain environment
    • Authors: Mohammadreza Radmanesh; Manish Kumar; Mohammad Sarim
      Pages: 168 - 179
      Abstract: Publication date: June 2018
      Source:Aerospace Science and Technology, Volume 77
      Author(s): Mohammadreza Radmanesh, Manish Kumar, Mohammad Sarim
      Unmanned Air Vehicles (UAVs), which have been popular in the military context, have recently attracted attention of many researchers because of their potential civilian applications. However, before UAVs can fly in civilian airspace, they need to be able to navigate safely to their goal while maintaining separation with other manned and unmanned aircraft during the transit. Algorithms for autonomous navigation of UAVs require access to accurate information about the state of the environment in order to perform well. However, this information is often uncertain and dynamically changing. In this paper, a Grey Wolf Optimization (GWO) based algorithm is proposed to find the optimal UAV trajectory in presence of moving obstacles, referred to as Intruder Aircraft (IAs), with unknown trajectories. The solution uses an efficient Bayesian formalism with a notion of cell weighting based on Distance Based Value Function (DBVF). The assumption is that the UAV is equipped with the Automatic Dependent Surveillance-Broadcast (ADS-B) and is provided with the position of IAs either via the ADS-B or ground-based radar. However, future trajectories of the IAs are unknown to the UAV. The proposed method is verified using simulations performed on multiple scenarios. The results demonstrate the effectiveness of the proposed method in solving the trajectory planning problem of the UAVs.

      PubDate: 2018-04-15T07:30:57Z
      DOI: 10.1016/j.ast.2018.02.031
      Issue No: Vol. 77 (2018)
       
  • Reachable set computation for spacecraft relative motion with
           energy-limited low-thrust
    • Authors: Sangjin Lee; Inseok Hwang
      Pages: 180 - 188
      Abstract: Publication date: June 2018
      Source:Aerospace Science and Technology, Volume 77
      Author(s): Sangjin Lee, Inseok Hwang
      For spacecraft in formation flight, the knowledge on a reachable set for their relative motion is crucial as it can be used to 1) assess their operational capability and plan tasks accordingly, and 2) predict the operational boundary of neighboring spacecraft, from which the collision probability can be effectively monitored, improving the level of situational awareness. In this paper, a new approach for reachable set computation is proposed that computes accurate inner and outer approximations of the reachable set for a spacecraft's relative motion with energy-limited low thrust. Finding the exact boundary of the reachable set requires solving an optimal control problem with infinitely many sets of initial and terminal conditions, which is intractable. To overcome this difficulty, an analytical solution to the optimal control problem is introduced, and an ellipsoidal approximation method is applied to the solution to find two inner and outer ellipsoids that approximate the exact boundary of the reachable set. The effectiveness of the proposed approach is demonstrated with illustrative numerical examples.

      PubDate: 2018-04-15T07:30:57Z
      DOI: 10.1016/j.ast.2018.02.034
      Issue No: Vol. 77 (2018)
       
  • Safe control of trailing UAV in close formation flight against actuator
           fault and wake vortex effect
    • Authors: Ziquan Yu; Yaohong Qu; Youmin Zhang
      Pages: 189 - 205
      Abstract: Publication date: June 2018
      Source:Aerospace Science and Technology, Volume 77
      Author(s): Ziquan Yu, Yaohong Qu, Youmin Zhang
      In the close formation flight, the wake vortex induced by the leading aircraft will have adverse effects on the safe flight of the trailing unmanned aerial vehicle (UAV). Hence, this paper investigates a difficult problem of safe control for the trailing UAV against actuator faults, input saturation, and wake vortex effect. By using disturbance observers, external wake vortex, disturbances, and internal actuator faults are estimated. Then, with the help of estimated knowledge of disturbance, backstepping control laws are developed for the longitudinal dynamics and the lateral-directional dynamics, respectively. One of the key features of the proposed strategy is that, the inherent problem, i.e., “explosion of complexity” in conventional backstepping control, is solved by the presented dynamic surface control scheme. Another key feature is that external wake vortex, disturbances, and internal actuator faults, input saturation are simultaneously considered. It is shown by Lyapunov stability analysis that the closed-loop system is uniformly ultimately bounded with safety requirements guaranteed even in the presence of wake vortex and actuator faults. The effectiveness of the proposed approach is further validated by simulation results.

      PubDate: 2018-04-15T07:30:57Z
      DOI: 10.1016/j.ast.2018.01.028
      Issue No: Vol. 77 (2018)
       
  • Equations of motion for optimal maneuvering with global aerodynamic model
    • Authors: Mauricio A.V. Morales; Flávio J. Silvestre; Antônio B. Guimarães Neto
      Pages: 206 - 216
      Abstract: Publication date: June 2018
      Source:Aerospace Science and Technology, Volume 77
      Author(s): Mauricio A.V. Morales, Flávio J. Silvestre, Antônio B. Guimarães Neto
      Aircraft point-mass equations of motion have been largely adopted to calculate optimal trajectories with local aerodynamic models, i.e. valid in a restricted domain. However, some optimal maneuvers may need aerodynamic models valid for a broader range of flight conditions. For this purpose, global aerodynamic models are attractive but their nonlinear structure can preclude obtaining optimal trajectories by an indirect method together with the point-mass equations of motion. To solve this impasse without resorting to direct methods the authors propose a new set of aircraft equations of motion. When compared to the point-mass equations, the proposed set permits the inclusion of the angular velocity in the evaluation of aerodynamic forces, making them more accurate. Another advantage of the proposed model over the point-mass one is that it allows a qualitative estimate of the control surface deflections after the trajectory is obtained, which enables to discard solutions with infeasible deflections. To verify consistency, the proposed equations of motion are compared by simulation to the point-mass and to the rigid-body equations. The use of the proposed set of equations is demonstrated by three optimizations of a 360 ∘ roll problem.

      PubDate: 2018-04-15T07:30:57Z
      DOI: 10.1016/j.ast.2018.02.037
      Issue No: Vol. 77 (2018)
       
  • Performance assessment of a multi-fuel hybrid engine for future aircraft
    • Authors: Feijia Yin; Arvind Gangoli Rao; Abhishek Bhat; Min Chen
      Pages: 217 - 227
      Abstract: Publication date: June 2018
      Source:Aerospace Science and Technology, Volume 77
      Author(s): Feijia Yin, Arvind Gangoli Rao, Abhishek Bhat, Min Chen
      This paper presents the performance assessment of a novel turbofan engine using two energy sources: Liquid Natural Gas (LNG) and kerosene, called Multi-Fuel Hybrid Engine (MFHE). The MFHE is a new engine concept consisting of several novel features, such as a contra-rotating fan to sustain distortion caused by boundary layer ingestion, a sequential dual-combustion system to facilitate “Energy Mix” in aviation and a Cryogenic Bleed Air Cooling System (CBACS) to cool the turbine cooling air. The MFHE has been envisaged as a propulsion system for a long-range Multi-Fuel Blended Wing Body (MFBWB) aircraft. In this research, we study the uninstalled characteristics of the MFHE covering three aspects: 1) the effects of CBACS on the High Pressure Turbine (HPT) cooling air requirement and its consequence on the engine cycle efficiency; 2) the cycle optimization of the MFHE; 3) the performance of the MFHE at a mission level. An integrated model framework consisting of an engine performance model, a sophisticated turbine-cooling model, and a CBACS model is used. The parametric analysis shows that using CBACS can reduce the bleed air temperature significantly (up to 400 K), thereby decreasing the HPT cooling air by more than 40%. Simultaneously, the LNG temperature increases by more than 200 K. The hybrid engine alone reduces the CO2 emission by about 27% and the energy consumption by 12% compared to the current state-of-the-art turbofan engine. Furthermore, the mission analysis indicates a reduction in NOx emission by 80% and CO2 emission by 50% when compared to the baseline aircraft B-777 200ER.

      PubDate: 2018-04-15T07:30:57Z
      DOI: 10.1016/j.ast.2018.03.005
      Issue No: Vol. 77 (2018)
       
  • Effect of extended necks on transmission loss performances of Helmholtz
           resonators in presence of a grazing flow
    • Authors: He Zhao; Zhengli Lu; Yiheng Guan; Zhiqiang Liu; Guoneng Li; Jun Liu; C.Z. Ji
      Pages: 228 - 234
      Abstract: Publication date: June 2018
      Source:Aerospace Science and Technology, Volume 77
      Author(s): He Zhao, Zhengli Lu, Yiheng Guan, Zhiqiang Liu, Guoneng Li, Jun Liu, C.Z. Ji
      Suppressing acoustic pulsations is a critical task in modern premixed combustion-involved gas turbines. As a classical acoustic noise damper, Helmholtz resonator is generally applied in gas turbine combustors to reduce the transmission of acoustic perturbations. However, the neck configuration of a Helmholtz resonator may be designed in different ways. To obtain an optimum design and to compare the noise damping performances of these different configurations of the resonator necks, comparison study is conducted via developing a 2D linearized Navier–Stokes model of a duct in the presence of a grazing flow. A Helmholtz resonator is implemented on the duct as a side branch. The model is in frequency domain and it is validated first by comparing the numerical results with the experimental measurements available in the literature. The effects of 1) 3 extended neck configurations, 2) extended neck length and 3) the grazing flow Mach number are evaluated. It is shown that higher Mach number of the grazing flow, lower transmission loss. As the extended neck is in different configurations, the resonant frequencies and the maximum transmission losses are dramatically different, especially as the grazing Mach number M u is greater than 0.05, i.e. M u ≥ 0.05 . Approximately 20% resonant frequency shift is observed. The conventional design of Helmholtz resonator without an extended neck is found to perform much less effective than that of with concentric extended neck. The optimum design of the resonator neck can lead to 5–11 dB more transmission loss over a broader frequency range, especially at higher Mach number. The noise damping mechanism is visualized by the formation of vortex shedding at the neck of the resonator and sound energy is converted into kinetic energy being dissipated by the surrounding air. The present work opens up a predictive means to optimize the design of a Helmholtz resonator.

      PubDate: 2018-04-15T07:30:57Z
      DOI: 10.1016/j.ast.2018.03.002
      Issue No: Vol. 77 (2018)
       
  • Optimal direct adaptive soft switching multi-model predictive control
           using the gap metric for spacecraft attitude control in a wide range of
           operating points
    • Authors: Saman Saki; Hossein Bolandi; Seyed Kamal-e-ddin Mousavi Mashhadi
      Pages: 235 - 243
      Abstract: Publication date: June 2018
      Source:Aerospace Science and Technology, Volume 77
      Author(s): Saman Saki, Hossein Bolandi, Seyed Kamal-e-ddin Mousavi Mashhadi
      In this paper, a new Multi-Input Multi-Output (MIMO) Multi-model Predictive Control (MMPC) with direct adaptive structure is used for spacecraft attitude control in a wide range of operating points and maneuvers. Because of the highly nonlinear dynamics, the linearized model characteristics are extremely depended to the operating points. In such cases, an expected performance, in the wide range of the operating points, never can be achieved using a single controller and single model (even instability may be anticipated). To handle this problem, in this paper, we divide the whole operating range of the spacecraft to construct sub-models (model bank) using a mathematical tool called as gap metric. Next, an adaptive MPC based on sub-models is designed. In this procedure, there are two problems: stability when switching among models and the minimum number of sub-models. Hard switching among models to update the controller's model causes extreme chattering on the control signal and reference tracking. The motivation of this paper is to present a new solution for the mentioned problems. To solve the first problem (remove chattering), an adaptive soft switching law to tune the controller parameters, when selecting new model, based on the Lyapunov theory is introduced as the main novelty of this paper. This guarantees the stability of the closed loop control system. For solution of the second problem, the number of optimal sub-models is obtained using different simulations. Finally, the effectiveness of the suggested method is proven via simulations.

      PubDate: 2018-04-15T07:30:57Z
      DOI: 10.1016/j.ast.2018.03.001
      Issue No: Vol. 77 (2018)
       
  • Dealiasing harmonic balance method for obtaining periodic solutions of an
           aeroelastic system
    • Authors: Honghua Dai; Xiaokui Yue; Jianping Yuan; Dan Xie
      Pages: 244 - 255
      Abstract: Publication date: June 2018
      Source:Aerospace Science and Technology, Volume 77
      Author(s): Honghua Dai, Xiaokui Yue, Jianping Yuan, Dan Xie
      The harmonic balance method with a dealiasing scheme, referred to as DHB method, is proposed to obtain the semi-analytical periodic solutions of a two dimensional subsonic airfoil. The equations of motion of this autonomous system are a set of integro-differential equations. To solve this problem, the original model is first transformed into a system of nonlinear ordinary differential equations by means of integral transformations, and consequently this system is transformed into nonlinear algebraic equations (NAEs) via the harmonic balance method. Previous study demonstrated that the harmonic balance method may yield mathematically aliased solutions. To remedy this drawback, a dealiasing scheme, based on shifting the high order harmonic coefficients to the correct low order positions and properly scaling the frequency, is proposed to effectively suppress mathematical aliasing. As for solving the resultant NAEs, a series of scalar homotopy methods (SHMs) are introduced. The SHM methods are robust to initial conditions, and do not require the calculation of the inverse of Jacobian matrix. Therefore, the computing cost can be reduced and the inaccuracy arising from inverting the ill-conditioned Jacobian can be eliminated. Finally, numerical examples are carried out to verify the efficiency and accuracy of the present method.

      PubDate: 2018-04-15T07:30:57Z
      DOI: 10.1016/j.ast.2018.03.008
      Issue No: Vol. 77 (2018)
       
  • Acceleration autopilot design for gliding guided projectiles with less
           measurement information
    • Authors: Qiuping Xu; Sijiang Chang; Zhongyuan Wang
      Pages: 256 - 264
      Abstract: Publication date: June 2018
      Source:Aerospace Science and Technology, Volume 77
      Author(s): Qiuping Xu, Sijiang Chang, Zhongyuan Wang
      In this paper, a novel acceleration autopilot is proposed to solve the problem of the design of the flight control system of gliding guided projectiles under the factors such as cross-coupling dynamics, uncertainties, and constraints of the sensors, actuators, and system complexity. Unlike the traditional two/three-loop autopilot, only the measured accelerations are directly adopted as feedback in the proposed autopilot to reduce the cost and the system complexity, and to improve the reliability. The feasible and effective of the proposed autopilot is verified through several case studies. Results indicate that the designed autopilot can achieve quick, accurate, and no-overshoot tracking of the given signals, with good active-disturbance-rejection and decoupling performance and strong robustness and adaptability. The control parameters are easy and systematic to tune, and not sensitive to the perturbations of aerodynamic parameters within a wide range. In addition, the canard deflection commands change slowly from zero at the initial stage, and also yield a smooth and gentle rather than sharp change after each switching of acceleration signals, which can effectively avoid the control saturation and oscillation and enhance the flight stability.

      PubDate: 2018-04-15T07:30:57Z
      DOI: 10.1016/j.ast.2018.03.007
      Issue No: Vol. 77 (2018)
       
  • Dynamic analysis of assembled aircraft structures considering interfaces
           with nonlinear behavior
    • Authors: Santiago Hernández; Edoardo Menga; Pablo Naveira; Daniel Freire; Carlos López; Miguel Cid Montoya; Simón Moledo; Aitor Baldomir
      Pages: 265 - 272
      Abstract: Publication date: June 2018
      Source:Aerospace Science and Technology, Volume 77
      Author(s): Santiago Hernández, Edoardo Menga, Pablo Naveira, Daniel Freire, Carlos López, Miguel Cid Montoya, Simón Moledo, Aitor Baldomir
      Dynamic analysis is an essential part of aircraft design and the behavior of structural components of aircraft systems needs to be evaluated undergoing time varying loads. This is done by creating Finite Elements (FE) models and carrying out dynamic analysis in the frequency or time domain. The first approach is faster but only can be used in linear theory, namely when all materials involved in the structural model have linear constitutive equations. But sometimes FE models of aircraft structures contain a small number of non-linearities usually localized in the connection points between the assembled components. In this paper an approach that allows to perform frequency analysis in structural models having several nonlinear springs is presented and its performance is compared with the results provided by time domain analysis methods. An application example that represents a quite realistic assembly between two fuselage sections and a real aircraft connection system are presented.

      PubDate: 2018-04-15T07:30:57Z
      DOI: 10.1016/j.ast.2018.03.004
      Issue No: Vol. 77 (2018)
       
  • Development of an optimized trend kriging model using regression analysis
           and selection process for optimal subset of basis functions
    • Authors: Hakjin Lee; Duck-Joo Lee; Hyungil Kwon
      Pages: 273 - 285
      Abstract: Publication date: June 2018
      Source:Aerospace Science and Technology, Volume 77
      Author(s): Hakjin Lee, Duck-Joo Lee, Hyungil Kwon
      Surrogate modeling, or metamodeling, is an efficient way of alleviating the high computational cost and complexity for iterative function evaluation in design optimization. Accuracy is significantly important because optimization algorithms rely heavily on the function response calculated by surrogate model and the optimum solution is directly affected by the quality of surrogate model. In this study, an optimized trend kriging model is proposed to improve the accuracy of the existing kriging models. Within the framework of the proposed model, regression analysis is carried out to approximate the unknown trend of the true function and to determine the order of the universal kriging model, which has a fixed form with a mean structure dependent on the order of model. In addition, the selection of an optimal basis function is conducted to separate the useful basis function terms from the full set of the basis function. The optimal subset of the basis function is selected with the global optimization algorithm; which can accurately represent the trend of true response surface. The mean structure of proposed model has been optimized to maximize the accuracy of kriging model depending on the trend of true function. Two- and three-dimensional analytic functions and a practical engineering problem are chosen to validate the proposed model. The results showed that the OTKG model yield the most accurate responses regardless of the number of initial sample points, and can conversed into well-trained model with few additional sample points.

      PubDate: 2018-04-15T07:30:57Z
      DOI: 10.1016/j.ast.2018.01.042
      Issue No: Vol. 77 (2018)
       
  • Low-speed preconditioning for strongly coupled integration of
           Reynolds-averaged Navier–Stokes equations and two-equation turbulence
           models
    • Authors: M.S. Campobasso; M. Yan; A. Bonfiglioli; F.A. Gigante; L. Ferrari; F. Balduzzi; A. Bianchini
      Pages: 286 - 298
      Abstract: Publication date: June 2018
      Source:Aerospace Science and Technology, Volume 77
      Author(s): M.S. Campobasso, M. Yan, A. Bonfiglioli, F.A. Gigante, L. Ferrari, F. Balduzzi, A. Bianchini
      Computational fluid dynamics codes using the density-based compressible flow formulation of the Navier–Stokes equations have proven to be very successful for the analysis of high-speed flows. However, solution accuracy degradation and, for explicit solvers, reduction of the residual convergence rates occur as the local Mach number decreases below the threshold of 0.1. This performance impairment worsens remarkably in the presence of flow reversals at wall boundaries and unbounded high-vorticity flow regions. These issues can be resolved using low-speed preconditioning, but there exists an outstanding problem regarding the use of this technology in the strongly coupled integration of the Reynolds-averaged Navier–Stokes equations and two-equation turbulence models, such as the k − ω shear stress transport model. It is not possible to precondition only the RANS equations without altering parts of the governing equations, and there did not exist an approach for preconditioning both the RANS and the SST equations. This study solves this problem by introducing a turbulent low-speed preconditioner of the RANS and SST equations that does not require any alteration of the governing equations. The approach has recently been shown to significantly improve convergence rates in the case of a one-equation turbulence model. The study focuses on the explicit multigrid integration of the governing equations, but most algorithms are applicable also to implicit integration methods. The paper provides all algorithms required for implementing the presented turbulent preconditioner in other computational fluid dynamics codes. The new method is applicable to all low- and mixed-speed aeronautical and propulsion flow problems, and is demonstrated by analyzing the flow field of a Darrieus wind turbine rotor section at two operating conditions, one of which is characterized by significant blade/vortex interaction. Verification and further validation of the new method is also based on the comparison of the results obtained with the developed density-based code and those obtained with a commercial pressure-based code.

      PubDate: 2018-04-15T07:30:57Z
      DOI: 10.1016/j.ast.2018.03.015
      Issue No: Vol. 77 (2018)
       
  • Design of a variable Mach number wind tunnel nozzle operated by a single
           jack
    • Authors: Zheng Lv; Jinglei Xu; Feng Wu; Pengfei Chen; Juanjuan Wang
      Pages: 299 - 305
      Abstract: Publication date: June 2018
      Source:Aerospace Science and Technology, Volume 77
      Author(s): Zheng Lv, Jinglei Xu, Feng Wu, Pengfei Chen, Juanjuan Wang
      This paper presents the detailed design process of the variable Mach number nozzle operated by a single jack, and then a design case that covers a Mach number range of 1.5 to 2.0 is investigated. The computational fluid dynamics (CFD) approach is employed to obtain the flow quality of the test area at different operating conditions. The results show that the elastic deflection of the flexible plate can be calculated using the boundary conditions associated with the traditional nozzle design method. The flexible plate length affects the flow quality of the test section significantly and is set at 560 mm in the current study, corresponding to the variable thickness coefficient of 1.213. In addition, the standard deviation and flow angularity of the working section throughout the Mach number range can satisfy the requirements of the National Military Standard of China, which indicates that the design method of the single-jack nozzle is feasible.

      PubDate: 2018-04-15T07:30:57Z
      DOI: 10.1016/j.ast.2018.03.011
      Issue No: Vol. 77 (2018)
       
  • Axisymmetric nonlinear vibration analysis of sandwich annular plates with
           FG-CNTRC face sheets based on the higher-order shear deformation plate
           theory
    • Authors: R. Ansari; J. Torabi; E. Hasrati
      Pages: 306 - 319
      Abstract: Publication date: June 2018
      Source:Aerospace Science and Technology, Volume 77
      Author(s): R. Ansari, J. Torabi, E. Hasrati
      In this paper, a comprehensive numerical study is presented on the large-amplitude free vibration of sandwich annular plates integrated with functionally graded carbon nanotube-reinforced composite (FG-CNTRC) face sheets resting on elastic foundation. The sandwich plate is made of a homogeneous core and two FG-CNTRC face sheets whose material properties are estimated through a micromechanical model. Since the fundamental vibrational mode shapes of annular plates are axisymmetric, the governing equations are derived assuming the axisymmetric formulation. For this purpose, the quadratic form of total potential energy of the structure is presented based on the higher-order shear deformation theory (HSDT) of plates along with von-Karman nonlinear kinematic relations. The numerical differential and integral operators are then employed to discretize the energy functional in space and time domains. Finally, using the response of linear analysis and applying the pseudo-arc length continuation method, the nonlinear frequencies are obtained. After validating the results of proposed approach, detailed numerical results are given to analyze the effects of geometrical and material parameters on the nonlinear vibration of FG-CNTRC sandwich annular plates.

      PubDate: 2018-04-15T07:30:57Z
      DOI: 10.1016/j.ast.2018.01.010
      Issue No: Vol. 77 (2018)
       
  • Identification of the formation of resonant tones in compressible cavity
           flows
    • Authors: David Bacci; Alistair J. Saddington; Derek Bray
      Pages: 320 - 331
      Abstract: Publication date: June 2018
      Source:Aerospace Science and Technology, Volume 77
      Author(s): David Bacci, Alistair J. Saddington, Derek Bray
      Identification of the fluid dynamic mechanisms responsible for the formation of resonant tones in a cavity flow is challenging. Time-frequency non-linear analysis techniques were applied to the post-processing of pressure signals recorded on the floor of a rectangular cavity at a transonic Mach number. The results obtained, confirmed that the resonant peaks in the spectrum were produced by the interaction of a carrier frequency (and its harmonics) and a modulating frequency. High-order spectral analysis, based on the instantaneous wavelet bi-coherence method, was able to identify, at individual samples in the pressure–time signal, that the interaction between the fundamental frequency and the amplitude modulation frequency was responsible for the creation of the Rossier–Heller tones. The same technique was also able to correlate the mode switching phenomenon, as well as the deactivation of the resonant tones during the temporal evolution of the signal.

      PubDate: 2018-04-15T07:30:57Z
      DOI: 10.1016/j.ast.2018.03.013
      Issue No: Vol. 77 (2018)
       
  • Optimisation of adaptive shock control bumps with structural constraints
    • Authors: Edward Jinks; Paul Bruce; Matthew Santer
      Pages: 332 - 343
      Abstract: Publication date: June 2018
      Source:Aerospace Science and Technology, Volume 77
      Author(s): Edward Jinks, Paul Bruce, Matthew Santer
      This paper presents the results from a study to design an adaptive shock control bump for a transonic aerofoil. An optimisation framework comprising aerodynamic and structural computational tools has been used to assess the performance of candidate adaptive bump geometries based on a novel surface-pressure-based performance metric. The geometry of the resultant design is a unique feature of its adaptivity; being strongly influenced by the (passive) aerodynamic pressure forces on the flexible surface as well as the (active) displacement constraints. This optimal geometry bifurcates the shock-wave and carefully manages the recovering post-shock flow to maximise pressure-smearing in the shock-region with only a small penalty in L / D for the aerofoil. Short adaptive bumps (with small imposed displacements) generally perform better than taller ones, and maintain their performance advantage for a wide range of bump positions, suggesting good robustness to variations in shock position, which are an inevitable feature of a real-world flight application. Such devices may offer advantages over conventional (fixed geometry) shock control bumps, where optimal performance is achieved with taller devices, at the expense of poor robustness to variations in shock position.

      PubDate: 2018-04-15T07:30:57Z
      DOI: 10.1016/j.ast.2018.03.018
      Issue No: Vol. 77 (2018)
       
  • Generation and control of monodisperse bubble suspensions in microgravity
    • Authors: Pau Bitlloch; Xavier Ruiz; Laureano Ramírez-Piscina; Jaume Casademunt
      Pages: 344 - 352
      Abstract: Publication date: June 2018
      Source:Aerospace Science and Technology, Volume 77
      Author(s): Pau Bitlloch, Xavier Ruiz, Laureano Ramírez-Piscina, Jaume Casademunt
      A new experimental setup for the generation of homogeneous, monodisperse bubble suspensions in turbulent duct flows in microgravity has been designed and tested in drop tower experiments. The setup provides independent control of bubble size, void fraction and degree of turbulence. The device combines several slug-flow injectors that produce monodisperse bubble jets, with a turbulent co-flow that ensures homogeneous spatial spreading. Bubble separation in the scale of the most energetic eddies of the flow, and bubble size sufficiently smaller, ensure that turbulence is most efficient as a mechanism for spatial spreading of bubbles while preventing coalescence, thus optimizing the homogeneous and monodisperse character of the suspension. The setup works in a regime for which bubbles are spherical, but sufficiently large compared to the turbulent dissipative scales to allow for two-way coupling between bubbles and carrying flow. The volume fraction is kept relatively small to facilitate particle tracking techniques. To illustrate the potential uses of the method we characterize the statistics of bubble velocity fluctuations in steady regimes and we characterize the transient relaxation of the buoyancy-driven pseudo-turbulence when gravity is switched-off.

      PubDate: 2018-04-15T07:30:57Z
      DOI: 10.1016/j.ast.2018.03.009
      Issue No: Vol. 77 (2018)
       
  • Hybrid reliability analysis and optimization for spacecraft structural
           system with random and fuzzy parameters
    • Authors: Chong Wang; Hermann G. Matthies; Menghui Xu; Yunlong Li
      Pages: 353 - 361
      Abstract: Publication date: June 2018
      Source:Aerospace Science and Technology, Volume 77
      Author(s): Chong Wang, Hermann G. Matthies, Menghui Xu, Yunlong Li
      In aerospace engineering, the incomplete environment conditions have been realized to be a significant factor affecting the system safety assessment. Considering the hybrid random and fuzzy uncertainties in system inputs, the paper proposes a novel numerical procedure for reliability analysis and reliability-based optimization. Random parameters are adopted to denote the aleatory uncertainties with sufficient sample information; whereas fuzzy parameters are used to quantify the epistemic uncertainties associated with expert opinions. Using the level-cut operation, fuzzy parameters are converted into interval variables, and a satisfaction degree-based interval ranking strategy is utilized to precisely quantify the interval safety possibility. Then the system safety possibility is calculated by the multiple integral, where cut levels of different fuzzy parameters are treated as independent variables. Subsequently based on the given safety index, a hybrid reliability optimization model is established. To avoid the huge computational burden caused by nested-loop optimization, a modified interval Monte Carlo method (MIMC) is proposed for limit state function evaluation. Eventually, a numerical example about the refractory ceramics tile of spacecraft verifies the feasibility of proposed method for hybrid reliability analysis and optimization design in practical aerospace engineering.

      PubDate: 2018-04-15T07:30:57Z
      DOI: 10.1016/j.ast.2018.03.014
      Issue No: Vol. 77 (2018)
       
  • Refractive sail and its applications in solar sailing
    • Authors: Shahin Firuzi; Shengping Gong
      Pages: 362 - 372
      Abstract: Publication date: June 2018
      Source:Aerospace Science and Technology, Volume 77
      Author(s): Shahin Firuzi, Shengping Gong
      Radiation pressure can be generated by interactions of electromagnetic (EM) waves with matter. Conventional in-space photonic propulsion systems like solar sails or solar photon thrusters operate by reflection of EM waves. This paper introduces a new type of solar sail which generates thrust by means of refraction of light through a thin film composed of micro-prisms. The main feature of the proposed refractive sail is its relatively large tangential radiation pressure, generated at near-normal radiation incidence. A method for computation of radiation pressure, by having the direction and power of input and output light beams, was introduced. Then a simple analytical approach for optimal design of the refractive film was presented, and ray tracing was utilized for computation of the radiation pressure to a good approximation. A refractive sail can be utilized in applications which a tangential force, especially at near-normal radiation incidence, is required. By utilizing this sail for orbit raising from low-Earth orbit (LEO), the minimum possible altitude for solar sailing can be reduced to about 500 km under the mean solar activity. Attitude control of solar sails along the sail's normal axis is another possible application of refractive films. Refractive films can also be utilized as solar collector (Fresnel lens) in space, which besides the convenience of their shape keeping, can be designed to be passively stable at the Sun-pointing attitude.

      PubDate: 2018-04-15T07:30:57Z
      DOI: 10.1016/j.ast.2018.03.016
      Issue No: Vol. 77 (2018)
       
  • Performance of radial–axial clearance rim seal in realistic working
           conditions
    • Authors: Xingyun Jia; Hai Zhang; Yuting Jiang
      Pages: 373 - 387
      Abstract: Publication date: June 2018
      Source:Aerospace Science and Technology, Volume 77
      Author(s): Xingyun Jia, Hai Zhang, Yuting Jiang
      Performance of radial–axial clearance rim seal (RACS) in realistic working condition is investigated and compared with axial-clearance rim seal (ACS), radial-clearance rim seal (RCS) in this paper.
      Authors use the numerical method of conjugate heat transfer for calculation to accurately take heat transfer between the rotor–stator cavity flow and the solid discs into account. Results show that seal effectiveness and cooling effectiveness of RACS are the best when compared with ACS and RCS, the minimum mass flow rate for seal of RACS is 75% of that of RCS, and 34.6% of ACS. RACS has higher air-cooled aerodynamic efficiency, minimizing the mainstream performance penalty when compared with ACS and RCS. Corresponding to the respective minimum mass flow rate for seal, the air-cooled aerodynamic efficiency of RACS is 23.71% higher than that of ACS, and 12.79% higher than the RCS. Finite element analysis in turbine disc shows that RACS minimizes the flow rate of cooling air required for suppressing the radial growth of turbine disc in the three rim seal types. Mass flow rate of the required cooling air of RACS is approximately 40.9–52.9% of that of ACS, and 70–75% of RCS.

      PubDate: 2018-04-15T07:30:57Z
      DOI: 10.1016/j.ast.2018.03.024
      Issue No: Vol. 77 (2018)
       
  • Rectangular and skew shear buckling of FG-CNT reinforced composite skew
           plates using Ritz method
    • Authors: Y. Kiani; M. Mirzaei
      Pages: 388 - 398
      Abstract: Publication date: June 2018
      Source:Aerospace Science and Technology, Volume 77
      Author(s): Y. Kiani, M. Mirzaei
      Present research deals with the shear buckling behaviour of composite skew plates reinforced with aligned single walled carbon nanotubes (CNTs). Distribution of CNTs across the thickness of the skew plate are assumed to be uniform or functionally graded. Two different types of shear loads are considered. The case of rectangular shear which produces pure shear and the case of skew shear which results in a combined uniform shear and uniaxial tension/compression. Suitable for moderately thick plates, first order shear deformation plate theory is used to estimate the displacement field of the plate. The equivalent properties of the composite media are obtained by means of the refined rule of mixtures approach which contains efficiency parameters to capture the size dependent properties of the CNTs. With the aid of the Hamilton principle, transformation of the orthogonal coordinate system to an oblique one and the conventional Ritz method whose shape functions are constructed according to the Gram–Schmidt process, the stability equations of the plate are established and solved for two different types of loading, namely rectangular and skew shear loads. As shown, through introduction of a proper functionally graded pattern, i.e., FG-X pattern, the buckling load of the plate may be increased, significantly.

      PubDate: 2018-04-15T07:30:57Z
      DOI: 10.1016/j.ast.2018.03.022
      Issue No: Vol. 77 (2018)
       
  • Numerical study on the aerodynamic coupling effects of spinning and coning
           motions for a finned vehicle
    • Authors: Tianyu Lu; Xiaosheng Wu; Juanmian Lei; Jintao Yin
      Pages: 399 - 408
      Abstract: Publication date: June 2018
      Source:Aerospace Science and Technology, Volume 77
      Author(s): Tianyu Lu, Xiaosheng Wu, Juanmian Lei, Jintao Yin
      A numerical method was used to compute the flow over a four-finned vehicle under supersonic conditions to study the aerodynamic characteristics of a spinning finned vehicle in a coning motion. For each flow condition, the side force and side moment of a combined spinning and coning motion were computed at roll angles of 0°–90°. The results of the non-rolling, coning, and spinning motions were superposed by using a nonlinear aerodynamic model, and the superposed side force and side moment were compared with those of the combined motion. The results indicated significant aerodynamic coupling effects with the combined motion at low Mach numbers and high coning and spin rates. Flow analysis showed that the coupling effects were mainly produced by the shockwaves and expansion waves around fins.

      PubDate: 2018-04-15T07:30:57Z
      DOI: 10.1016/j.ast.2018.03.027
      Issue No: Vol. 77 (2018)
       
  • Nonlinear thermomechanical buckling and post-buckling response of porous
           FGM plates using Reddy's HSDT
    • Authors: Pham Hong Cong; Trinh Minh Chien; Nguyen Dinh Khoa; Nguyen Dinh Duc
      Pages: 419 - 428
      Abstract: Publication date: June 2018
      Source:Aerospace Science and Technology, Volume 77
      Author(s): Pham Hong Cong, Trinh Minh Chien, Nguyen Dinh Khoa, Nguyen Dinh Duc
      This work presents an analytical approach to investigate buckling and post-buckling behavior of FGM plate with porosities resting on elastic foundations and subjected to mechanical, thermal and thermomechanical loads. The formulations are based on Reddy's higher-order shear deformation plate theory taking into consideration Von Karman nonlinearity, initial geometrical imperfections, and Pasternak type of elastic foundations. By applying Galerkin method, closed-form relations of buckling loads and post-buckling equilibrium paths for simply supported plates are determined. Numerical results are carried out to show the effects of porosity distribution characteristics (Porosity-I and Porosity-II), geometrical parameters, material properties and elastic foundations on the mechanical, thermal and thermomechanical buckling loads and post-buckling resistance capacity of the porous FGM plates.

      PubDate: 2018-04-15T07:30:57Z
      DOI: 10.1016/j.ast.2018.03.020
      Issue No: Vol. 77 (2018)
       
  • Attitude tracking control for a space moving target with high dynamic
           performance using hybrid actuator
    • Abstract: Publication date: July 2018
      Source:Aerospace Science and Technology, Volume 78
      Author(s): Yun-Hua Wu, Feng Han, Mo-Hong Zheng, Feng Wang, Bing Hua, Zhi-Ming Chen, Yue-Hua Cheng
      Attitude tracking control for a space moving target (such as debris or malfunctioning satellite) is investigated in this paper, which is different from the traditional agile attitude maneuvering and tracking control, and is a challenging problem for attitude control system, requiring agility, large control torque output, and high dynamic accuracy, etc. The rapidly moving target and spacecraft pose several tough issues such as agile attitude tracking control and actuator configuration design. A novel attitude tracking strategy is proposed to tackle the dynamic imaging process, including three phases, earth observation, attitude adjustment and dynamic tracking phase. With the accomplishment of attitude adjustment, the spacecraft will point toward the target to start the imaging task. For the maneuvers in the attitude adjustment and tracking phases, a combined control strategy consisting of saturation controller and backstepping controller is proposed. The former one constrains the attitude angular velocity as well as the required momentum on the actuators during the initial phase, while the backstepping controller guarantees the control accuracy with high dynamic performance in the imaging phase. A hybrid momentum exchanging actuator consisting of Control Moment Gyro (CMG) and Reaction Wheel (RW) is introduced to satisfy the great control torque demand. Null motion strategy is derived for the hybrid actuator to deal with CMG singularity and RW saturation simultaneously. Numerical simulations have demonstrated the advantages of the hybrid actuator and the proposed attitude control strategy, which not only enables the spacecraft to maneuver rapidly but also guarantees the tracking accuracy.

      PubDate: 2018-04-23T15:32:25Z
       
  • Numerical study on the nonlinear resonant dynamics of carbon
           nanotube/fiber/polymer multiscale laminated composite rectangular plates
           with various boundary conditions
    • Abstract: Publication date: July 2018
      Source:Aerospace Science and Technology, Volume 78
      Author(s): Raheb Gholami, Reza Ansari, Yousef Gholami
      This work deals with the numerical investigation of the geometrically nonlinear resonant dynamics of carbon nanotube/fiber/polymer multiscale laminated composite (CNT-FPMLC) rectangular plates with different boundary conditions. It is assumed that a uniform distributed harmonic excitation load in the transverse direction is applied to the CNT-FPMLC plates. The material properties of multiscale composite are estimated by means of the fiber micromechanics and Halpin–Tsai relations. Furthermore, it is assumed that the carbon nanotubes (CNTs) are distributed uniformly and oriented arbitrarily through the epoxy resin matrix. Based upon the Mindlin plate theory and using the von Kármán hypotheses, the governing equations of motion for the in-plane and out-of-plane motions including the effects of geometric nonlinearity, rotary inertia and shear deformation are achieved by means of the Hamilton's principle. In the solution process, the nonlinear partial differential equations of motions and associated boundary conditions are discretized via the generalized differential quadrature (GDQ) and afterward converted into a Duffing-type nonlinear time-varying set of ordinary differential equations via a numerical Galerkin approach. Then, the time periodic discretization method and the pseudo-arc length continuation technique are employed to solve the obtained equations in order to achieve the frequency–response curves associated with nonlinear free and forced resonances for the CNT-FPMLC rectangular plates with various edge supports. Finally, the influences of important design parameters including the weight percentage of single-walled and multi-walled CNTs, volume fraction of fibers, CNT aspect ratio, plate geometry and boundary conditions on the nonlinear resonant dynamics and linear natural frequencies of CNT-FPMLC rectangular plate are investigated in the numerical results.

      PubDate: 2018-04-23T15:32:25Z
       
  • Morphing and growing micro unmanned air vehicle: Sizing process and
           stability
    • Abstract: Publication date: July 2018
      Source:Aerospace Science and Technology, Volume 78
      Author(s): M. Hassanalian, A. Quintana, A. Abdelkefi
      An optimized and comprehensive method is proposed in order to design an efficient micro unmanned air vehicle with morphing and growing capabilities. In the sizing process, to select the optimum wing shape, three different shapes are compared based on an aerodynamic analysis, and a tapered wing is selected for the compressed mode. Then, since the wingspan, wing area, wing loading, and other parameters are changing as function of time, a transition analysis is carried-out during the sizing process. By using the calculated surface area and considered aspect ratio for compression and expansion modes, the wingspan is determined as function of time. Considering the estimated weight, the required lift coefficient is calculated and then two types of airfoils are selected. Finally, after completing the optimal geometric design, aerodynamic analyses are carried out to investigate the performance of the growing drone. The proposed strategy for designing efficient micro unmanned air vehicles for a well-defined mission can be utilized and extended to design other growing micro unmanned systems depending on the mission.

      PubDate: 2018-04-23T15:32:25Z
       
 
 
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