ALBERT

Description: 〈p〉Publication date: October 2019〈/p〉 〈p〉〈b〉Source:〈/b〉 Aerospace Science and Technology, Volume 93〈/p〉 〈p〉Author(s): Zeyang Yin, Afzal Suleman, Jianjun Luo, Caisheng Wei〈/p〉 〈h5〉Abstract〈/h5〉 〈div〉〈p〉This work investigates the attitude tracking control problem of spacecraft under strong external disturbances and parameter uncertainties. A novel appointed-time stable control scheme is proposed with guaranteed transient and steady-state performance. First, an appointed-time reachable performance function (ARPF) is presented, and its reach time can be arbitrarily selected by the users. Then, a double-ARPFs strategy is introduced, that is, by imposing two ARPFs on the attitude and the system output, respectively, all system states will be appointed-time stable. Furthermore, a robust controller with implementable structure is proposed to guarantee the performance functions under strong external disturbances and parameter uncertainties. And the attitude tracking errors as well as the angular velocity errors are proved to be appointed-time stable. Last, three groups of simulations are organized to verify the effectiveness, robustness and appointed-time stability of the proposed control scheme.〈/p〉〈/div〉

Description: 〈p〉Publication date: October 2019〈/p〉 〈p〉〈b〉Source:〈/b〉 Aerospace Science and Technology, Volume 93〈/p〉 〈p〉Author(s): Fan Liu, Hong Yan, Wangjie Zhan, Yunpeng Xue〈/p〉 〈div xml:lang="en"〉 〈h5〉Abstract〈/h5〉 〈div〉〈p〉The effects of the steady and pulsed arc discharge on the oblique shock wave control are explored through an experimental study in a supersonic wind tunnel with the maximum design Mach number of 2.5. The oblique shock is formed by a compression ramp with an angle of 7 degree mounted on the floor of the wind tunnel. Four electrodes are placed upstream of the compression ramp and spaced equally in the spanwise direction to generate the discharge arcs. The steady discharge is generated between the electrodes and downstream ramp corner, forming streamwise arcs. The arc length can be modulated between 10 mm and 40 mm by moving the electrodes in the streamwise direction. With a High Frequency Switch (HFS), the pulsed discharge is achievable with a frequency range from 5 kHz to 50 kHz. The pulsed discharge is generated between two adjacent electrodes, forming so-called transversal arcs, which are blown downstream by the main flow. Results show that the steady discharge arcs act like a uniformly distributed conductor. With an increase of the arc length, the arc power increases, and the weakening effect on the shock is enhanced. For the pulsed discharge arcs, the shorter arcs are observed with higher discharge frequency, which implies a lower arc power. Overall, the weakening effect of the steady arcs on the shock is more effective with shock strength reduced by 4%, compared to the pulsed ones with only 0.35% reduction.〈/p〉〈/div〉 〈/div〉

Description: 〈p〉Publication date: October 2019〈/p〉 〈p〉〈b〉Source:〈/b〉 Aerospace Science and Technology, Volume 93〈/p〉 〈p〉Author(s): Jiachao Li, Guozhu Liang, Pingping Zhu, Xi Wang〈/p〉 〈h5〉Abstract〈/h5〉 〈div〉〈p〉In order to accurately predict the whole operating process of a liquid hydrogen tank under gaseous hydrogen pressurization, a 2-D axial symmetry Volume-of-Fluid (VOF) based numerical simulation method is established. Phase change and turbulence models are included in the numerical simulation. The variations of physical parameters such as the ullage mass, temperature and pressure, are carefully analyzed. The different effects are given based on simulations with and without phase change, and the comparison between feedback pressurization and open pressurization is also given. Compared with the NASA's experiment under the feedback pressurization, the simulation results show that the deviation of pressurant gas masses consumption is 11.0% during the whole operating process. The deviation of the total ullage mass is 〈math xmlns:mml="http://www.w3.org/1998/Math/MathML" altimg="si1.svg"〉〈mo linebreak="badbreak" linebreakstyle="after"〉−〈/mo〉〈mn〉0.8〈/mn〉〈mtext〉%〈/mtext〉〈/math〉, 1.4% and 7.6% for the ramp period, the hold period and the expulsion period, respectively. The deviation of phase change mass is 7.5% and 〈math xmlns:mml="http://www.w3.org/1998/Math/MathML" altimg="si2.svg"〉〈mo linebreak="badbreak" linebreakstyle="after"〉−〈/mo〉〈mn〉21.5〈/mn〉〈mtext〉%〈/mtext〉〈/math〉 for the ramp period and the expulsion period, respectively. The simulation results also reach an agreement with the experiment on the energy absorption proportions and demonstrate that most of the energy addition from the external environment and the pressurizing gas is absorbed by the tank wall. The liquid gains the least energy during the expulsion period. Temperature stratification appears along the axial direction in the surface liquid region and the ullage region, and the bulk liquid is in a subcooled state. The location of phase change mainly appears near the vapor-liquid interface, where the net condensation appears during the ramp period and the hold period, while the net vaporization appears during the expulsion period. The phase change increases the amplitude of temperature oscillation. The open pressurization has an ullage pressure peak and an average ullage temperature peak, which lead to large impacts on the tank structure, but the control of the inlet mass flow rate is easy to implement. The feedback pressurization could maintain a steady ullage pressure, but more pressurant gas masses are consumed, and the control of inlet mass flow rate becomes more complicated. The simulation results can be used as references for design optimization of the pressurization systems of cryogenic liquid launch vehicles in order to save pressurant gas masses and decrease the ullage pressure peak which could reduce the tank wall thickness and enhance the carrying capacity of liquid launch vehicles.〈/p〉〈/div〉

Description: 〈p〉Publication date: October 2019〈/p〉 〈p〉〈b〉Source:〈/b〉 Aerospace Science and Technology, Volume 93〈/p〉 〈p〉Author(s): Yongle Du, Yanchen Liu, John A. Ekaterinaris〈/p〉 〈h5〉Abstract〈/h5〉 〈div〉〈p〉High-order, low-dissipation and low-dispersion time-integration schemes are critical for efficient large-eddy simulations of sensitive time-dependent wave propagation problems, such as aeroacoustic phenomena. Based on the relative error analysis of the model equation 〈math xmlns:mml="http://www.w3.org/1998/Math/MathML" altimg="si1.svg"〉〈mi〉d〈/mi〉〈mi〉y〈/mi〉〈mo stretchy="false"〉/〈/mo〉〈mi〉d〈/mi〉〈mi〉t〈/mi〉〈mo linebreak="goodbreak" linebreakstyle="after"〉=〈/mo〉〈mo stretchy="false"〉(〈/mo〉〈mi〉μ〈/mi〉〈mo linebreak="badbreak" linebreakstyle="after"〉+〈/mo〉〈mi〉i〈/mi〉〈mi〉λ〈/mi〉〈mo stretchy="false"〉)〈/mo〉〈mi〉y〈/mi〉〈/math〉, a comprehensive analysis on the optimal low-dispersion low-dissipation diagonally implicit Runge-Kutta schemes is presented. Similar studies on optimal implicit Runge-Kutta schemes were proposed in existing publications to minimize the integrated error of the amplification factor in a pre-defined range of 〈em〉λ〈/em〉Δ〈em〉t〈/em〉 on the imaginary axis. In contrast, this study introduces the concept of “regions with acceptable amplification and phase-shift” of the numerical solutions in the complex 〈math xmlns:mml="http://www.w3.org/1998/Math/MathML" altimg="si2.svg"〉〈mo stretchy="false"〉(〈/mo〉〈mi〉μ〈/mi〉〈mo linebreak="badbreak" linebreakstyle="after"〉+〈/mo〉〈mi〉i〈/mi〉〈mi〉λ〈/mi〉〈mo stretchy="false"〉)〈/mo〉〈mi mathvariant="normal"〉Δ〈/mi〉〈mi〉t〈/mi〉〈/math〉 plane. As a result, the optimization aims to maximize the radii of the regions with acceptable dispersion and dissipation errors. When applied to the 2- and 3-stage diagonally implicit Runge-Kutta schemes, optimal A-stable or not-A-stable schemes with the low-dispersion and low-dissipation property and the 2nd- upto 4th-order accuracy are derived with equally good or better performances as compared to existing implicit Runge-Kutta schemes. Numerical experiments consistently demonstrate that the proposed criteria provide better indicators for the dispersion and dissipation errors of the time-integration schemes. Furthermore, the optimal schemes achieve the design order of accuracy with reasonably large time steps.〈/p〉〈/div〉

Description: 〈p〉Publication date: October 2019〈/p〉 〈p〉〈b〉Source:〈/b〉 Aerospace Science and Technology, Volume 93〈/p〉 〈p〉Author(s): Qing Wang, Ligang Gong, Chaoyang Dong, Kewei Zhong〈/p〉 〈h5〉Abstract〈/h5〉 〈div〉〈p〉This paper investigates the control problem of a morphing aircraft with variable sweep wings based on switched nonlinear systems and adaptive dynamic programming (ADP). The longitudinal altitude motion of the morphing aircraft is first modeled as switched nonlinear systems in lower triangular form. Then, the designed controller is comprised of the basic part and supplementary part. For the basic part, the backstepping technique is applied and a modified dynamic surface is introduced to overcome the ‘explosion of complexity’ problem. Disturbance observers inspired from the idea of extended state observer are designed to obtain the estimations of the internal uncertainties and external disturbances. The common virtual control laws of the backstepping method are developed by the disturbance observers and radial basis function neural networks. On the other hand, for the supplementary part, an ADP approach with the name of action-dependent heuristic dynamic programming is used to further decrease the altitude tracking error, which generates an additional control input by observing the differences between the actual and desired values in the backstepping design. Finally, comparative simulations are conducted to demonstrate the improved control performance of the proposed approach.〈/p〉〈/div〉

Description: 〈p〉Publication date: October 2019〈/p〉 〈p〉〈b〉Source:〈/b〉 Aerospace Science and Technology, Volume 93〈/p〉 〈p〉Author(s): J. Pinho, L. Peveroni, M.R. Vetrano, J.-M. Buchlin, J. Steelant, M. Strengnart〈/p〉 〈h5〉Abstract〈/h5〉 〈div〉〈p〉In the present study, a cryogenic valve used in launch vehicle liquid propulsion systems is experimentally and numerically characterized. Two independent measurement campaigns are performed with liquid nitrogen and water as working fluids. Two equivalent facilities have been designed and manufactured to perform the cryogenic and the water tests. The characteristic relationship between the volumetric flow rate and pressure drop across the test valve is obtained for both the fluids. As far as cryogenic tests are concerned, temperature measurements at the test valve inlet and outlet are presented as well as visualizations of the flow upstream the test section. The experimental results show that the test valve flow coefficient is independent of the working fluid provided single phase flow conditions. To further validate this result a numerical study is conducted using the commercial code CFD-ACE+. A good agreement between numerical and experimental results is found. Furthermore, test cases in the semi-critical and critical flow conditions are simulated using the so-called full cavitation model. The computed liquid recovery factor is shown to be also independent from the working fluid nature.〈/p〉〈/div〉

Description: 〈p〉Publication date: October 2019〈/p〉 〈p〉〈b〉Source:〈/b〉 Aerospace Science and Technology, Volume 93〈/p〉 〈p〉Author(s): Chengen Li, Guobiao Cai, Pengfei Wang, Hui Tian〈/p〉 〈h5〉Abstract〈/h5〉 〈div〉〈p〉A throttleable annular-gap axial injector, which could automatically adjust the injection gap area, is proposed to decrease injection pressure drop variation amplitude during wide range thrust regulation process in hybrid rocket motor. With the annular-gap injector adopted, liquid oxidizer is unhomogeneously injected into combustion chamber through a narrow annular injection gap. In conventional hybrid rocket motor, the oxidizer is usually homogeneously injected through full-inlet injection method. The injection area covers the whole fuel grain port. Difference of the injection methods greatly influences the oxidizer flow characteristics. Consequently, combustion and heat transfer characteristics of the motor are significantly changed. This paper is aimed to analyze the two-phase combustion flow field and coupled injector heat transfer characteristics of a lab-scale hybrid rocket motor with annular-gap injector through two-dimensional axisymmetric steady numerical simulations. The motor adopts 98% hydrogen peroxide and polyethylene as the propellants. Numerical analysis reveals that position of the injection gap influences the regression rate distribution in the first half of solid fuel grain but has little effect on that in the second half. The regression rate is relatively high when the injection gap is close to the fuel inner surface. In addition, the flowing liquid hydrogen peroxide in the injector could cool the chamber head. Sharp turns that produces vortex in oxidizer flow channel decreases the cooling effect and increases the overheating risk. Smooth bend could improve the injector heat transfer characteristics and eliminate the risk.〈/p〉〈/div〉

Description: 〈p〉Publication date: October 2019〈/p〉 〈p〉〈b〉Source:〈/b〉 Aerospace Science and Technology, Volume 93〈/p〉 〈p〉Author(s): Seyed Morteza Sajadmanesh, Mohammad Mojaddam, Arman Mohseni, Ali Nikparto〈/p〉 〈h5〉Abstract〈/h5〉 〈div〉〈p〉The flow-field inside a gas turbine engine, especially in the low-pressure turbine, is very complicated as it is normally accompanied by unsteady flow structures, strong and rapidly changing pressure gradients, intermittent transition of boundary layer, and flow separation and reattachment, especially during off-design performance. In this article, flow separation and reattachment on the suction side of an ultra-high-lift low-pressure turbine blade is studied and characterized using 3D Unsteady Reynolds-Averaged Navier-Stokes (URANS) equations. For turbulence modeling, transitional-SST method (〈em〉γ〈/em〉-〈math xmlns:mml="http://www.w3.org/1998/Math/MathML" altimg="si1.svg"〉〈msub〉〈mrow〉〈mi mathvariant="normal"〉Re〈/mi〉〈/mrow〉〈mrow〉〈mi〉θ〈/mi〉〈/mrow〉〈/msub〉〈/math〉) is adopted. The simulations are performed at the exit Reynolds numbers of 200,000 and 60,000, and at a constant isentropic exit Mach number of 0.4. The shape and extent of the separation bubble are primarily dependent on large vortical structures due to the Kelvin-Helmholtz instability and spanwise vortex tube shedding. Therefore, a better prediction of these phenomena could result in a more realistic separation bubble identification and consequently more accurate profile loss assessment. In order to better capture the transitional flow characteristics, which are not often readily available from conventional computational fluid dynamics simulations, the method of Proper Orthogonal Decomposition (POD) is used in this study. Non-coherent structures in the main flow, such as separation bubble, are investigated and studied. The POD modes of pressure-field are analyzed to clarify the generation of spanwise vortex tubes after separation point. In the higher Reynolds number, low-energy small-scale structures in the separation zone and downstream of the trailing edge are observed from the POD analysis. In the lower Reynolds number, high-energy large-scale structures shed from the separated shear layer are identified, which are responsible for increasing turbulent kinetic energy as well as increasing profile losses. This study also shows that the combination of URANS and POD can successfully be used to identify the separation bubble.〈/p〉〈/div〉

Description: 〈p〉Publication date: October 2019〈/p〉 〈p〉〈b〉Source:〈/b〉 Aerospace Science and Technology, Volume 93〈/p〉 〈p〉Author(s): Hanzhi Zhang, Ce Yang, Changmao Yang, Hang Zhang, Leilei Wang, Jiang Chen〈/p〉 〈h5〉Abstract〈/h5〉 〈div〉〈p〉Inlet bent torsional (BT) pipes are applied to complex turbocharger systems and industrial centrifugal compressors owing to space and weight constraints. To determine the interaction mechanism between the inlet BT pipe and outlet volute, and its effect on the centrifugal compressor performance and stability, three compressor models (model with clean inlet (M0 model) and models with distorted inlets induced by BT pipes (M1 and M2 models)) were chosen to conduct a performance experiment and static pressure measurement of the casing wall. The results show that the M1 and M2 cause a 6.4% increment and 22% reduction in the stable operating range, respectively. At the near-choke and peak efficiency points, both M1 and M2 can change the casing static pressure distribution evidently and thus influence the compressor pressure ratio and efficiency significantly. At the near-stall point, the high static pressure at the inlet region induced by the BT pipe of M1 can weaken the pressure peak strip induced by the volute. Therefore, both the considerable increase in circumferential pressure uniformity and the positive pre-whirl distortion of M1 are responsible for the delay of compressor stall. Finally, the interaction relationship of the inlet/outlet distortions in the entire operating range was sketched.〈/p〉〈/div〉

Description: 〈p〉Publication date: October 2019〈/p〉 〈p〉〈b〉Source:〈/b〉 Aerospace Science and Technology, Volume 93〈/p〉 〈p〉Author(s): Raj Nangia, Mehdi Ghoreyshi, Michel P.C. van Rooij, Russell M. Cummings〈/p〉 〈h5〉Abstract〈/h5〉 〈div〉〈p〉This article describes the aerodynamic design and assessment of a UCAV wing developed under the NATO-STO AVT-251 Multi-disciplinary Design Task Group. The wing design rationale and process is described, as well as how the idealized wing design produced by the design/optimization process is turned into a realizable aircraft geometry. Both the baseline and designed wings are analyzed in detail using several Navier-Stokes computational fluid dynamics flow solvers. Analysis at low speed and transonic conditions, both in the pitch plane and at sideslip conditions, is presented. Conclusions about the design are drawn and further improvements are suggested.〈/p〉〈/div〉

Description: 〈p〉Publication date: October 2019〈/p〉 〈p〉〈b〉Source:〈/b〉 Aerospace Science and Technology, Volume 93〈/p〉 〈p〉Author(s): Tong Liu, Kun Qu, Jinsheng Cai, Shucheng Pan〈/p〉 〈h5〉Abstract〈/h5〉 〈div〉〈p〉This paper develops a three-dimensional aircraft ice accretion model based on the numerical solution of the unsteady Stefan problem. In this model, for both rime and glaze ice situation, the heat conduction within the ice layer is considered as an unsteady process and the evolution of temperature distribution is numerically solved by Variable Space Grid (VSG) method. Unlike the Myers' state-of-the-art model and its extended model which introduces a posteriori correction factor 〈em〉α〈/em〉 by assuming a linear temperature profile, current model does not need to tune the parameter 〈em〉α〈/em〉 to match the experimental results. A hierarchical overset grid strategy is applied to ease the grid generation and shorten the workload on constructing complex icing configurations. Only the grid around the iced surface needs to be regenerated at each ice accretion step and complex flow characteristics near the ice horn can be captured accurately with such high-quality grid. First, simulations on NACA 0012 airfoil are conducted to validate current ice accretion model. Under rime and glaze ice condition, the predicted ice shapes agree well with experimental results. The predicted temperature evolution at iced surface reveals that ice accretion is a complex process influenced by heat and mass transfer. Then the computations on GLC-305 swept wing are performed to study the ice accretion process on the three-dimensional surface. The results show a good agreement with the experimental data. At the leading edge, the overall convective heat transfer coefficient and droplet collection efficiency increase along the direction from the wing root to the wing tip. This leads to very distinct ice horns at different wing sections.〈/p〉〈/div〉

Description: 〈p〉Publication date: Available online 27 March 2019〈/p〉 〈p〉〈b〉Source:〈/b〉 Aerospace Science and Technology〈/p〉 〈p〉Author(s): Runpei Jiang, Yun Tian, Peiqing Liu〈/p〉 〈h5〉Abstract〈/h5〉 〈div〉〈p〉The transonic buffet over large aircraft wings seriously affects flight safety and ride quality, thus it has been a focus issue for researchers to control transonic buffet through active or passive means. With traditional supercritical airfoil and wing, shock foot bubble and trailing edge separation usually exist in the pre-buffet conditions. The merging of the shock foot bubble and trailing edge separation is associated with buffet onset. In order to prevent shock foot bubble from merging with the trailing edge separation, this paper presents a method to alleviate the transonic buffet by applying the Buffet Breather, which is made of a punched hole near the trailing edge at the RAE2822 airfoil and Wing1 supercritical wing published by Drag Prediction Workshop-III, respectively. The buffet characteristics of the airfoil and wing with and without the Buffet Breather are analyzed with Unsteady Reynolds Averaged Navier-Stokes method. The results show that the Buffet Breather can not only eliminate the buffet phenomenon under two-dimensional condition, but also weaken the pressure oscillation on the wing. In terms of the static aerodynamic characteristics, the Buffet Breather will reduce the lift coefficient under lower angle of attack and deteriorate the lift-to-drag ratio. In the actual application, a control valve can be considered to be installed, to open and close the Buffet Breather in order to achieve the best control effect.〈/p〉〈/div〉

Description: 〈p〉Publication date: Available online 27 March 2019〈/p〉 〈p〉〈b〉Source:〈/b〉 Aerospace Science and Technology〈/p〉 〈p〉Author(s): Jacopo Serafini, Giovanni Bernardini, Roberto Porcelli, Pierangelo Masarati〈/p〉 〈h5〉Abstract〈/h5〉 〈div〉〈p〉This paper presents an original approach to structural health monitoring of helicopter rotors based on strain measurement on the blades. Three algorithms are presented, one in the time domain and two in the frequency domain. They are based on the analysis of the discrepancies between the strains on damaged and undamaged blades. Two damage types are considered: a mass unbalance at the tip, and a localized stiffness reduction. The performance of the proposed methods is assessed by numerical simulation using a multibody dynamic solver for comprehensive aeroelastic analysis of rotorcraft. The numerical investigation has highlighted the capability of all the presented techniques to detect damages in the blades, even of small entity, in both steady-flight and soft maneuvers. The reduced prediction accuracy in aggressive maneuvered flight suggests the use of these methods only in steady or quasi-steady flight conditions.〈/p〉〈/div〉

Description: 〈p〉Publication date: May 2019〈/p〉 〈p〉〈b〉Source:〈/b〉 Aerospace Science and Technology, Volume 88〈/p〉 〈p〉Author(s): Hodei Urrutxua, Claudio Bombardelli, José Manuel Hedo〈/p〉 〈div xml:lang="en"〉 〈h5〉Abstract〈/h5〉 〈div〉〈p〉The “ion-beam shepherd” (IBS) is a contactless active space debris removal technique, also applicable to asteroid deflection. In the design of an IBS mission many constraints need to be considered, which involve multiple trade-offs. These constraints can be expressed analytically as a function of certain design parameters, and conveniently displayed in a two-dimensional parameter space or design space. This construction yields a “matching chart”, which effectively provides a feasible design envelope, and enables to find graphically a suitable design point that satisfies all applicable operational constraints simultaneously, thus providing a powerful tool tailored for the preliminary design of an IBS mission.〈/p〉〈/div〉 〈/div〉

Description: 〈p〉Publication date: Available online 27 March 2019〈/p〉 〈p〉〈b〉Source:〈/b〉 Aerospace Science and Technology〈/p〉 〈p〉Author(s): Yushu Jin, Xu Xu, Qingchun Yang, Shaohua Zhu〈/p〉 〈h5〉Abstract〈/h5〉 〈div〉〈p〉Much attention has been drawn to liquid-propellant rocket engine with thrust that can be varied on demand in recent years. A GO〈sub〉2〈/sub〉/kerosene deep-throttling variable thrust rocket engine using the newly proposed gas-liquid pintle injector is presented. Three-dimensional numerical simulations are conducted to investigate the flame appearance and heat flux distribution of the engine. The SST k-omega model is used for modeling turbulence and single-step finite-rate kinetics are used for modeling combustion. A grid sensitivity study is performed to examine the validity of the simulation results. The effects of total mass flow-rate, kerosene droplet injection velocity and kerosene droplet size distribution on flame appearance and heat flux distribution of the combustor are studied in detail. Three entirely different flame appearances are observed and it correspondingly causes different heat flux distributions. Details of the flow field structure and mass fraction contour inside the combustor are discussed to explore the cause of heat flux distribution. The work will provide a reference for the design of thermal protection system of the variable-thrust rocket engine operating with a wide range of space.〈/p〉〈/div〉

Description: 〈p〉Publication date: Available online 30 March 2019〈/p〉 〈p〉〈b〉Source:〈/b〉 Aerospace Science and Technology〈/p〉 〈p〉Author(s): Peixu Guo, Zhenxun Gao, Zhengyuan Wu, Hongpeng Liu, Chongwen Jiang, Chunhian Lee〈/p〉 〈h5〉Abstract〈/h5〉 〈div〉〈p〉One of the current problems of the detached eddy simulation (DES) and delayed DES (DDES) methods is the inaccuracy in prediction of the separated shear layer flows, which mainly reflects in the delayed development of the shear layer. To correctly simulate the separated shear flow, various modified definitions of the DDES length scale are investigated and the improved definition of the subgrid eddy viscosity is introduced to deal with the delayed development problem for the first time. In this paper, the shear layer without the upstream boundary layer is studied in order to assess the subgrid model of the DDES in the simulation of shear layers. Large eddy simulation (LES) and experiment results are introduced to test the performance of different definitions of the DDES parameters. According to the results, both the definitions of the length scale and subgrid eddy viscosity coefficient should be improved to get the best prediction. DDES with the LES length scale and vorticity-correlated length scale as well as the modified definition of the subgrid eddy viscosity coefficient predict a more accurate development of the shear layer than DDES with the original definitions, which provides a possible solution to the delay development problem of the shear layer. Furthermore, the LES mode of DDES performs considerably differently from the Smagorinsky-LES, which shows the former subgrid model is essentially different from that of Smagorinsky-LES. Theoretical analyses are made in order to find out the quantitative difference of the subgrid eddy viscosity between the LES mode of DDES and Smagorinsky-LES.〈/p〉〈/div〉

Description: 〈p〉Publication date: Available online 28 March 2019〈/p〉 〈p〉〈b〉Source:〈/b〉 Aerospace Science and Technology〈/p〉 〈p〉Author(s): Ioannis Goulos, Marco Bonesso〈/p〉 〈h5〉Abstract〈/h5〉 〈div〉〈p〉The concepts of variable rotor speed and active blade twist constitute promising technologies in terms of improving the operational performance and environmental impact of rotorcraft. Modern civil helicopters typically operate using nearly constant main and tail rotor speeds throughout their operational envelope. However, previous research has shown that decreasing the main rotor speed within salient points of the operational envelope can result in a notable reduction of rotor power requirement, resulting in more efficient aircraft. This work aims to develop an integrated approach able to evaluate the potential improvements in fuel economy and environmental impact through optimum implementation and scheduling of variable rotor speed combined with active blade twist. A comprehensive rotorcraft analysis method is utilized, comprising models applicable to flight dynamics, rotor blade aeroelasticity, engine performance, gaseous emission prediction, and flight path analysis. A holistic optimization strategy comprising methods for Design of Experiment (DOE), Gaussian Process-based (GP) surrogate-modeling, and genetic optimization is developed. The combined framework is used to predict globally optimum variable rotor speed and active blade twist schedules resulting in minimum fuel consumption. The overall method is employed to assess the impact of the investigated concepts for a representative Twin-Engine Light (TEL) helicopter operating within realistic mission scenarios. The optimizations carried out suggest that variable rotor speed combined with active blade twist may result in mission fuel consumption and nitrogen oxides emission (〈math xmlns:mml="http://www.w3.org/1998/Math/MathML" altimg="si1.gif" overflow="scroll"〉〈mi〉N〈/mi〉〈msub〉〈mrow〉〈mi〉O〈/mi〉〈/mrow〉〈mrow〉〈mi〉x〈/mi〉〈/mrow〉〈/msub〉〈/math〉) reductions of the order of 5% and 8%, relative to the fixed rotor speed case. The developed method constitutes an enabling technology for the investigation of novel technologies at multiple levels of assessment including aircraft-engine and mission levels.〈/p〉〈/div〉

Description: 〈p〉Publication date: February 2019〈/p〉 〈p〉〈b〉Source:〈/b〉 Aerospace Science and Technology, Volume 85〈/p〉 〈p〉Author(s): Xiaoxu Kan, Songtao Wang, Ling Yang, Jingjun Zhong〈/p〉 〈h5〉Abstract〈/h5〉 〈div〉〈p〉In general, understanding the cascade flow loss mechanism is crucial for the design and optimisation of the cascade from the basically essential. The research objective of this study was to develop a highly loaded compressor linear cascade, and a numerical simulation was performed to obtain both the vortex structures and the flow loss of the cascade during the corner stall process. The results indicate that the curved blade affects the influence range of the vortex structures by changing the pressure gradients of the cascade, and subsequently changes the transport process of the vortex structure to low-energy fluid clusters, thus affecting the weight distribution of the flow loss. Additionally, the weight coefficients of the passage vortex and concentrated shedding vortex are merged together, and accounts for half of the total flow losses. Finally, the innovation of this paper is to propose a topological analysis method based on the accurate and quantitative identification of the singular points' positions serves as a useful research method to reveal the vortex dynamic mechanism of the weight distribution of the flow loss affected by the curved blade, and the vortex criterion of the corner stall is proposed additionally.〈/p〉〈/div〉

Description: 〈p〉Publication date: February 2019〈/p〉 〈p〉〈b〉Source:〈/b〉 Aerospace Science and Technology, Volume 85〈/p〉 〈p〉Author(s): Yue Meng, Wei Wang, Hao Han, Jingxuan Ban〈/p〉 〈h5〉Abstract〈/h5〉 〈div〉〈p〉This paper presents a visual/inertial integrated guidance method for UAV shipboard landing. The airborne vision system is utilized to track infrared cooperated targets on the ship and output their center coordinates in the image. Meanwhile, the attitude information of the UAV is obtained from airborne Inertial Measurement Unit (IMU). Extended Kalman Filter (EKF) is chosen to fuse the visual and inertial information. The UAV's position, attitude, and velocity relative to the runway and the ship motion information are estimated. The ship motion information is utilized to predict the position of the intended landing point at the touchdown moment through the Autoregressive algorithm. And the motion information of the UAV is used to calculate the deviation from the intended landing path. All the information is delivered to the flight control system to calculate control command. Simulation shows that the visual/inertial integrated landing guidance system achieves satisfied estimation and prediction results.〈/p〉〈/div〉

Description: 〈p〉Publication date: February 2019〈/p〉 〈p〉〈b〉Source:〈/b〉 Aerospace Science and Technology, Volume 85〈/p〉 〈p〉Author(s): Jisong Zhao〈/p〉 〈h5〉Abstract〈/h5〉 〈div〉〈p〉Wall shear stress (skin friction) is one of the fundamental surface quantities in aerodynamics but its measurement remains challenging. This paper studies the measurement of wall shear stress vector fields in supersonic flows with shock waves using the shear-sensitive liquid crystal coating (SSLCC) technique. A previously developed SSLCC technique was improved and used for the measurement of wall shear stresses in a supersonic jet flow. Experimental results show that the flow structures including shock waves in the supersonic jet flow were visualized by the SSLCC technique; furthermore, the shear stress vector field over a planar surface induced by the supersonic jet flow was measured by the SSLCC technique. The shock waves and the compression-expansion repetitions in the supersonic jet flow were captured, respectively, by analysing the shear direction field and the skin friction line pattern. This work demonstrates the capabilities of the SSLCC technique to visualize and measure wall shear stresses in supersonic flows with complex shock waves.〈/p〉〈/div〉

Description: 〈p〉Publication date: February 2019〈/p〉 〈p〉〈b〉Source:〈/b〉 Aerospace Science and Technology, Volume 85〈/p〉 〈p〉Author(s): Yong Xi, Yao Meng〈/p〉 〈h5〉Abstract〈/h5〉 〈div〉〈p〉This paper addresses the longitudinal control problem of air-breathing hypersonic vehicles subject to actuator faults, full state constraints and backlash-like hysteresis. To accomplish the full state constraints, the modified nonlinear mappings are proposed to transform the altitude subsystem to a new system without state constraints. An adaptive backstepping controller is developed for this new system, in which the command filters are embedded to eliminate the problem of “explosion of terms”. Then, the controller parameters are updated to compensate the dysfunction of failed actuators and attenuate the unfavorable effects resulted from the backlash-like hysteresis. It is proved that the proposed controller can guarantee that the tracking errors converge to an arbitrarily small residual set and the full state constraints are not violated. Simulation results are carried out to demonstrate the effectiveness of the proposed control scheme.〈/p〉〈/div〉

Description: 〈p〉Publication date: February 2019〈/p〉 〈p〉〈b〉Source:〈/b〉 Aerospace Science and Technology, Volume 85〈/p〉 〈p〉Author(s): Francisco J. Arias, Salvador De Las Heras〈/p〉 〈h5〉Abstract〈/h5〉 〈div〉〈p〉The basis of a novel braking technique by using the Phobos sands for landing large payloads on Mars is outlined. Here consideration is given to the utilization of the Phobos or Deimos regolith as material for aerobraking by discharging a load of sand at certain distance in front of the spacecraft during the descent manoeuver. Although immediately after getting rid the load of sand in front of the spacecraft they have a null relative velocity with the spacecraft, however, because the stronger atmospheric drag acting on the tiny particles of sand they will be promptly decelerated. As a result, the particles of sand will impact onto the front of the spacecraft with a velocity close to the terminal velocity of the spacecraft itself. By using a pusher-disc – or akin damping system, in front of the spacecraft the momentum exchange from the sand collisions will result in a braking force acting on the spacecraft. Due to the very small delta-v budget required to lift material from the surface of Phobos or Deimos to their transfer orbits, then a small amount of dedicated rocket chemical propellant brought from Earth could be transformed into a huge amount of sand lifted from the surface of Phobos of Deimos to their transfer orbits. The large thrust generated by the Sandbraking makes this technique propitious for landing of planetary bodies struggling against gravity.〈/p〉〈/div〉

Description: 〈p〉Publication date: February 2019〈/p〉 〈p〉〈b〉Source:〈/b〉 Aerospace Science and Technology, Volume 85〈/p〉 〈p〉Author(s): Hongkang Liu, Chao Yan, Yatian Zhao, Sheng Wang〈/p〉 〈h5〉Abstract〈/h5〉 〈div〉〈p〉The restart performance of the inlet is crucial to a hypersonic air-breathing propulsion vehicle. An active flow control method based on energy addition for the restart of hypersonic inlets is brought forward in this paper. Using Reynolds Averaged Navier–Stokes equations method, the effects of energy addition parameters on the restart performances of hypersonic inlets are investigated and the restart process is presented. This study verifies that the present method is feasible to improve the restarting capability of hypersonic inlets. The restart performances rely heavily on the energy addition parameters. Using energy addition properly, it usually yields a better inlet performance and more importantly enables the inlet restart again; otherwise, the large separation bubble still exists. Investigations on energy addition parameters further reveal that energy addition always increases the mass flow rate, and its center far from the cowl facilitates restarting the inlet and decreasing the stagnation pressure losses. It also appears advantageous to work with energy addition at a moderate power consumption and effective radius. Besides, the restart process indicates that the parabolic shock induced by heated region plays an important role on the restart performance. It alters the separation bubble, broadens the flow passage and enables more deflected air into the inlet. The adverse pressure gradients in the entrance are greatly changed as well. Finally, results also suggest that an appropriate effective time for energy addition is sufficient to restart the inlet. The variations of hysteresis loops show that energy addition prompts the separated bubble to disappear and the inlet to restart in advance. As a result, the restarting Mach number declines and the operation range of hypersonic inlets prominently enlarges.〈/p〉〈/div〉

Description: 〈p〉Publication date: February 2019〈/p〉 〈p〉〈b〉Source:〈/b〉 Aerospace Science and Technology, Volume 85〈/p〉 〈p〉Author(s): Xiangyu Wang, Xin Yu, Shihua Li, Jiyu Liu〈/p〉 〈h5〉Abstract〈/h5〉 〈div〉〈p〉In this paper, the position and yaw angle trajectory tracking control problem is studied for unmanned helicopters subject to both matched and mismatched disturbances. To achieve the trajectory tracking goal, a feedforward-feedback composite control scheme is proposed based on the combination of the generalized proportional integral observer and the block backstepping control techniques. The controller design process mainly consists of two stages. In the first stage, some generalized proportional integral observers are developed for the helicopter system to estimate the mismatched, matched disturbances and their (higher-order) derivatives. In the second stage, the composite controller is designed by integrating the block backstepping control method and the disturbance estimates together. The proposed composite scheme guarantees asymptotic tracking performances for the position and yaw angle of the helicopter to the desired trajectories even in the presence of fast time-varying disturbances. Numerical simulations demonstrate the effectiveness of the proposed composite control scheme.〈/p〉〈/div〉

Description: 〈p〉Publication date: March 2019〈/p〉 〈p〉〈b〉Source:〈/b〉 Aerospace Science and Technology, Volume 86〈/p〉 〈p〉Author(s): Daniel T. Banuti, Martin Grabe, Klaus Hannemann〈/p〉 〈div xml:lang="en"〉 〈h5〉Abstract〈/h5〉 〈div〉〈p〉We investigate the flow in planar microscale nozzles and find that design and analysis paradigms based on the assumption of a dominant isentropic core with moderate viscosity corrections are not valid. Instead, the flow downstream of the throat is dominated by boundary layers that may choke the flow to subsonic velocities. The geometrical expansion ratio is found to be essentially irrelevant, instead, the length from throat to exit plane is found to be a much more important design parameter. Full 3D simulations are required to predict the flow topology; thermophysical modeling of the expanding gas has a noticeable impact on predicted performance. An analytical estimation of the Knudsen number in the expanding flow is given, allowing to determine its values from the expansion pressure ratio. An axial thrust analysis suggests truncation of the nozzle, resulting in a predicted 30% increase in thrust and 30% increase in specific impulse compared to the baseline configuration. The work has been carried out within the European Commission co-funded PRECISE project which was focused on designing and testing a micro chemical propulsion system thruster prototype using catalytically decomposed hydrazine as propellant.〈/p〉〈/div〉 〈/div〉

Description: 〈p〉Publication date: October 2019〈/p〉 〈p〉〈b〉Source:〈/b〉 Aerospace Science and Technology, Volume 93〈/p〉 〈p〉Author(s): Guido Voss, Dominik Schaefer, Cyrille Vidy〈/p〉 〈h5〉Abstract〈/h5〉 〈div〉〈p〉In the present work, flutter stability studies of an unmanned flying-wing configuration are presented. For this purpose, different fidelity modeling methods (DLM, CFD-Euler and CFD-RANS) are considered. The dependence of flutter speeds on altitude and Mach number is examined, showing that aerodynamic potential-based methods cannot predict aerodynamic phenomena such as flow detachment occurring at high angles of attack. In this respect, it is important that flutter investigations industry-oriented calculation methods are compared with the results obtained by high-fidelity CFD methods.〈/p〉〈/div〉

Description: 〈p〉Publication date: November 2019〈/p〉 〈p〉〈b〉Source:〈/b〉 Aerospace Science and Technology, Volume 94〈/p〉 〈p〉Author(s): Yunzhi Chen, Ling Yang, Jingjun Zhong〈/p〉 〈h5〉Abstract〈/h5〉 〈div〉〈p〉This study investigates the secondary flow control effect of the single diminutive streamwise endwall fence with varying geometrical parameters in a highly-loaded compressor. To determine the effects of the fence parameters upon the flow field structure and aerodynamic performance, numerous fenced cases are numerically simulated at the aerodynamic design point. The results demonstrate that the optimum endwall fence obstructs the migration of the endwall boundary layer from the pressure side (PS) to the suction side (SS), which leads to a measurable reduction in the separation line of corner reverse flow. Moreover, the counter-rotating fence vortex (FV) induced by the endwall fence can suppress the passage vortex (PV) and enhance the momentum exchange between the endwall boundary layer and high momentum freestream fluid. For the negative side, the fence with a larger surface area (the thicker, taller or longer fence) is proved to bring greater additional loss to the cascade, which offsets its positive flow control effect. In this study, the fence device with a width of 0.5 mm, a height of 10% the inflow boundary layer thickness and a length of 75% blade chord provides a 1.55% reduction on the cascade total pressure loss, as well as optimizes secondary flow structure by weakening passage vortex and dissipating corner vortex (CV).〈/p〉〈/div〉

Description: 〈p〉Publication date: November 2019〈/p〉 〈p〉〈b〉Source:〈/b〉 Aerospace Science and Technology, Volume 94〈/p〉 〈p〉Author(s): Guoshuai Li, Takahiro Ukai, Konstantinos Kontis〈/p〉 〈h5〉Abstract〈/h5〉 〈div〉〈p〉This paper proposes and demonstrates a novel shock tube driven by commercially available detonation transmission tubing in a safe, repeatable, and controllable manner for laboratory scale experiments. A circular cross-sectional open-ended shock tube (inner-diameter 〈math xmlns:mml="http://www.w3.org/1998/Math/MathML" altimg="si1.svg"〉〈mi〉D〈/mi〉〈mo linebreak="goodbreak" linebreakstyle="after"〉=〈/mo〉〈mn〉22〈/mn〉〈mspace width="0.25em"〉〈/mspace〉〈mtext〉mm〈/mtext〉〈/math〉) driven by detonation transmission tubing was used to investigate the working principle of this novel shock tube using a dynamic pressure transducer and time-resolved shadowgraph photography. Specifically, the shock Mach number, repeatability, and flow structure generated from the tube exit were characterized. The experimental result shows that the flow structure including an initial shock wave, a vortex ring, an embedded shock, and an oblique shock pattern is similar to that of the conventional compressed-gas driven shock tubes. Furthermore, the shock tube has good repeatability of less than 2% with a shock Mach number up to 1.58. The versatile and cost-effective nature of the shock tube driven by detonation transmission tubing opens a new horizon for shock wave-assisted interdisciplinary applications.〈/p〉〈/div〉

Description: 〈p〉Publication date: October 2019〈/p〉 〈p〉〈b〉Source:〈/b〉 Aerospace Science and Technology, Volume 93〈/p〉 〈p〉Author(s): C. Grillo, F. Montano〈/p〉 〈h5〉Abstract〈/h5〉 〈div〉〈p〉One of the most important problem of autonomous flight for UAS is the wind identification, especially for small scale vehicles. This research focusses on an identification methodology based on the Extended Kalman Filter (EKF). In particular authors focus their attention on the filter tuning problem. The proposed procedure requires low computational power, so it is very useful for UAS. Besides it allows a robust wind component identification even when, as it is usually, the measurement data set is affected by noticeable noises.〈/p〉〈/div〉

Description: 〈p〉Publication date: October 2019〈/p〉 〈p〉〈b〉Source:〈/b〉 Aerospace Science and Technology, Volume 93〈/p〉 〈p〉Author(s): S. Zenkner, M. Trost, R. Becker, C. Voß〈/p〉 〈div xml:lang="en"〉 〈h5〉Abstract〈/h5〉 〈div〉〈p〉In this study, aspects of the propulsion design for a highly swept wing configuration with military application are presented. These include the design of the thermodynamic cycle and the automated optimization of the inlet geometry. First, the thermodynamic design process of a conventional turbofan engine is described, which is intended to meet the requirements of a generic UCAV configuration. In addition to the thrust requirements derived from the mission profile, other constraints such as maximum fuel consumption, available installation space and aerodynamic and structural limits dimension the model in preliminary design. The resulting engine data are used as boundary conditions for the CFD simulation of the associated engine intake. This CFD calculation is part of an optimization process chain for determining the inlet duct geometry with low total pressure loss and low engine visibility. In order to estimate the effects of the different total pressure losses on the mission fuel consumption, the results obtained using CFD calculations are fed back into the performance simulation. This allows quantifying the influence of the inlet geometry on the global engine parameter. Based on the investigations and obtained results, the final configurations for the engine and intake are selected.〈/p〉〈/div〉 〈/div〉

Description: 〈p〉Publication date: November 2019〈/p〉 〈p〉〈b〉Source:〈/b〉 Aerospace Science and Technology, Volume 94〈/p〉 〈p〉Author(s): Yang Liu, Yonggang Gao, Zexin Chai, Zhichao Dong, Chunbo Hu, Xiaojing Yu〈/p〉 〈h5〉Abstract〈/h5〉 〈div〉〈p〉Based on the previous experimental results, in order to improve the understanding of the mixing and heat release process of the internal flow field in the solid-fuel scramjet combustor, and to provide theoretical reference for the layout of the combustor configuration in the next step, the numerical simulation method of DES combined with the gas mixing degree, mixing layer thickness and mixing fraction was used to analyze quantitatively the microscopic structures and the mixing combustion enhancement mechanism in the combustor at different lobe sweep angles based on the influence of the change of lobe sweep angle on engine performance, and then the influence law was explained from the microscopic point of view. Results demonstrate that 1) based on the engine configuration in this paper, in the design process of staggered support plate structure, the lobe sweep angle should be based on the width and the straight section length of the support plate. And the engine performance improves with the decrease of the lobe sweep angle. 2) The growth of the reaction mixture layer is in four typical stages: the initial growth zone, the rapid growth zone, the saturated growth zone and the full development zone. 3) The downstream part of combustor is in the full development zone of the reaction mixing layer. The full development zone as the main area determining the heat release performance of the engine, the mixing degree increases with the decrease of the lobe sweep angle in the behind part of the combustor. Therefore, the pressure in the downstream flow passage of the combustor and the space of combustor occupied by the high pressure area increase with the decrease of the lobe sweep angle.〈/p〉〈/div〉

Description: 〈p〉Publication date: November 2019〈/p〉 〈p〉〈b〉Source:〈/b〉 Aerospace Science and Technology, Volume 94〈/p〉 〈p〉Author(s): Lei Wang, Chuang Xiong〈/p〉 〈h5〉Abstract〈/h5〉 〈div〉〈p〉Various uncertainties, which are usually time-dependent, affect the reliability of complicated engineering systems seriously. Considering the fact that only limited sample data of the uncertain variables can be obtained in engineering practice during the whole in-service time of multidisciplinary systems, the interval process model is introduced to model the time-dependent uncertain variables, and a non-probabilistic time-dependent reliability estimation model is proposed. In addition, a sequential multidisciplinary optimization and non-probabilistic time-dependent reliability assessment (SMO_NTRA) approach is developed to decouple the time-dependent reliability analysis from the multidisciplinary design optimization (MDO). In the framework of SMO_NTRA, the deterministic MDO and non-probabilistic time-dependent reliability analysis are executed in a sequential manner. Thus the computationally expensive double level optimization problem can be avoided and the efficiency can be greatly improved. Furthermore, the shifting distance of the constraint is calculated by bi-section method. Both numerical and engineering examples are employed to demonstrate the validity of the proposed method.〈/p〉〈/div〉

Description: 〈p〉Publication date: Available online 14 October 2019〈/p〉 〈p〉〈b〉Source:〈/b〉 Aerospace Science and Technology〈/p〉 〈p〉Author(s): Zhiyuan Cao, Xi Gao, Bo Liu〈/p〉 〈h5〉Abstract〈/h5〉 〈div〉〈p〉In order to exploit the control mechanism of endwall profiling (EP) on flow field and performance of compressors, this paper carried out investigations on non-axisymmetric endwall profiling (NAEP), axisymmetric endwall profiling (AEP) and the comparison with bowed blading in a highly-loaded compressor cascade. Firstly, the influence mechanism of NAEP on corner separation and its design strategies were investigated. Pattern effect of NAEP, axial effect and height effect of concave/convex of NAEP were studied. Secondly, the influence mechanism of AEP on corner separation and its design strategies were studied. Thirdly, the comparison of flow mechanism between endwall profiling (EP) and bowed blading was carried out. Results show that although all the NAEP patterns can reduce the cross-passage pressure gradient, they exhibit distinct different control effects on corner separation. It is indicated that the radially inward pressure gradient, rather than the cross-passage pressure gradient near the endwall, is the key to control corner separations. The endwall cross-passage pressure gradient remained after AEP. However, the corner separation was also eliminated effectively by AEP with a concave curvature, with loss coefficient reduced by 20.07%. At last, compared with bowed blading, the cascade with EPs exhibits a higher overall performance but a lower control effect on corner separation.〈/p〉〈/div〉

Description: 〈p〉Publication date: December 2019〈/p〉 〈p〉〈b〉Source:〈/b〉 Aerospace Science and Technology, Volume 95〈/p〉 〈p〉Author(s): Junjie Liu, Mingwei Sun, Zengqiang Chen, Qinglin Sun〈/p〉 〈h5〉Abstract〈/h5〉 〈div〉〈p〉This paper presents a novel fixed-time extended state observer (FXTESO)-based fixed-time output feedback control scheme for aircraft with thrust vector at high angle of attack. In order to enhance the robustness of the closed-loop control system, a FXTESO is designed to estimate and compensate model uncertainties, external disturbance and the strong coupling among different channels. Furthermore, theoretical analysis shows that the convergence time of FXTESO is independent of initial estimation errors. For the double integrator systems obtained by FXTESO feedback linearization, a continuous output feedback control is utilized to obtain expected performance and fixed-time stability. The stability analysis for the closed-loop aircraft control system is conducted by Lyapunov theory. The daisy chain method is adopted to realize the control allocation. Finally, several numerical simulations are provided to demonstrate the effectiveness and robustness of the proposed methodology.〈/p〉〈/div〉

Description: 〈p〉Publication date: December 2019〈/p〉 〈p〉〈b〉Source:〈/b〉 Aerospace Science and Technology, Volume 95〈/p〉 〈p〉Author(s): Shipeng Dong, Liang Li, Dingguo Zhang〈/p〉 〈h5〉Abstract〈/h5〉 〈div〉〈p〉The dynamic modeling and free vibrations of rotating functionally graded (FG) tapered cantilever beams with hollow circular cross-section are studied in this paper. To capture the additional dynamic stiffening terms, the axial shrinkage of the beam caused by the transverse displacement is considered. The dynamic equations of the system governing stretching motion, flapwise bending motion, and chordwise bending motion are derived via employing assumed modes method and Lagrange's equations. Based on the first order approximate coupling (FOAC) dynamic model, natural frequencies and mode shapes of the beam system are calculated by solving eigenvalue problem of the deduced dimensionless vibration equations. Influences of the angular speed, the hub radius, the slenderness ratio, the ratio of hollow radius to the root radius, the taper ratio, and the functional gradient index on natural frequencies are studied. Frequency veering and mode shape interaction are discussed when the bending-stretching mode coupling effect of the beam is considered.〈/p〉〈/div〉

Description: 〈p〉Publication date: December 2019〈/p〉 〈p〉〈b〉Source:〈/b〉 Aerospace Science and Technology, Volume 95〈/p〉 〈p〉Author(s): Fang Fang, Lei Qiu, Shenfang Yuan, Yuanqiang Ren〈/p〉 〈h5〉Abstract〈/h5〉 〈div〉〈p〉This paper proposes a dynamic probability modeling-based aircraft Structural Health Monitoring (SHM) framework which provides a modular hierarchical SHM architecture for the development of applicable aircraft SHM techniques. The problem of reliable damage monitoring under time-varying conditions, which is the main application obstacle, is fully considered in the framework by the double probability models combined with the short term and long term dynamic update of the models. To realize the SHM capability of the framework, an adaptive constructing method of Gaussian mixture model is proposed for stable and efficient probability modeling and the probability similarity between dynamically updated models is measured for normalized damage detection. The framework is realized by combining with the guided wave SHM technique and validated in a full-scale aircraft fatigue test which is an in-flight test simulated on ground. The cracks of the right landing gear spar and the left wing panel on the aircraft structure are monitored reliably under the fatigue load conditions.〈/p〉〈/div〉

Description: 〈p〉Publication date: December 2019〈/p〉 〈p〉〈b〉Source:〈/b〉 Aerospace Science and Technology, Volume 95〈/p〉 〈p〉Author(s): Korosh Khorshidi, Mahdi Karimi〈/p〉 〈h5〉Abstract〈/h5〉 〈div〉〈p〉This paper presents an analytical model for flutter analysis of sandwich plates with functionally graded face sheets in the thermal environment. The material properties of the face sheets are supposed to be temperature-dependent by a nonlinear distribution satisfying one-dimensional heat equation, and vary according to power law distribution in terms of the volume fractions along the thickness of the plate. The vibration of the sandwich plate is modeled on the basis of different plate theories including Mindlin theory, classical theory, exponential theory, third-order theory, sinusoidal theory, hyperbolic theory and fifth-order theory. Modified shear deformation theories applied herein are capable of considering inertia effects and transverse shear stresses. First order piston theory is utilized to model the aerodynamic load due to supersonic flow. The coupled governing equations are derived using Hamilton's principle and, Galerkin approach is applied to obtain vibrational characteristics and critical dynamic pressure of the system. Some comparisons with available results in the literature are performed to validate the present modeling, and excellent agreement is observed. Our attention is focused on analyzing the effects of different parameters such as thickness ratio, aspect ratio, thermal load, the thickness of face sheet and power law index on dynamic stability of the sandwich plate.〈/p〉〈/div〉

Description: 〈p〉Publication date: December 2019〈/p〉 〈p〉〈b〉Source:〈/b〉 Aerospace Science and Technology, Volume 95〈/p〉 〈p〉Author(s): Jialin Zheng, Juntao Chang, Jicheng Ma, Daren Yu〈/p〉 〈div xml:lang="en"〉 〈h5〉Abstract〈/h5〉 〈div〉〈p〉Mode transition control is a critical issue of Turbine-Based-Combined Cycle (TBCC) engines when the primary thrust provider changes from gas turbine engines to ramjets/scramjets. Improper control laws might incur unexpected propulsion performance and even lead to mission abortion. Existed control laws are designed from the perspective of the combined engine itself without aerodynamic/propulsive couplings considered. However, like scramjet-powered aircraft, those with TBCC engines propelled also have intricate couplings between the engine and the airframe, especially during TBCC mode transition. In this paper, a nonlinear longitudinal TBCC-powered dynamic vehicle model is derived from first principles with extra aerodynamic effects and adjustable surfaces involved when a typical TBCC engine is integrated into the airframe. Then, a control-oriented model with six control inputs is obtained by approximately expressing aerodynamic coefficients through curve-fitting methods. Trim results indicate that the total thrust required to maintain a steady flight increases at different steady stages during mode transition of the TBCC engine. One feasible control law is designed with dynamics of the integrated system considered for a typical mode transition process to reveal a way that control inputs should be regulated to maintain a steady flight during mode transition of the TBCC engine.〈/p〉〈/div〉 〈/div〉

Description: 〈p〉Publication date: December 2019〈/p〉 〈p〉〈b〉Source:〈/b〉 Aerospace Science and Technology, Volume 95〈/p〉 〈p〉Author(s): Zhe Hui, Yang Zhang, Gang Chen〈/p〉 〈h5〉Abstract〈/h5〉 〈div〉〈p〉The excellent flight ability of birds is closely related not only to the morphing skeleton structure that can cause large-scale geometrical changes of their wings but also to the discrete or discontinuous wing structure composed of many feathers. In this study, a bio-inspired morphing discrete wing inspired from a pigeon's wing structure was designed with bionic feathers, with the explicit aim of improving the aerodynamic performance of an unmanned aerial vehicle. The bio-inspired discrete wing structure, controlled by a morphing skeleton structure, can actively morph into different swept-wing configurations similar to the wing postures of the pigeon and maintain a discrete wing surface similar to the pigeon wing surface at the same time. The results reveal that the bio-inspired morphing UAV can always maintain an optimal lift-to-drag ratio at three different Reynolds numbers utilizing the symmetrical wing morphing. The asymmetrical wing morphing can well achieve rolling control of the UAV. Furthermore, compared with a continuous wing surface structure, the bio-inspired discrete wing surface structure not only can achieve the induced drag reduction of the UAV through effectively decreasing the wing-tip vortex strength but also improve the lateral stability of the UAV.〈/p〉〈/div〉

Description: 〈p〉Publication date: December 2019〈/p〉 〈p〉〈b〉Source:〈/b〉 Aerospace Science and Technology, Volume 95〈/p〉 〈p〉Author(s): Aykut Tamer, Vincenzo Muscarello, Pierangelo Masarati, Giuseppe Quaranta〈/p〉 〈h5〉Abstract〈/h5〉 〈div〉〈p〉This work presents the use of a modern helicopter simulation environment for the evaluation of the combined performance of several systems for helicopter ride quality assessment. The proposed framework can handle increasingly detailed aeroservoelastic helicopter models while providing great flexibility and versatility in modeling human biodynamic models for vibration evaluation as well as models of the vibration attenuation devices. A numerical model representative of a medium weight helicopter is used to demonstrate the approach. Lumped parameter models of seat-cushion and human biodynamics are dynamically coupled to the helicopter model to provide a more realistic estimate of the actual vibratory level experienced by the occupants. Two performance indicators are formulated, based on the acceleration of the seat locations and using the ISO-2631 standard: i) qualitative criteria and related vibration dose values of the individuals seated at prescribed locations of a fully occupied helicopter, and ii) an overall rating of the occupants inside the cabin, considering the most and least comfortable seating distributions as the number of occupants changes. To demonstrate the proposed method, three configurations of helicopter-specific passive vibration absorbers are considered.〈/p〉〈/div〉

Description: 〈p〉Publication date: December 2019〈/p〉 〈p〉〈b〉Source:〈/b〉 Aerospace Science and Technology, Volume 95〈/p〉 〈p〉Author(s): Di Liu, Xiyuan Chen, Yuan Xu, Xiao Liu, Chunfeng Shi〈/p〉 〈div xml:lang="en"〉 〈h5〉Abstract〈/h5〉 〈div〉〈p〉For SINS/CNS integrated navigation system, the CKF can perform well for state estimation in Gaussian noise. However, its performances are likely to degrade significantly under non-Gaussian noise conditions. To improve the robustness of the CKF against non-Gaussian noise, we propose an improved cubature Kalman filter, called the maximum correntropy generalized high-degree CKF (MCGHCKF). In the MCGHCKF, the generalized high-degree cubature rule is used to improve the filtering performance, and the maximum correntropy criterion is utilized to reduce the influence of non-Gaussian noise on state estimation. Simulation experiments illustrate the effectiveness and robustness of our algorithm.〈/p〉〈/div〉 〈/div〉

Description: 〈p〉Publication date: November 2019〈/p〉 〈p〉〈b〉Source:〈/b〉 Aerospace Science and Technology, Volume 94〈/p〉 〈p〉Author(s): Changchang Che, Huawei Wang, Qiang Fu, Xiaomei Ni〈/p〉 〈div xml:lang="en"〉 〈h5〉Abstract〈/h5〉 〈div〉〈p〉The development of airborne sensor monitoring and artificial intelligence technologies provides effective tools for precise prognostic and health management (PHM) of aircraft. This paper presents a PHM model which combines multiple deep learning algorithms for condition assessment, fault classification, sensor prediction, and remaining useful life (RUL) estimation of aircraft systems. A long short-term memory (LSTM) based recurrent network is used to predict multiple multivariate time series of sensors, and deep belief network (DBN) is applied to assess system condition and classify faults of aircraft systems. Then, the RUL can be estimated through the integration of condition assessment and sensor prediction. Finally, the proposed algorithm is validated experimentally using NASA's C-MAPSS dataset, and the results showed a lower error rate and deviation than traditional models.〈/p〉〈/div〉 〈/div〉

Description: 〈p〉Publication date: December 2019〈/p〉 〈p〉〈b〉Source:〈/b〉 Aerospace Science and Technology, Volume 95〈/p〉 〈p〉Author(s): Changzhu Wei, Xiaozhe Ju, Rong Wu, Yanfeng He, Yin Diao〈/p〉 〈h5〉Abstract〈/h5〉 〈div〉〈p〉The typical return flight profile of vertical take-off/vertical landing (VTVL) reusable launch vehicles (RLVs) mainly comprises flip maneuver phase, boost back phase, grid fins deploy phase, among which the boost back phase plays a determinant role in accurate landing of reusable launch vehicles. In order to achieve pinpoint terminal precision in the boost back phase, a geometry and time updaters-based arbitrary-yaw iterative explicit guidance method is presented in this paper. By adopting an analytic motion predictor and abandoning small yaw angle hypothesis, the arbitrary-yaw iterative guidance law is formulated to deal with large yaw guidance problem caused by initial deviations and long-time flight. To compensate for the guidance errors due to the ignorance of terminal position constraint in the baseline guidance, a geometry updater is developed to update the target based on analytical geometry relationship. Furthermore, considering that the terminal target continuously moves following the earth, the time updater is designed to determine the candidate virtual orbit according to the estimated time-to-go. Simulations under various nominal flight trajectories and different initial deviations as well as the Monte Carlo simulation are carried out. Results illustrate that the proposed guidance algorithm performs well, showing high precision, strong adaptability and robustness.〈/p〉〈/div〉

Description: 〈p〉Publication date: December 2019〈/p〉 〈p〉〈b〉Source:〈/b〉 Aerospace Science and Technology, Volume 95〈/p〉 〈p〉Author(s): Yan Qing Wang, Mei Wen Teng〈/p〉 〈h5〉Abstract〈/h5〉 〈div〉〈p〉This paper focuses on the free vibration of circular and annular three-dimensional graphene foam (3D-GrF) plates under various boundary conditions. The Chebyshev-Ritz method is developed to solve the present problem. Different types of foam distribution are considered and the effective elastic modulus and mass density vary along the thickness or radial directions of the plates. The Kirchhoff plate theory is employed to derive the energy equations of the 3D-GrF plates. The numerical results show that the developed method has good accuracy and stability for analyzing free vibration problem of circular and annular 3D-GrF plates. It is also found that the increase in foam coefficient leads to the decrease in natural frequencies of 3D-GrF plates. Among different types of foam distribution, the 3D-GrF-I results in the highest natural frequency while the 3D-GrF-II corresponds to the lowest natural frequency of circular and annular 3D-GrF plates for most cases, depending on the specific boundary condition and foam coefficient. Moreover, the foam coefficient, the boundary condition, and the foam distribution interact with each other and have coupled effect on free vibration characteristics of 3D-GrF plates.〈/p〉〈/div〉

Description: 〈p〉Publication date: December 2019〈/p〉 〈p〉〈b〉Source:〈/b〉 Aerospace Science and Technology, Volume 95〈/p〉 〈p〉Author(s): Qin Zhao, Guangren Duan〈/p〉 〈h5〉Abstract〈/h5〉 〈div〉〈p〉This paper investigates the post-capture control problem associated with trajectory tracking and inertia property identification for a six degree-of-freedom (6DOF) combined spacecraft system subject to input saturation. Post-capture of non-cooperative target will cause a large shift in the dynamics of combined spacecraft mainly resulting from the change of inertia properties. An adaptive tracking control law for combined spacecraft is designed based on the terminal sliding mode and dynamic surface control techniques, and the inertia properties can be identified simultaneously. To estimate mass and inertia matrix, the expression of parameter estimation error is obtained by introducing a group of auxiliary filtered variables. A saturation compensator is employed to deal with the input saturation. Within the Lyapunov framework, the proposed controller is proved to guarantee the finite-time convergence of both trajectory tracking and inertia property identification driven by continuous control forces. Numerical simulations are finally performed to demonstrate the effectiveness of the designed control law.〈/p〉〈/div〉

Description: 〈p〉Publication date: December 2019〈/p〉 〈p〉〈b〉Source:〈/b〉 Aerospace Science and Technology, Volume 95〈/p〉 〈p〉Author(s): Rongqiao Wang, Xi Liu, Dianyin Hu, Jianxing Mao〈/p〉 〈div xml:lang="en"〉 〈h5〉Abstract〈/h5〉 〈div〉〈p〉This paper presents a probabilistic analysis framework for the reliability evaluation of turbine disc considering the correlation of multi-failure modes. A system-level zone division method is first applied to decompose the whole structure into different serial zones. Due to the same input random variables, there is correlation between failure modes. Thus, a mathematical copula function method is introduced to quantify the correlation between failure modes after reliability calculation of separate zones, during which process, dependent random variables are transformed to independent ones using Nataf transformation method. Meanwhile, to guarantee the accuracy and efficiency of calculation, adaptive surrogate model based on local radial point interpolation method (LRPIM) is established in each zone. Two main failure modes, i.e., low cycle fatigue and creep-fatigue are considered during the reliability analysis on a turbine disc. The results reveal that the reliability of the turbine disc changes with the correlation between failure modes. Also, sensitivity analysis shows that rotating speed and maximum temperature are two dominant factors affecting the turbined disc's reliability. Finally, the comparisons among three methods including the proposed method, the zone-based method without considering correlation and Monte Carlo (MC) method based on physics of failure (POF) of correlation are conducted. It is demonstrated that the proposed method in this study is more efficient and accurate for evaluating structural reliability with multi-failure modes coupling. Moreover, the proposed method provides an available prospect for reliability-based design optimization of multiple failure structure, contributing to enhance reliability in mechanical design.〈/p〉〈/div〉 〈/div〉

Description: 〈p〉Publication date: December 2019〈/p〉 〈p〉〈b〉Source:〈/b〉 Aerospace Science and Technology, Volume 95〈/p〉 〈p〉Author(s): Chuanzhi Sun, Baosheng Wang, Yongmeng Liu, Xiaoming Wang, Chengtian Li, Hongye Wang, Jiubin Tan〈/p〉 〈div xml:lang="en"〉 〈h5〉Abstract〈/h5〉 〈div〉〈p〉Coaxiality and cylindricity are the important geometric parameters of the low-pressure turbine (LPT) shaft. And the measurement accuracy of coaxiality and cylindricity directly affect the rotary characteristics of the aero engine. Therefore, a cylindrical profile measurement model with five systematic errors is designed to improve the coaxiality and cylindricity measurement accuracy of the low-pressure turbine shaft in this paper, in which eccentricity, probe offset, probe radius, geometric axis tilt and guide rail tilt are considered. Besides, the influence of systematic error and the shaft radius on the residual error for the stepped low-pressure turbine shaft is analyzed as well. The evaluation results of coaxiality and cylindricity are obtained based on different measurement strategies and models. In order to verify the effectiveness of the cylindrical profile measurement model with five systematic errors in the paper, a rotary measuring instrument with high precision is built. Compared with the traditional cylindrical profile measurement model with two systematic errors, the measurement accuracy of the coaxiality and cylindricity by the cylindrical profile measurement model with five systematic errors proposed in this paper are improved by 2.9 μm and 8.18 μm, respectively in the condition of the optimal measurement strategy for the LPT shaft with large radius. The proposed method is suitable for small probe radius and large eccentricity error, probe offset error, geometric axis tilt error and guide rail tilt error, especially for the LPT shaft with large radius. The proposed method can be applied to error separation and tolerance allocation for multistage rotor.〈/p〉〈/div〉 〈/div〉

Description: 〈p〉Publication date: December 2019〈/p〉 〈p〉〈b〉Source:〈/b〉 Aerospace Science and Technology, Volume 95〈/p〉 〈p〉Author(s): Mingfang Zheng, Hongwei Ma, Yan Lyu, Cunfu He, Chao Lu〈/p〉 〈h5〉Abstract〈/h5〉 〈div〉〈p〉The formula system of the state-vector and Legendre polynomials hybrid method (SV-LPHM) was applied to produce the dispersion curves and mode shape for the general anisotropic multilayer composite cylinders with whatever the dissimilarities of the layer material properties. According to relevant literature's reports, traditional Legendre polynomial method was only able to deal with the multilayer system where the material properties of adjacent layers are not significantly changed. To overcome the drawback, we introduce the state vector method to reshape the wave equation, boundary conditions and interface continuous conditions of the multilayer hollow cylinder in a consistent manner. Moreover, expanding the displacement field by the Legendre polynomials, and then a complete and concise state matrix form is formed for dispersion equation after complex algebraic transformation. All the matrices involving the mass and stiffness have been deduced analytically by the recurrence relation and orthogonality of the Legendre polynomial. The abovementioned operation overcomes the cumbersome when applying traditional Legendre polynomial method to dealing with the interface displacement and stress continuity. Firstly, we outlined the derivation of the formalism of the hybrid method (SV-LPHM) for guided waves propagating in the anisotropic multilayer hollow cylinder with an arbitrary number of layers. And then we demonstrated the SV-LPHM technique on the composite pipe of isotropic material and anisotropic material, where the key-factors of phase speed, displacement, and stress distribution were assessed and elaborated thoroughly. The results confirm the exponential convergence of the SV-LPHM compared with finite element methods.〈/p〉〈/div〉

Description: 〈p〉Publication date: December 2019〈/p〉 〈p〉〈b〉Source:〈/b〉 Aerospace Science and Technology, Volume 95〈/p〉 〈p〉Author(s): Chenxi Li, Jing Wang, Zhendong Guo, Liming Song, Jun Li〈/p〉 〈h5〉Abstract〈/h5〉 〈div〉〈p〉An aero-mechanical multidisciplinary optimization was carried out for a high-speed centrifugal impeller, SRV2-O, by integrating a Self-adaptive Multi-Objective Differential Evolution (SMODE) algorithm, RANS solver technique, Finite Element Method (FEM), and data mining technique of analysis of variance (ANOVA). Specifically, the optimization of the impeller was conducted for the maximization of isentropic efficiency and the minimization of maximum stress. During optimization, constraints were imposed on the total pressure ratio at the optimization point and the mass flow rate at the choked point as well. Where, the former constraint intends to guarantee the working capability of the impeller while the latter tries to fix the working range. After optimization, six optimal Pareto solutions are finally obtained. The isentropic efficiency of the optimal solutions is increased by 2.07% at most while the maximum stress is decreased by 6.36% at most among the Pareto solutions. The better performance of optimal designs was demonstrated through detailed aerodynamic and mechanical analysis. Then, the ANOVA is used to explore the effects of variables on performance function in design space, it is found that, the design variables located at the meridian channel and the leading edge of the full blade have significant effect on aerodynamic performance. These variables are crucial to reduce the loss caused by shock wave at impeller inlet and the leakage flow at blade tip. Meanwhile, the design variables located near the leading edge of full blade root section have great effect on strength performance, as they are effective to decrease the bend of blades and thus reduce the maximum stress. Thereby, the better aeromechanical performance can be achieved by dedicated adjustment on the curves of both the shroud of meridional channel and leading edge of full blade. The results of ANOVA are consistent with the aero-mechanical analysis. Therefore, the effectiveness of aero-mechanical optimization and data mining framework is demonstrated.〈/p〉〈/div〉

Description: 〈p〉Publication date: May 2019〈/p〉 〈p〉〈b〉Source:〈/b〉 Aerospace Science and Technology, Volume 88〈/p〉 〈p〉Author(s): Yayun Shi, Tihao Yang, Junqiang Bai, Lei Lu, Hui Wang〈/p〉 〈h5〉Abstract〈/h5〉 〈div〉〈p〉A transonic wind tunnel transition test is implemented on a fuselage-wing configuration with the sweep angle 〈math xmlns:mml="http://www.w3.org/1998/Math/MathML" altimg="si1.gif" overflow="scroll"〉〈msup〉〈mrow〉〈mn〉35〈/mn〉〈/mrow〉〈mrow〉〈mo〉∘〈/mo〉〈/mrow〉〈/msup〉〈/math〉 for commercial aircraft. The wide range of angle of attack from 〈math xmlns:mml="http://www.w3.org/1998/Math/MathML" altimg="si2.gif" overflow="scroll"〉〈mo〉−〈/mo〉〈msup〉〈mrow〉〈mn〉3.69〈/mn〉〈/mrow〉〈mrow〉〈mo〉∘〈/mo〉〈/mrow〉〈/msup〉〈/math〉 to 〈math xmlns:mml="http://www.w3.org/1998/Math/MathML" altimg="si3.gif" overflow="scroll"〉〈msup〉〈mrow〉〈mn〉3.07〈/mn〉〈/mrow〉〈mrow〉〈mo〉∘〈/mo〉〈/mrow〉〈/msup〉〈/math〉 assures that with increasing angle of attack, the laminar to turbulent transition dominant factor varies from cross-flow (CF) vortices to Tollmien-Schlichting (TS) waves. With linear stability theory, the limiting N-factors are calibrated based on the pressure distribution by experiment or the Reynolds Averaged Navier-Stokes (RANS) solver using the fixed experimental transition location. The pressure distribution of the RANS solver agrees well with the experiment in general except some small discrepancies, which causes deviation by 0.6 for the limiting TS N-factor. The RANS solver and the stability analysis provide the limiting N-factors of 7.0 and 8.7 for CF-vortices and TS-waves at the two sides, respectively. In the between, the TS value decays with the CF value due to their interaction. Thus, the transition criterion for limiting N-factors is established for the laminar prediction tool of 〈math xmlns:mml="http://www.w3.org/1998/Math/MathML" altimg="si4.gif" overflow="scroll"〉〈msup〉〈mrow〉〈mi〉e〈/mi〉〈/mrow〉〈mrow〉〈mi〉N〈/mi〉〈/mrow〉〈/msup〉〈/math〉 method at similar transonic wind tunnel. With the transition criterion, the transition location difference for 95% cases between the simulation and the experimental data is lower than 5% chord. The good match illustrates that the transition tool is accurate and robust for engineering applications, and also verifies the reasonability of the limiting N-factors. Therefore, the transition criteria at similar transonic conditions and well-performed 〈math xmlns:mml="http://www.w3.org/1998/Math/MathML" altimg="si4.gif" overflow="scroll"〉〈msup〉〈mrow〉〈mi〉e〈/mi〉〈/mrow〉〈mrow〉〈mi〉N〈/mi〉〈/mrow〉〈/msup〉〈/math〉 transition tool can be applied for the future laminar wing design.〈/p〉〈/div〉

Description: 〈p〉Publication date: May 2019〈/p〉 〈p〉〈b〉Source:〈/b〉 Aerospace Science and Technology, Volume 88〈/p〉 〈p〉Author(s): Aqib Khan, Rohit Panthi, Rakesh Kumar, S. Mohammed Ibrahim〈/p〉 〈h5〉Abstract〈/h5〉 〈div〉〈p〉Plug and ramp nozzles offer many advantages over the conventional converging-diverging counterpart and have received much attention in the recent decades. Variants of plug nozzles have been investigated for a wide range of operating conditions. The performance of these nozzles is dependent on the flow development on the plug or ramp surface, which in turn is greatly influenced by the cowl geometry. Experiments are conducted to study the planar plug nozzle flowfield for Mach 1.8 and 2.2 using half nozzle geometry. The work primarily focuses on the influence of the cowl length on the flow evolution on the plug surface at different nozzle pressure ratios (NPRs). The cowl of the outer nozzle is extended to 10, 30, 50 and 100% of the full plug length. This allows the supersonic flow to partially expand internally ahead of the throat section. Schlieren images are used to visualize the wave structure at different pressure ratios for different cowl lengths. For low NPRs, the nozzle with extended cowl behaves more like a conventional planar nozzle with strong shock waves. It is observed that the cowl length influences the pressure distribution on the plug surface only for low NPRs. The effect of side walls on the flow field of planar plug nozzles is also studied.〈/p〉〈/div〉

Description: 〈p〉Publication date: April 2019〈/p〉 〈p〉〈b〉Source:〈/b〉 Aerospace Science and Technology, Volume 87〈/p〉 〈p〉Author(s): Xuerui Wang, Sihao Sun, Erik-Jan van Kampen, Qiping Chu〈/p〉 〈h5〉Abstract〈/h5〉 〈div〉〈p〉This paper proposes an Incremental Sliding Mode Control driven by Sliding Mode Disturbance Observers (INDI-SMC/SMDO), with application to a quadrotor fault tolerant control problem. By designing the SMC/SMDO based on the control structure of the sensor-based Incremental Nonlinear Dynamic Inversion (INDI), instead of the model-based Nonlinear Dynamic Inversion (NDI) in the literature, the model dependency of the controller and the uncertainties in the closed-loop system are simultaneously reduced. This allows INDI-SMC/SMDO to passively resist a wider variety of faults and external disturbances using continuous control inputs with lower control and observer gains. When applied to a quadrotor, both numerical simulations and real-world flight tests demonstrate that INDI based SMC/SMDO has better performance and robustness over NDI based SMC/SMDO, in the presence of model uncertainties, wind disturbances, and sudden actuator faults. Moreover, the implementation process is simplified because of the reduced model dependency and smaller uncertainty variations of INDI-SMC/SMDO. Therefore, the proposed control method can be easily implemented to improve the performance and survivability of quadrotors in real life.〈/p〉〈/div〉

Description: 〈p〉Publication date: May 2019〈/p〉 〈p〉〈b〉Source:〈/b〉 Aerospace Science and Technology, Volume 88〈/p〉 〈p〉Author(s): Yang Lu, Taoyong Su, Renliang Chen, Pan Li, Yu Wang〈/p〉 〈h5〉Abstract〈/h5〉 〈div〉〈p〉The aerodynamic environment during the flight of a helicopter is complex because of the severe aerodynamic interaction between various components. To fully consider the effect of aerodynamic interaction in the initial stages of helicopter design and to eliminate or reduce its adverse effects, a comprehensive design optimization method for the aerodynamic layout of a helicopter that is capable of reducing the adverse effects of aerodynamic interaction is developed in this paper. To satisfy the requirements for precision and efficiency in the calculation model, an aerodynamic interaction analysis model of various helicopter components was established based on a viscous vortex particle and the unsteady panel hybrid method. To simultaneously consider the influences of the position and shape of the aerodynamic components on the aerodynamic interaction during the optimization process, parameter modeling of the helicopter's shape was performed based on the class function/shape function transformation (CST) method. A Kriging surrogate model of the objective function was further developed and combined with a hybrid sequential quadratic algorithm and genetic algorithm optimization strategy to establish a comprehensive optimization flow for the aerodynamic layout of a helicopter that reduces the adverse effects of aerodynamic interaction. Verification was carried out based on a fuselage shape derived from UH-60 helicopter. The optimization results showed that the use of the comprehensive optimization method for the aerodynamic layout of a helicopter can effectively reduce the adverse effects of aerodynamic interaction. Based on the optimization objectives, the efficiency of hovering increased by 4.7%, the hovering ceiling increased by 3.48%, the speed stability derivative increased by 264.7%, and the angle of attack stability derivative decreased by 26.4%.〈/p〉〈/div〉

Description: 〈p〉Publication date: May 2019〈/p〉 〈p〉〈b〉Source:〈/b〉 Aerospace Science and Technology, Volume 88〈/p〉 〈p〉Author(s): Chengyu Zhang, Volker Gümmer〈/p〉 〈div xml:lang="en"〉 〈h5〉Abstract〈/h5〉 〈div〉〈p〉Incorporating recuperators into gas turbines shows considerable potential for lower emissions and fuel consumption. Nowadays the technology readiness of advanced compact heat exchanger has provided a solid foundation for the availability of lightweight, higher efficient recuperators which would find good acceptance on the rotorcraft without penalizing the operational capabilities. To understand the impact of recuperator on the whole system for further development of future recuperated helicopter, it is proposed to evaluate the potential of recuperated helicopter turboshaft engines with emphasis placed on highly effective primary surface recuperator. This paper presents a comprehensive multidisciplinary simulation framework, and the aircraft configuration selected is a generic helicopter, which is similar to the helicopter Bo105, equipped with two Allison 250-C20B turboshaft engine variants. The improved part-load performance against the reference non-recuperated cycle is discussed first, followed by the analysis and evaluation of two representative flight missions. The study is finally extended to quantify the flight time required to compensate for the additional recuperator weight under the flight condition of 0-250 km/h and 0-3000 m for different recuperator design effectiveness values. It is suggested that the selection of recuperator effectiveness should be dependent on the most commonly involved mission profile and flight duration, in order to offset the added parasitic weight of the recuperator. The established rotorcraft multidisciplinary framework proves to be an effective tool to conduct a comprehensive assessment for the recuperated helicopter under a wide range of flight conditions as well as at mission level.〈/p〉〈/div〉 〈/div〉

Description: 〈p〉Publication date: May 2019〈/p〉 〈p〉〈b〉Source:〈/b〉 Aerospace Science and Technology, Volume 88〈/p〉 〈p〉Author(s): Srisha M.V. Rao, S.K. Karthick〈/p〉 〈h5〉Abstract〈/h5〉 〈div〉〈p〉Dynamic modal analysis enables new insights into the spatio-temporal dynamics of complex flow scenarios. Time resolved schlieren imaging provides significant information in compressible flow scenarios on flow structures and their evolution. We conduct a systematic study using synthetic images and experimental schlieren images on the effect of image acquisition parameters on the modal analysis by dynamic mode decomposition (DMD). We consider the effect of two important capture parameters – the capture rate (〈math xmlns:mml="http://www.w3.org/1998/Math/MathML" altimg="si1.gif" overflow="scroll"〉〈msub〉〈mrow〉〈mi〉f〈/mi〉〈/mrow〉〈mrow〉〈mi〉s〈/mi〉〈/mrow〉〈/msub〉〈/math〉) and the exposure time (〈math xmlns:mml="http://www.w3.org/1998/Math/MathML" altimg="si2.gif" overflow="scroll"〉〈msub〉〈mrow〉〈mi〉t〈/mi〉〈/mrow〉〈mrow〉〈mi〉e〈/mi〉〈mi〉x〈/mi〉〈mi〉p〈/mi〉〈/mrow〉〈/msub〉〈/math〉). Analysis is carried out on two sets of synthetic images, SI-I, an unsteady wavy interface created using a linear combination of sinusoids, and SI-II – hypothetical shock oscillations. Finally, a flapping supersonic jet is observed using high-speed schlieren with a nano-pulsed laser light source with three different imaging parameters. We find that among the two parameters the effect of exposure time on modal analysis and its interpretation is more pronounced than capture rate. An exposure time of 5% of maximum exposure produces 8% reduction in mode amplitude, and in case of long exposure the dynamic significance of modes undergoes complete change. If the flow images are instantaneous, then the spatial mode shapes of dominant modes remain the same irrespective of the capture rate. Aliasing has to be considered in sub-Nyquist capture rates, however, the actual frequencies can be suitably resolved.〈/p〉〈/div〉

Description: 〈p〉Publication date: May 2019〈/p〉 〈p〉〈b〉Source:〈/b〉 Aerospace Science and Technology, Volume 88〈/p〉 〈p〉Author(s): Rei Yamashita, Kojiro Suzuki〈/p〉 〈h5〉Abstract〈/h5〉 〈div〉〈p〉This paper describes the world's first successful simulation for lateral cutoff phenomena of sonic boom far from the flight path due to variation in atmospheric temperature with altitude. A flow field around an axi-symmetric paraboloid has been analyzed by the full-field simulation method that solves the three-dimensional Euler equations with a gravity term to create a horizontally stratified atmosphere. A solution-adapted structured grid is constructed to align the grid lines with the front and rear shock-wave surfaces in the entire domain, including the near field around a supersonic body and far field reaching the ground beyond lateral cutoff. The flight is assumed to have a speed of Mach 1.2 at an altitude of 10 km, and the computational domain ranges over a distance of 30 km from the axis of symmetry. The computational results show that the evanescent wave in the shadow zone beyond lateral cutoff decays exponentially and changes into a progressive rounding waveform. The characteristics of the waveform transition are in good agreement with those observed in the flight tests. Therefore, the full-field simulation is recognized as a promising approach for investigating sonic boom strength in the full extent of sonic boom noise, including lateral cutoff and evanescent waves. Moreover, the computational results clarify that sonic boom focusing occurs above the ground, except for the vicinity of the ground, and the focusing strength along the lateral cutoff curve detected from the three-dimensional shock-wave surface increases with altitude. The results of ray tracing analysis collaborate the reasonability of the simulation results, and the caustic of downward convex agrees well with the lateral cutoff curve. In the shadow zone, the magnitude of exponential decay increases with altitude, and the lateral distance where the pressure rise decreases rapidly shortens with altitude.〈/p〉〈/div〉

Description: 〈p〉Publication date: April 2019〈/p〉 〈p〉〈b〉Source:〈/b〉 Aerospace Science and Technology, Volume 87〈/p〉 〈p〉Author(s): Emanuele L. de Angelis, Fabrizio Giulietti, Goele Pipeleers, Gianluca Rossetti, Ruben Van Parys〈/p〉 〈h5〉Abstract〈/h5〉 〈div〉〈p〉This paper addresses a methodology for autonomous motion planning of multirotor aircraft in obstructed environments. The control strategy allows the vehicle to online generate quasi-optimal trajectories with limited computational load while performing collision avoidance tasks. The problem is formulated in a model-predictive control architecture in which motion planning and trajectory tracking processes are solved separately. The first process is based on a spline path planning approach to generate smooth and safe trajectories. The second process elaborates trajectory inputs in terms of commanded thrust magnitude and desired attitude rates in order to steer the vehicle during the mission task. Results of both numerical simulations and, for the first time, an experimental validation are provided in order to assess the performance of the approach in the presence of external disturbances and unmodeled dynamics, provided adequate time horizon and update frequency are selected for the numerical optimization algorithm.〈/p〉〈/div〉

Description: 〈p〉Publication date: April 2019〈/p〉 〈p〉〈b〉Source:〈/b〉 Aerospace Science and Technology, Volume 87〈/p〉 〈p〉Author(s): Qing Wang, Qijun Zhao〈/p〉 〈h5〉Abstract〈/h5〉 〈div〉〈p〉In order to improve the effective load of an unmanned helicopter, a new blade shape with a new airfoil section, non-linear negative twist, varied chord length and double-swept blade tip was designed by employing a CFD method coupled with an optimization method. Firstly, a new airfoil is designed with improved aerodynamic characteristics. By comparing the numerical data, the new airfoil has better lift-drag ratio, i.e., the maximum lift-drag ratio increases from 63.03 to 66.29 at Mach number of 0.3 and from 70.17 to 72.13 at Mach number of 0.4, compared with the original NACA8H12 airfoil. In addition, a new blade shape is designed based on the new airfoil. By comparing the numerical data, the 〈math xmlns:mml="http://www.w3.org/1998/Math/MathML" altimg="si1.gif" overflow="scroll"〉〈msub〉〈mrow〉〈mi〉C〈/mi〉〈/mrow〉〈mrow〉〈mi〉T〈/mi〉〈/mrow〉〈/msub〉〈/math〉 of the design rotor increases from 〈math xmlns:mml="http://www.w3.org/1998/Math/MathML" altimg="si2.gif" overflow="scroll"〉〈mn〉4.79〈/mn〉〈mo lspace="0em" rspace="0em"〉×〈/mo〉〈msup〉〈mrow〉〈mn〉10〈/mn〉〈/mrow〉〈mrow〉〈mo〉−〈/mo〉〈mn〉3〈/mn〉〈/mrow〉〈/msup〉〈/math〉 to 〈math xmlns:mml="http://www.w3.org/1998/Math/MathML" altimg="si3.gif" overflow="scroll"〉〈mn〉5.07〈/mn〉〈mo lspace="0em" rspace="0em"〉×〈/mo〉〈msup〉〈mrow〉〈mn〉10〈/mn〉〈/mrow〉〈mrow〉〈mo〉−〈/mo〉〈mn〉3〈/mn〉〈/mrow〉〈/msup〉〈/math〉, and the maximum FM increases from 0.67 to 0.72 at the design state. After that, a verification test is performed in hovering flight. The test data indicated the a maxumu thrust increase of 3.18% at the design state. Meanwhile, the FM of design rotor is improved about 3.41% compared with the original rotor.〈/p〉〈/div〉

Description: 〈p〉Publication date: April 2019〈/p〉 〈p〉〈b〉Source:〈/b〉 Aerospace Science and Technology, Volume 87〈/p〉 〈p〉Author(s): Zhenbo Wang〈/p〉 〈h5〉Abstract〈/h5〉 〈div〉〈p〉Hypersonic trajectory optimization has been intensively investigated through different approaches; however, the normal-load-optimal entry problems were barely studied and reported in the literature. Finding the optimal trajectories with maximum or minimum peak normal load is essential to evaluate the maneuverability and structural strength of the vehicle. In this paper, both the maximum and minimum peak-normal-load entry trajectories are explored using convex optimization. Based on the previous work, the maximum-peak-normal-load entry problem is firstly addressed by a Big-M method and a line-search approach. Through successive relaxations, the nonconvex discrete-event optimal control problem associated with maximum-peak-normal-load entry is transformed into a sequence of mixed-integer convex optimization problems. Then, a line-search technique is introduced to improve the convergence of the proposed method. Additionally, a sequential convex programming method is designed to solve the minimum-peak-normal-load entry problem to comprehensively analyze the normal load during the entry flight. There are efficient solvers that can solve each relaxed convex subproblem with a global optimum if the feasible set of the subproblem is nonempty. The convergence and accuracy of the proposed methodologies are demonstrated by numerical simulations, and the feasibility of the converged solutions is discussed based on an entry-corridor approach.〈/p〉〈/div〉

Description: 〈p〉Publication date: Available online 12 March 2019〈/p〉 〈p〉〈b〉Source:〈/b〉 Aerospace Science and Technology〈/p〉 〈p〉Author(s): Peng Jun, Zhang Zijian, Hu Zongmin, Jiang Zonglin〈/p〉 〈h5〉Abstract〈/h5〉 〈div〉〈p〉In this paper, we study the high-temperature effects on the reflection of shock waves in hypersonic flows by using analytical and computational approaches. First, a theoretical approach is established to solve the shock relations which are further applied to develop the shock polar analytical method for high-temperature air. Then, a comparative investigation using ideal gas model and real gas model considering vibration excitation indicates that the high-temperature effects cause an obvious change to the overall profile of the shock polar. The post-shock pressure increases within the strong branch of the shock polar while decreases within the weak branch due to vibration excitation of air molecules. A more notable phenomenon is the increase in the maximum deflection angle of the shock polar which can significantly influence the detachment criterion of shock reflection transition in high-temperature air flows. The shock polar analysis of shock reflection shows that the high-temperature effects result in an obvious increase to the detachment criterion while a slight increase to the von Neumann criterion. A series of computations are conducted to confirm the above analytical findings on the shock reflection considering high-temperature effects. A slight difference of transition criterion between the theory and computations is found to be caused by the existence of the expansion fan which is an inherent flow structure. The proposed shock polar analytical method is proved to be an effective but simple approach for the study of shock wave reflections in hypersonic flows.〈/p〉〈/div〉

Description: 〈p〉Publication date: May 2019〈/p〉 〈p〉〈b〉Source:〈/b〉 Aerospace Science and Technology, Volume 88〈/p〉 〈p〉Author(s): Pengfei Fu, Lingyun Hou, Zhuyin Ren, Zhen Zhang, Xiaofang Mao, Yusong Yu〈/p〉 〈h5〉Abstract〈/h5〉 〈div〉〈p〉Droplet impact on a solid wall contributes to droplet vaporization and cooling on the inner wall. A new droplet/wall impact model is developed and fitted from the experimental data, which takes both kinematic parameters of the impinging droplets and thermal parameters of the solid wall into consideration. The model describes the behavior of droplets after impact and identifies five representative regimes, i.e., stick/spread, suspend, rebound, boiling including breakup, and splash, through the critical Weber number of droplets and wall temperature. To assess the new model, numerical simulations were performed of propellant droplets impacting the wall of a bipropellant rocket engine chamber. A comparison with experimental data shows better predictions of the wall temperature of the chamber than obtained from previous models. A larger number of large-sized monomethylhydrazine and nitrogen tetroxide droplets in the boiling induced breakup and splash regimes break into smaller droplets when using the new model. There are also temperature peaks near the impact points and near the throat. Especially in the throat of the combustion chamber, the new model predicts the wall temperature distribution accurately, offering improved prediction of the combustion chamber and assessing thermal protection of the throat.〈/p〉〈/div〉

Description: 〈p〉Publication date: May 2019〈/p〉 〈p〉〈b〉Source:〈/b〉 Aerospace Science and Technology, Volume 88〈/p〉 〈p〉Author(s): Zhankui Song, Kaibiao Sun〈/p〉 〈div xml:lang="en"〉 〈h5〉Abstract〈/h5〉 〈div〉〈p〉This paper investigates the trajectory tracking problem for a small coaxial-rotor unmanned aerial vehicle (CRUAV) with partial loss of actuator effectiveness. The CRUAV model is decomposed into a dual loop structure based on back-stepping design idea. First, certain performance function specified a priori by the designer is introduced into the position loop such that the original position tracking error is transformed into an equivalent constrained variable providing for the performance judgment of the position loop. Then, a fault-tolerant control scheme is proposed based on adaptive strategies for compensating the effect caused by various adverse factors. Subsequently, an attitude control loop is derived by incorporating adaptive compensation method. It is proved that the proposed dual-loop structure is able to guarantee the satisfaction of the pre-specified constraint on the transformed errors. Finally, the effectiveness and benefits of the design dual-loop control system are validated via computer simulation.〈/p〉〈/div〉 〈/div〉

Description: 〈p〉Publication date: May 2019〈/p〉 〈p〉〈b〉Source:〈/b〉 Aerospace Science and Technology, Volume 88〈/p〉 〈p〉Author(s): Hyeonsoo Yeo〈/p〉 〈div xml:lang="en"〉 〈h5〉Abstract〈/h5〉 〈div〉〈p〉This paper reviews lessons learned from the compound helicopter studies performed by NASA and the US Army Aviation Development Directorate to support the NASA Heavy Lift Rotorcraft Systems Investigation and the US Army's Joint Heavy Lift (JHL), Joint Multi-Role Technology Demonstrator (JMR-TD), and Future Vertical Lift (FVL) programs. These studies explored performance potential of advanced rotorcraft and investigated the impact of key modern-technologies in performance, weight, and aerodynamics on rotorcraft. The compound helicopter configurations considered in this paper represent a wide range of sizes, gross weight, rotor systems, and operating conditions. A brief description of design and aeromechanics analysis tools and methodologies is provided. Rotor performance correlation results at high advance ratio, which are critical for the accurate design and analysis of high-speed rotorcraft, are shown. Detailed aeromechanics analysis results, such as the effects of various compounding methods, lift share between rotor and wing, rotor rotational speed, blade twist, aircraft drag, and rotor/wing interference on aircraft performance, are presented.〈/p〉〈/div〉 〈/div〉

Description: 〈p〉Publication date: May 2019〈/p〉 〈p〉〈b〉Source:〈/b〉 Aerospace Science and Technology, Volume 88〈/p〉 〈p〉Author(s): Aaron D. Koch〈/p〉 〈h5〉Abstract〈/h5〉 〈div〉〈p〉Currently, algorithms exist for the optimal staging of serially staged rockets under loss-free conditions. These algorithms are based on the Tsiolkovsky rocket equation. Here, one variant is extended to include velocity losses and/or fairing separation. Instead of simply adding the velocity losses to the required loss-free Δ〈em〉v〈/em〉 and freely distributing the total amount among all stages, a two-step process is implemented. First, the loss-free solution is obtained to determine the optimal velocity gain for each stage. Then, the rocket's stages are iteratively scaled up, starting with the uppermost stage and continuing downwards. The size of each stage is enlarged so that it generates Δ〈em〉v〈/em〉 equal to the optimal velocity gain plus the losses occurring during its flight. Also, each stage accounts for the increased mass of the stages on top of it. Adding the velocity losses after the optimization step ensures that they are allocated to the correct stages. Ariane 40 is used as an example for a three-stage rocket. In this case, the proposed method produced realistic payload ratios, in contrast to the old idea of adding the velocity losses directly to the required loss-free Δ〈em〉v〈/em〉. As the method requires inputs that stem from a trajectory analysis, it works best when iteratively coupled to a trajectory optimizer. In doing so, Ariane 40's total payload ratio was increased, while taking both velocity losses and fairing separation into account.〈/p〉〈/div〉

Description: 〈p〉Publication date: May 2019〈/p〉 〈p〉〈b〉Source:〈/b〉 Aerospace Science and Technology, Volume 88〈/p〉 〈p〉Author(s): Robert Dibble, Vaclav Ondra, Branislav Titurus〈/p〉 〈div xml:lang="en"〉 〈h5〉Abstract〈/h5〉 〈div〉〈p〉Varying the rotational speed of the main rotor is one method being considered to improve the performance of future rotorcraft. However, changes in rotor speeds often lead to resonant interactions between rotor blade modes and the rotor's excitation frequencies which increase the vibratory loads in the rotor. This research investigates the use of a compressive load to reduce a blade's natural frequencies and its potential to be used as a resonance avoidance technique by improving separation between the natural and excitation frequencies of a blade. The research presented herein describes and validates a model of a pretwisted rotating beam with non-coincident mass and elastic axes with an applied compressive load. The compressive load is applied at the elastic axis at the tip of the beam and is orientated towards the root of the beam. The beam model is then used in a case study to represent the rotor blade of a typical mid-sized civilian helicopter. The case study is performed to calculate the natural frequencies of a compressed blade for a reduction in rotor speed of up to 40% and evaluate the performance of the compressive load resonance avoidance technique. The results of the case study show that the compressive load improves the separation between natural and excitation frequencies over the full range of rotor speeds evaluated. The improved separation allows the rotor to operate safely with a reduction in rotor speed of up to 19%.〈/p〉〈/div〉 〈/div〉

Description: 〈p〉Publication date: May 2019〈/p〉 〈p〉〈b〉Source:〈/b〉 Aerospace Science and Technology, Volume 88〈/p〉 〈p〉Author(s): Lei Sun, Yong Huang, Ruixiang Wang, Xiang Feng, Zhilin Liu, Jiaming Wu〈/p〉 〈h5〉Abstract〈/h5〉 〈div〉〈p〉The lean blowout (LBO) limit is a crucial performance for aircraft engine combustors. It is essential to obtain the LBO limit during the design stage of the aircraft engine combustors. The semi-empirical correlation is an important tool for quick prediction of the LBO limits. Among all the semi-empirical correlations for the prediction of the LBO limits, Lefebvre's LBO model is widely used for the swirl stabilized combustors. The Flame Volume (FV) model was proposed based on Lefebvre's LBO model to accommodate the effects of the geometry of the flame tube on the LBO. Meanwhile, the multi-point lean direct injection (MPLDI) combustor whose geometry of the dome is different from the traditional combustors is a promising low NOx emission combustor. Up to now, there are few existing semi-empirical correlations to predict the LBO limit for the MPLDI combustors although the prediction of the LBO limit is critical for them. Based on the FV concept and new physics-based analysis, the FV-MP (Flame Volume for the Multi-Point) model is derived to predict the LBO limits for the MPLDI combustors. The FV-MP model could accommodate the effects of the fuel staging and recessed pilot stage, in addition to the operating conditions, on the LBO limits of the MPLDI combustors and achieve better prediction accuracy than both the FV and Lefebvre's LBO models within the range of corresponding validation experiments. Compared with Lefebvre's LBO model, the FV model could double the prediction accuracy. Compared with the FV model, the FV-MP model could further double the prediction accuracy.〈/p〉〈/div〉

Description: 〈p〉Publication date: May 2019〈/p〉 〈p〉〈b〉Source:〈/b〉 Aerospace Science and Technology, Volume 88〈/p〉 〈p〉Author(s): Pu Zhang, Jinglei Xu, Yang Yu, Wei Cui〈/p〉 〈h5〉Abstract〈/h5〉 〈div〉〈p〉Influences of adverse pressure gradient (APG) on supersonic turbulent boundary layers are numerically studied using Reynolds-averaged Navier–Stokes (RANS) equations. The RANS methodology is validated by comparing the numerical results with the existing experimental data. Although the flame is restricted in the boundary layer, the heat flux is reduced rather than increased for the suppressed turbulence momentum transportation ability. Moreover, APG contributes more to the heat flux than combustion heat release does. Compared with no-injection case, a large skin friction reduction can be obtained by boundary layer combustion, and further reduction can be achieved in APG state. Additionally, the effects of combustion in APG on the velocity laws of the wall show that White's velocity law is close to the numerical results in the outer region of boundary layer, and Nichols's velocity law is appropriate in the whole boundary layer unless the APG is too strong. Turbulence intensity influences are analyzed in the end. Results show that the additional reduction of skin friction due to induced combustion cannot offset the skin friction increase caused by high turbulence intensity.〈/p〉〈/div〉

Description: 〈p〉Publication date: April 2019〈/p〉 〈p〉〈b〉Source:〈/b〉 Aerospace Science and Technology, Volume 87〈/p〉 〈p〉Author(s): Fei Sun, Kamran Turkoglu〈/p〉 〈h5〉Abstract〈/h5〉 〈div〉〈p〉In this paper, based on real-time nonlinear receding horizon control methodology, a novel scheme is developed for multi-agent nonlinear consensus problem under jointly connected switching topologies. The consensus problem is converted into a family of finite horizon optimization control process and is solved numerically to generate distributed control protocols in real-time. The stability is proved without the assumption that the topology is connected for all the time. Two benchmark examples on nonlinear chaotic systems provide validated results which demonstrate the significant outcomes of such methodology.〈/p〉〈/div〉

Description: 〈p〉Publication date: May 2019〈/p〉 〈p〉〈b〉Source:〈/b〉 Aerospace Science and Technology, Volume 88〈/p〉 〈p〉Author(s): X. Mao, P.N.A.M. Visser, J. van den IJssel〈/p〉 〈h5〉Abstract〈/h5〉 〈div〉 〈p〉Baseline determination for the European Space Agency Swarm magnetic field mission is investigated. Swarm consists of three identical satellites -A, -B and -C. The Swarm-A and -C form a pendulum formation whose baseline length varies between about 30 and 180 km. Swarm-B flies in a higher orbit, causing its orbital plane to slowly rotate with respect to those of Swarm-A and -C. This special geometry results in short periods when the Swarm-B satellite is adjacent to the other Swarm satellites. Ten 24-hr periods around such close encounters have been selected, with baseline lengths varying between 50 and 3500 km. All Swarm satellites carry high-quality, dual-frequency and identical Global Positioning System receivers not only allowing precise orbit determination of the single Swarm satellites, but also allowing a rigorous assessment of the capability of precise baseline determination between the three satellites. These baselines include the high-dynamic baselines between Swarm-B and the other two Swarm satellites.〈/p〉 〈p〉For all orbit determinations, use was made of an Iterative Extended Kalman Filter approach, which could run in single-, dual-, and triple-satellite mode. Results showed that resolving the issue of half-cycle carrier phase ambiguities (present in original release of GPS RINEX data) and reducing the code observation noise by the German Space Operations Center converter improved the consistency of reduced-dynamic and kinematic baseline solutions for both the Swarm-A/C pendulum pair and other combinations of Swarm satellites. All modes led to comparable consistencies between the computed orbit solutions and satellite laser ranging observations at a level of 2 cm. In addition, the consistencies with single-satellite ambiguity fixed orbit solutions by the German Space Operations Center are at comparable levels for all the modes, with reduced-dynamic baseline consistency at a level of 1-3 mm for the pendulum Swarm-A/C formation and 3-5 mm for the high-dynamic Swarm-B/A and -B/C satellite pairs in different directions.〈/p〉 〈/div〉

Description: 〈p〉Publication date: Available online 18 March 2019〈/p〉 〈p〉〈b〉Source:〈/b〉 Aerospace Science and Technology〈/p〉 〈p〉Author(s): Anmin Zhao, Hui Zou, Haichuan Jin, Dongsheng Wen〈/p〉 〈div xml:lang="en"〉 〈h5〉Abstract〈/h5〉 〈div〉〈p〉A whole adaptive variable camber wing (AVCW) equipped with an innovative double rib sheet (DRS) structure is experimentally and numerically studied in this work. The DRS structure adopts the surface contact mode for the force transmission of changeable camber wing instead of the conventional rigid hinge joint contact. The whole AVCW design allows to adjust the shape of wing in a real-time at various flight conditions, which is of great interest for Unmanned Aerial Vehicle (UAV) applications. The flight-test experiments demonstrate that the total AVCW carrying the developed adaptive control system (ACS) can enhance UAV flight efficiency by 14.1% comparing to a traditional fixed-wing of Talon UAV. In addition, it indicates that employing the whole AVCW structure can sustain a larger flight load, without increasing the weight of entire wing structure except for the actuator device and adhesives, which is promising for future engineering applications.〈/p〉〈/div〉 〈/div〉

Description: 〈p〉Publication date: Available online 15 March 2019〈/p〉 〈p〉〈b〉Source:〈/b〉 Aerospace Science and Technology〈/p〉 〈p〉Author(s): Xiaomeng Yin, Bo Wang, Lei Liu, Yongji Wang〈/p〉 〈h5〉Abstract〈/h5〉 〈div〉〈p〉In this study, hypersonic vehicle (HV) tracking control in the presence of uncertainties and external disturbance is investigated. As the exact bounds of uncertainties and disturbances are usually unknown during flight, a disturbance observer (DOB)-based gain adaptation high-order sliding mode control (HOSMC) method is proposed for HVs. To mitigate the chattering effect while maintaining strong robustness, a method combining the advantages of the DOB and gain adaptation is introduced into the HOSMC. The DOB is employed to estimate and reject the uncertainties and external disturbance. Additionally, an adaptive control law is developed to compensate for estimation errors. By combining the DOB and the adaptive HOSMC, the unnecessarily large gain for maintaining robustness is reduced; thus, the chattering is mitigated. The effectiveness of the proposed control method is validated via simulation, in which a strong robustness, high tracking performance, and reduced chattering effect are achieved under uncertainties and external disturbance.〈/p〉〈/div〉

Description: 〈p〉Publication date: May 2019〈/p〉 〈p〉〈b〉Source:〈/b〉 Aerospace Science and Technology, Volume 88〈/p〉 〈p〉Author(s): Davide Cinquegrana, Raffaele Votta, Carlo Purpura, Eduardo Trifoni〈/p〉 〈h5〉Abstract〈/h5〉 〈div〉〈p〉A facility characterization test campaign was performed for NASA in SCIROCCO Plasma Wind Tunnel, with increasing enthalpy and constant reservoir pressure. In the frame of numerical rebuilding of the experimental data, an important gap among measured and numerically predicted values of heat flux led to deep analyse the numerical methods and the models usually employed for test rebuilding. The Navier–Stokes model with chemical reacting non-equilibrium flows, denoted a lack of physical accuracy due to local rarefaction effects, as certified by means of the continuum breakdown parameter. Furthermore, the low stagnation pressure environment could influence the surface catalytic behaviour of the hemisphere copper probe, and the chemical contribution to the stagnation heat flux. At the end, a set of direct simulation Monte Carlo with partial catalytic behaviour of the probe was performed, in order to address both critical phenomena highlighted and close the gap with the measured heat fluxes, understanding the actual test chamber environment.〈/p〉〈/div〉

Description: 〈p〉Publication date: May 2019〈/p〉 〈p〉〈b〉Source:〈/b〉 Aerospace Science and Technology, Volume 88〈/p〉 〈p〉Author(s): Giuseppe Petrone, Giacomo Melillo, Aurelio Laudiero, Sergio De Rosa〈/p〉 〈h5〉Abstract〈/h5〉 〈div〉 〈p〉Comfort plays an increasingly important role in the interior design of airplanes. In general, comfort is defined as ‘freedom from pain, well-being’; in scientific literature, indeed, it is defined as a pleasant state of physiological, psychological and physical harmony between a human being and the environment or a sense of subjective well-being. Cabin noise in passenger aircraft is one of the comfort parameter, which creates straightaway discomfort when exceeding personal thresholds. In general the cabin noise varies by the seat position and changes with flight condition. It is driven by several source types, which are transmitted through different transfer paths into the cabin. In the forward area the noise is mainly dominated by the turbulent boundary layer described by pressure vortexes traveling along the fuselage surface.〈/p〉 〈p〉In this paper evaluation of the Sound Pressure level, for the medium-high frequency range, of an aircraft fuselage section at different stations and locations inside the cabin has been performed numerically by using Statistical Energy Analysis (SEA) method. Different configurations have been considered for the analysis: from the “naked” cabin (only primary structure) up to “fully furnished” (primary structure with interiors and noise control treatments) one. These results are essential to understand which are the main parameters affecting the noise insulation. Furthermore the Power Inputs evaluation has been determined to see the contribution of each considered aeronautic component on the acoustic insulation. Finally, the effect of a viscoelastic damping layer embedded in the glass window has been evaluated.〈/p〉 〈/div〉

Description: 〈p〉Publication date: May 2019〈/p〉 〈p〉〈b〉Source:〈/b〉 Aerospace Science and Technology, Volume 88〈/p〉 〈p〉Author(s): Devi Supriya, K.V. Nagaraja, T.V. Smitha, Sarada Jayan〈/p〉 〈h5〉Abstract〈/h5〉 〈div〉〈p〉This paper presents automatically generated higher order curved triangular meshes around airfoil design using MATLAB code. This work shows a valuable basis for the finite element procedures involved in evaluating aerodynamic performances. Finite element method (FEM) effectively solves all computational fluid dynamics problems around the airfoil and for that region around the airfoil that has been discretized with unstructured curved triangular elements. Meshes have been formed on the basis of subparametric transformation created for the curved triangular element obtained from the nodal relations of parabolic arcs. This scheme can be used to obtain the output data of node coordinates, element connectivity and boundary values for all discretized elements over the airfoil design. A spectacular work done on linear triangular element meshing over a domain by Persson and Gilbert Strang is the basis of present meshing scheme. The proposed meshing scheme presents a refined higher order (HO) curved triangular discretization of few airfoil designs namely NACA0012, NACA0015 and NACA0021 inscribed inside a circle. The approach of the suggested meshing scheme described in this paper can be applied to numerous aerospace applications such as computing pressure gradients, understanding atmospheric nature study, evaluating laminar viscous compressible flow around the airfoil shape, etc. The element and nodal information gained from this discretization is useful for the numerical solutions of FEM and for the aerodynamic portrayal. This paper is aimed at the innovative discretization scheme which can be extended to all kinds of NACA airfoil designs. We have provided the MATLAB code 〈em〉AirfoilHOmesh2d〈/em〉 for HO curved meshing around an airfoil with a cubic order triangular element. The mathematical explanation of this along with the description and implementation of it on few airfoil designs is described. The flowchart of the MATLAB code for cubic order meshing over airfoil design has been provided. This implementation supports many applications in an aerodynamic performance that have been elaborated in this paper. Two applications for the analysis of potential flow around airfoil and computation of pressure coefficient (〈math xmlns:mml="http://www.w3.org/1998/Math/MathML" altimg="si1.gif" overflow="scroll"〉〈msub〉〈mrow〉〈mi〉C〈/mi〉〈/mrow〉〈mrow〉〈mi〉p〈/mi〉〈/mrow〉〈/msub〉〈/math〉) on the surface of the airfoil design have been performed. It has been verified and found that the present HO curved meshing technique efficiently gives converging solution.〈/p〉〈/div〉

Description: 〈p〉Publication date: May 2019〈/p〉 〈p〉〈b〉Source:〈/b〉 Aerospace Science and Technology, Volume 88〈/p〉 〈p〉Author(s): Natsuki Tsushima, Tomohiro Yokozeki, Weihua Su, Hitoshi Arizono〈/p〉 〈div xml:lang="en"〉 〈h5〉Abstract〈/h5〉 〈div〉〈p〉In this paper, an integrated geometrically nonlinear aeroelastic framework to analyze the static nonlinear aeroelastic response of morphing composite wing with orthotropic materials has been developed. A flat plate/shell finite element, which can model plate-like wings, has been accommodated to model composite/corrugated panels to investigate effects of different laminate orientations and corrugations. A corotational approach is used to consider the geometrical nonlinearity due to large deformation produced by wing morphing. An unsteady vortex-lattice method is implemented to couple with the structural model subject to the large deformations. A homogenization method is also implemented to model corrugated panels as equivalent orthotropic plates. Individual structural, aerodynamic, and corrugated panel models, as well as the complete nonlinear aeroelastic framework, are verified. Numerical studies explore the static aeroelastic responses of a flat wing with composite/corrugated panels. This work helps to understand the nonlinear aeroelastic characteristics of composite/corrugated wings and demonstrates the capability of the framework to analyze the nonlinear aeroelasticity of such morphing wings.〈/p〉〈/div〉 〈/div〉

Description: 〈p〉Publication date: May 2019〈/p〉 〈p〉〈b〉Source:〈/b〉 Aerospace Science and Technology, Volume 88〈/p〉 〈p〉Author(s): Mingying Huo, Giovanni Mengali, Alessandro A. Quarta, Naiming Qi〈/p〉 〈h5〉Abstract〈/h5〉 〈div〉〈p〉The aim of this paper is to propose a shape-based method in which the concept of Bezier curve is used to efficiently design the three-dimensional interplanetary trajectory of a spacecraft whose primary propulsion system is an Electric Solar Wind Sail. The latter is a propellantless propulsion concept that consists of a spinning grid of tethers, kept at a high positive potential by a power source and maintained stretched by the centrifugal force. The proposed approach approximates the time variation of the components of the spacecraft position vector using a Bezier curve function, whose geometric coefficients are calculated by optimizing the total flight time with standard numerical methods and enforcing the boundary conditions of a typical interplanetary rendezvous mission. The paper also discusses a geometrical approach to include, in the optimization process, the propulsive acceleration vector constraints obtained with the latest Electric Solar Wind Sail thrust model.〈/p〉〈/div〉

Description: 〈p〉Publication date: May 2019〈/p〉 〈p〉〈b〉Source:〈/b〉 Aerospace Science and Technology, Volume 88〈/p〉 〈p〉Author(s): Steven F.T. Apirana, Christopher M. James, Robert Eldridge, Richard G. Morgan, Steven W. Lewis〈/p〉 〈h5〉Abstract〈/h5〉 〈div〉〈p〉The X2 expansion tube facility at The University of Queensland is capable of simulating entry into most of the planetary bodies in our solar system, producing test conditions with stagnation enthalpies in excess of 100 MJ/kg. Models used in X2 are typically made from steel and, consequently, manufacture is often constrained to conventional methods, with associated long lead times, and high cost. Through an experimental campaign, the survivability and applicability of Rapid Prototype models was investigated. Three prototyping methods were investigated; Selective Laser Sintering, Stereolithography, and Fused Deposition Modelling, and these were compared to a steel baseline. A computational stress analysis was used to design an internally hollow test model geometry. All models survived the experimental test-time, one was destroyed by the post-experiment flow. It was thought that test-models might ablate during the experimental test time and this was investigated by using filtered imaging to capture Cyanogen radiation occurring in the model's boundary layer. Ablation was seen in all Rapid Prototype models, most strongly observed in the Selective Laser Sintered and Acrilonitrile Butadiene Styrene models. These models may be suitable for the study of non-equilibrium radiative emission just behind the shock wave, or ablation phenomena in a model's boundary layer.〈/p〉〈/div〉

Description: 〈p〉Publication date: May 2019〈/p〉 〈p〉〈b〉Source:〈/b〉 Aerospace Science and Technology, Volume 88〈/p〉 〈p〉Author(s): Feng Liu, Feng Gao, Weiwei Zhang, Bai Zhang, Jianhua He〈/p〉 〈h5〉Abstract〈/h5〉 〈div〉〈p〉The task of Integrated Power and Attitude Control (IPAC) of a spacecraft can be implemented by Variable Speed Control Moment Gyros (VSCMGs). The Integrated Power and Attitude Control System (IPACS) singularity problem is the key factor for the spacecraft to successfully perform IPAC task, which can be overcome by rational designing steering law. The singularity characteristic and the steering law should be considered during the process of design parameters for the VSCMG cluster with IPACS task. There is no research report in this area at present. The steering results of weighted pseudo-inverse and null motion of Weighted Pseudo-Inverse with Null Motion (WPINM) can be canceled by each other under some certain condition. So the flywheel torque requirement of the WPINM steering law can be greatly increased, which is contradictory to the original design intention of the weighted matrix. A steering law with minimum requirement of flywheel power and torque is introduced from existing research result. Then, the constraint of the IPACS singularity problem and the SGCMGs singularity problem cannot be encountered during the whole process of IPAC task is given. At last, the parameters design problem of VSCMGs for the IPAC task is cast as a multi-objective optimization problem with minimum whole system power and maximum utilization ratio of flywheel momentum under the condition of consideration of the steering. The intelligent algorithm of Non-dominated Sorting Genetic Algorithm (NSGA) is used to solve the nonlinear multi-objective problem. The flywheel power can be greatly reduced by the new parameters design method.〈/p〉〈/div〉

Description: 〈p〉Publication date: May 2019〈/p〉 〈p〉〈b〉Source:〈/b〉 Aerospace Science and Technology, Volume 88〈/p〉 〈p〉Author(s): Rodrigo C. Palharini, Wilson F.N. Santos〈/p〉 〈h5〉Abstract〈/h5〉 〈div〉〈p〉A computational investigation has been carried out to examine a non-reacting rarefied hypersonic flow over cavities by employing the Direct Simulation Monte Carlo (DSMC) method. The work focuses on the effects on the aerodynamic surface quantities due to variations in the cavity length-to-depth (〈math xmlns:mml="http://www.w3.org/1998/Math/MathML" altimg="si1.gif" overflow="scroll"〉〈mi〉L〈/mi〉〈mo stretchy="false"〉/〈/mo〉〈mi〉H〈/mi〉〈/math〉) ratio. The results highlight the sensitivity of the heat transfer, pressure and skin friction coefficients due to changes to the cavity 〈math xmlns:mml="http://www.w3.org/1998/Math/MathML" altimg="si1.gif" overflow="scroll"〉〈mi〉L〈/mi〉〈mo stretchy="false"〉/〈/mo〉〈mi〉H〈/mi〉〈/math〉 ratio. The 〈math xmlns:mml="http://www.w3.org/1998/Math/MathML" altimg="si1.gif" overflow="scroll"〉〈mi〉L〈/mi〉〈mo stretchy="false"〉/〈/mo〉〈mi〉H〈/mi〉〈/math〉 ratio ranged from 1 to 4, which corresponds to the transition flow regime based on an overall Knudsen number 〈math xmlns:mml="http://www.w3.org/1998/Math/MathML" altimg="si2.gif" overflow="scroll"〉〈mi〉K〈/mi〉〈msub〉〈mrow〉〈mi〉n〈/mi〉〈/mrow〉〈mrow〉〈mi〉L〈/mi〉〈/mrow〉〈/msub〉〈/math〉. The analysis showed that the aerodynamic quantities acting on the cavity surface rely on the 〈math xmlns:mml="http://www.w3.org/1998/Math/MathML" altimg="si1.gif" overflow="scroll"〉〈mi〉L〈/mi〉〈mo stretchy="false"〉/〈/mo〉〈mi〉H〈/mi〉〈/math〉 ratio. It was found that pressure load and heating load to the cavity surfaces presented peak values along the forward face, more precisely in the vicinity of the cavity shoulder. Moreover, these loads are much higher than those found in a smooth surface, for the conditions investigated.〈/p〉〈/div〉

Description: 〈p〉Publication date: May 2019〈/p〉 〈p〉〈b〉Source:〈/b〉 Aerospace Science and Technology, Volume 88〈/p〉 〈p〉Author(s): Chen Xu, Yijun Mao, Zhiwei Hu〈/p〉 〈h5〉Abstract〈/h5〉 〈div〉〈p〉Various passive and active methods have been developed to control flow separation from bluff bodies. However, these methods require adjusting features of the solid surface, such as modifying its geometry or porosity, or applying external force or momentum. This paper develops an off-body-based method, without adjusting any features of the solid surface, for controlling the unsteady flow separation by fixing porous materials downstream of bluff bodies. Numerical study on flow past a circular cylinder at a subcritical Reynolds number is performed, and the result indicates that the added downstream porous material changes flow in the wake and re-laminarizes the turbulent flow around the curved cylinder surface, reducing the wall pressure fluctuation around the cylinder. Therefore, the associated aerodynamic noise is reduced greatly.〈/p〉〈/div〉

Description: 〈p〉Publication date: Available online 26 June 2019〈/p〉 〈p〉〈b〉Source:〈/b〉 Aerospace Science and Technology〈/p〉 〈p〉Author(s): Ziyang Zhen, Gang Tao, Chaojun Yu, Yixuan Xue〈/p〉 〈h5〉Abstract〈/h5〉 〈div〉〈p〉This paper studies a multivariable model reference adaptive control (MRAC) scheme for the automatic carrier-landing control problem of unmanned aerial vehicles (UAVs) with system dynamics of nonlinearity, multivariable coupling and parametric uncertainty. A complete automatic carrier landing system (ACLS) for carrier-based UAVs is developed, which consists of a guidance subsystem and a flight control subsystem. The MRAC scheme is based on a state feedback for output tracking framework with relaxed design conditions, which guarantees the reference glide slope tracking. Simulation results of a nonlinear UAV model demonstrate that the multivariable MRAC based ACLS has a better carrier-landing performance than a fixed control based ACLS.〈/p〉〈/div〉

Description: 〈p〉Publication date: September 2019〈/p〉 〈p〉〈b〉Source:〈/b〉 Aerospace Science and Technology, Volume 92〈/p〉 〈p〉Author(s): A. Alaimo, C. Orlando, S. Valvano〈/p〉 〈h5〉Abstract〈/h5〉 〈div〉〈p〉In this work analytical damped free-vibration and frequency response solutions are obtained for the analysis of composite plates structures embedding viscoelastic layers. On the basis of the Principle of Virtual Displacements, Layer-Wise models related to linear up to fourth order variations of the unknown variables in the thickness direction are treated. Analytical solutions using the Navier procedure are presented for the analysis of isotropic, cross-ply composite and simply-supported plate structures. The modelization of multilayered structure materials takes into account the composite material properties and the frequency dependence of the viscoelastic material. Various external loads are considered: closed form solution for bi-sinusoidal pressure, constant distributed pressure and concentrated loads. Several analyses are carried out to validate and demonstrate the accuracy and efficiency of the present formulation for the study of viscoelastic plates, taking into account different lamination sequences and different plate aspect ratios.〈/p〉〈/div〉

Description: 〈p〉Publication date: June 2019〈/p〉 〈p〉〈b〉Source:〈/b〉 Aerospace Science and Technology, Volume 89〈/p〉 〈p〉Author(s): Xingling Shao, Linwei Wang, Jie Li, Jun Liu〈/p〉 〈div xml:lang="en"〉 〈h5〉Abstract〈/h5〉 〈div〉〈p〉This paper investigates the trajectory tracking and attitude stabilization problem with only position measurements for quadrotors subject to position constraints and uncertainties. By introducing the one-to-one nonlinear mapping (NM) to prevent position state violation, an output constrained trajectory tracking law is developed by transforming the original restricted translational dynamics into an equivalent unconstrained subsystem. To address the uncertainties arising from parametric deviations and external disturbances, with given model information incorporated into observer design, we develop a high-order extended state observer (ESO), capable of simultaneously online estimating the uncertainties and full-states of quadrotors. Then, an output feedback based trajectory tracking and attitude stabilization approach is synthesized by integrating NM and high-order ESO via dynamic surface control (DSC), leading to a much simpler control structure and reduced implementation costs. The salient feature is that position constraints, uncertainties as well as output feedback difficulties can be comprehensively handled with acceptable control performance. It is shown via Lyapunov stability that all signals in the closed-loop system are guaranteed to be uniformly ultimately bounded. Simulation results are provided to validate the benefits and effectiveness of the proposed method.〈/p〉〈/div〉 〈/div〉

Description: 〈p〉Publication date: Available online 9 April 2019〈/p〉 〈p〉〈b〉Source:〈/b〉 Aerospace Science and Technology〈/p〉 〈p〉Author(s): Zongxia Jiao, Bo Yu, Shuai Wu, Yaoxing Shang, Haishan Huang, Zhewen Tang, Renlei Wei, Chunfang Li〈/p〉 〈h5〉Abstract〈/h5〉 〈div〉〈p〉The design of flight control actuation system is facing major challenge due to the development of more electrical aircraft. The task is to find the combinations of power sources, actuators and computers, which becomes more complex because of the new power sources and actuator types of more electrical aircraft. It is impossible to determine optimal architecture by traditional trial-and-error method within acceptable time. Therefore, the need for new methodology for actuation system architecture design emerges. This study proposes an intelligent design method which has steps of design space exploration of actuation system architectures by constraint satisfaction problem (CSP) method, safety assessment process to exclude unsafety solution, multi-objectives optimization to get Pareto optimal front and comprehensive decision for final architecture via analytic hierarchy process. And the design method is implemented in python and a software platform is developed. Furthermore, within the paper a case study for A350 flight control actuation system is presented to testify the application of this methodology. Compared to the traditional hydraulic architecture, the optimal architecture is more competitive in weight, power and cost. At the same time, the optimal architecture is found in less than 30 minutes among 10〈sup〉75〈/sup〉 candidates, which greatly reduces the design cycle. This method deals with the problem in the design of flight control actuation system.〈/p〉〈/div〉

Description: 〈p〉Publication date: Available online 9 April 2019〈/p〉 〈p〉〈b〉Source:〈/b〉 Aerospace Science and Technology〈/p〉 〈p〉Author(s): N. Bartoli, T. Lefebvre, S. Dubreuil, R. Olivanti, R. Priem, N. Bons, J.R.R.A. Martins, J. Morlier〈/p〉 〈h5〉Abstract〈/h5〉 〈div〉〈p〉Surrogate models are often used to reduce the cost of design optimization problems that involve computationally costly models, such as computational fluid dynamics simulations. However, the number of evaluations required by surrogate models usually scales poorly with the number of design variables, and there is a need for both better constraint formulations and multimodal function handling. To address this issue, we developed a surrogate-based gradient-free optimization algorithm that can handle cases where the function evaluations are expensive, the computational budget is limited, the functions are multimodal, and the optimization problem includes nonlinear equality or inequality constraints. The proposed algorithm—super efficient global optimization coupled with mixture of experts (SEGOMOE)—can tackle complex constrained design optimization problems through the use of an enrichment strategy based on a mixture of experts coupled with adaptive surrogate models. The performance of this approach was evaluated for analytic constrained and unconstrained problems, as well as for a multimodal aerodynamic shape optimization problem with 17 design variables and an equality constraint. Our results showed that the method is efficient and that the optimum is much less dependent on the starting point than the conventional gradient-based optimization.〈/p〉〈/div〉

Description: 〈p〉Publication date: June 2019〈/p〉 〈p〉〈b〉Source:〈/b〉 Aerospace Science and Technology, Volume 89〈/p〉 〈p〉Author(s): S.M. Frolov, V.S. Aksenov, V.S. Ivanov, I.O. Shamshin, A.E. Zangiev〈/p〉 〈h5〉Abstract〈/h5〉 〈div〉〈p〉The air-breathing pulsed detonation thrust module (TM) for an aircraft designed for a subsonic flight at a speed of up to 120 m/s when operating on a standard aviation kerosene was developed using the analytical estimates and parametric multivariant three-dimensional (3D) calculations. The TM consists of an air intake with a check valve, a fuel supply system, a prechamber-jet ignition system and a combustion chamber with an attached detonation tube. An experimental sample of TM was fabricated, and its firing tests were carried out on a test rig with a thrust-measuring table. In firing tests, TM characteristics are obtained in the form of dependencies of effective thrust, aerodynamic drag and fuel-based specific impulse on fuel consumption at different speeds of the approaching air flow. It has been experimentally shown that the fuel-based specific impulse of the TM reaches 1000-1200 s, and the effective thrust developed by it reaches 180–200 N.〈/p〉〈/div〉

Description: 〈p〉Publication date: June 2019〈/p〉 〈p〉〈b〉Source:〈/b〉 Aerospace Science and Technology, Volume 89〈/p〉 〈p〉Author(s): Yu Wu, LeiLei Li, Xichao Su, Bowen Gao〈/p〉 〈h5〉Abstract〈/h5〉 〈div〉〈p〉Unmanned aerial-aquatic vehicle (UAAV) is a new type of aircraft that can navigate both in air and underwater. Considering that the diving motion of UAAV plays an important role in the performance of UAAV and is under exploration, the dynamics modeling and trajectory optimization problem are studied in this paper. The UAAV model used in this study is introduced firstly, and folded wings are adopted to reduce the drag in the diving process. Among the forces imposed on UAAV, fluid force is the most complicated and is calculated by the forces induced by ideal fluid and viscous fluid respectively. Based on the established dynamic model, the diving process is regarded as a free motion to avoid the instability during the control switch between air and water. Therefore, the trajectory of UAAV is determined by the initial states of diving process. To obtain the satisfactory trajectory under certain optimization index, an adaptive and global-best guided CS algorithm, named as improved cuckoo search (ICS) algorithm, is developed to strength the exploitation ability and search efficiency. Simulation results demonstrate that the established dynamical model of UAAV is rational and can reflect the characteristic of the diving motion. The proposed ICS algorithm performs better than the particle swarm optimization (PSO) algorithm and the standard CS algorithm both in optimizing the elapsed time of diving process and the terminal position error.〈/p〉〈/div〉

Description: 〈p〉Publication date: June 2019〈/p〉 〈p〉〈b〉Source:〈/b〉 Aerospace Science and Technology, Volume 89〈/p〉 〈p〉Author(s): Kunlin Cheng, Jiang Qin, Hongchuang Sun, Chaolei Dang, Silong Zhang, Xiaoyong Liu, Wen Bao〈/p〉 〈h5〉Abstract〈/h5〉 〈div〉〈p〉Hypersonic vehicle as next generation aircraft/spacecraft has broad applications, but its power supply and refrigeration are strictly limited by finite cold source. In this article, an integrated power generation and refrigeration system is developed, in which low-temperature fuel is utilized as cold source and high-temperature fuel is used as heat source. A novel combined generator based on closed-Brayton-cycle (CBC) and thermoelectric generator (TEG) is proposed to enhance electric power through extending the available temperature range of cold source. The integrated system model which consists of a refrigerator, a simple recuperated CBC and a three-stage TEG, is established to assess performance. Results indicate that the combined CBC-TEG generator has great potential in electric power enhancement. The power has an increase of 18.2% compared with single CBC. Power increase percentage of combined generator reduces with fuel outlet temperature in primary cooler. Moreover, decoupling the cooling and heating process of the combined generator is beneficial for the matching between its cold source and heat source. This research provides an innovative technical solution for the power generation and refrigeration on hypersonic vehicles under finite cold source conditions.〈/p〉〈/div〉

Description: 〈p〉Publication date: June 2019〈/p〉 〈p〉〈b〉Source:〈/b〉 Aerospace Science and Technology, Volume 89〈/p〉 〈p〉Author(s): Nader H. Al-Battal, David J. Cleaver, Ismet Gursul〈/p〉 〈h5〉Abstract〈/h5〉 〈div〉〈p〉Counter-flowing wall jets actuated on the upper surface of an airfoil are investigated for the purpose of controlling gust encounters. For periodic and transient actuations, phase-averaged lift force and particle image velocimetry measurements are presented for a NACA 0012 airfoil, at a Reynolds number of 660,000, for a range of reduced frequencies and three jet locations, 〈math xmlns:mml="http://www.w3.org/1998/Math/MathML" altimg="si1.gif" overflow="scroll"〉〈msub〉〈mrow〉〈mi〉x〈/mi〉〈/mrow〉〈mrow〉〈mi〉J〈/mi〉〈/mrow〉〈/msub〉〈mo stretchy="false"〉/〈/mo〉〈mi〉c〈/mi〉〈mo〉=〈/mo〉〈mn〉0.08〈/mn〉〈/math〉, 0.60 and 0.95. For periodic actuation, amplitude of lift oscillations decrease and phase delay increase with increasing reduced frequency. The effect of reduced frequency on the amplitude and phase is more significant for blowing locations near the leading-edge and with increasing angle of attack. Transient actuation reveals the slow response of the separated flow, and therefore lift, with the delay becoming more pronounced for blowing near the leading-edge. Estimated time constants are similar to previous observations for forced separation and reattachment.〈/p〉〈/div〉

Description: 〈p〉Publication date: June 2019〈/p〉 〈p〉〈b〉Source:〈/b〉 Aerospace Science and Technology, Volume 89〈/p〉 〈p〉Author(s): Luyao Zang, Defu Lin, Siyuan Chen, Hui Wang, Yi Ji〈/p〉 〈h5〉Abstract〈/h5〉 〈div〉〈p〉An on-line guidance algorithm is presented for high L/D hypersonic reentry vehicles using plane-symmetry bank-to-turn control method. The proposed guidance algorithm can generate a feasible trajectory at each guidance cycle during the reentry flight. In the longitudinal profile, height-range (H-R) and height-velocity (H-V) joint design method is proposed to rapidly generate the reference trajectory and the controlled bank angle to satisfy all constraints. Then, the quasi-equilibrium glide phenomenon is employed to extract the other controlled angle of attack corresponding to the reference trajectory. In the lateral profile, the bank angle reversal strategy is designed according to the cross-range angle threshold and the proportional navigation approach to reduce heading error. Simulation results for nominal and dispersed cases show that the proposed guidance algorithm is capable of generating a feasible trajectory rapidly that satisfies both path and terminal constraints.〈/p〉〈/div〉

Description: 〈p〉Publication date: Available online 26 June 2019〈/p〉 〈p〉〈b〉Source:〈/b〉 Aerospace Science and Technology〈/p〉 〈p〉Author(s): Huafei Du, Mingyun Lv, Jun Li, Weiyu Zhu, Lanchuan Zhang, Yifei Wu〈/p〉 〈h5〉Abstract〈/h5〉 〈div〉〈p〉Station-keeping endurance of high altitude balloon is the foundation of the assignments of environmental monitoring and communicational relaying. Exploiting the natural wind-field that varies with altitude by the altitude control system to extend the station-keeping endurance is proposed in this paper. A Matlab program is developed based on the theoretical model to simulate the station-keeping performance of the balloon in the real wind field. The trajectories of the balloon in different wind fields and the states of the balloon caused by the venting and pumping processes are discussed in detail. The results show that with the altitude control system it is possible to retain the balloon within the designated district for few days to a week. This can serve as a guideline for the design and initial flight tests of the serviceable high altitude balloon.〈/p〉〈/div〉

Description: 〈p〉Publication date: September 2019〈/p〉 〈p〉〈b〉Source:〈/b〉 Aerospace Science and Technology, Volume 92〈/p〉 〈p〉Author(s): Hongliang Wang, Xingling Shao, Jie Li, Jun Liu〈/p〉 〈div xml:lang="en"〉 〈h5〉Abstract〈/h5〉 〈div〉〈p〉This paper presents a velocity-free desired compensation trajectory tracking strategy integrated with adaptive robust integral of the sign of the error (RISE) feedback mechanism for quadrotors concerning parametric uncertainties and external disturbances. The original cascaded dynamics of quadrotor is derived in a strict form with additive disturbances. Then, the adaptive RISE feedback controllers are respectively constructed in position and attitude loop, in which the control gains in RISE terms are adaptively updated online to ensure the robustness against uncertainties. In addition, to alleviate the measurement noise effect arising from the actual velocity signals, the velocity states in the model-based feedforward control are replaced with their desired values, then the desired compensation adaptive RISE controllers that depend on the desired trajectory and output tracking errors are synthesized, where the design conservatism on selecting the control gain in RISE is eliminated without knowing the prior bound of uncertainties, and enhanced performance robustness is also retained in the absence of velocity information. It is shown via Lyapunov analysis that the proposed method can guarantee the tracking errors to converge to the origin with asymptotic performance despite of bounded disturbances. The effectiveness and superiority of proposed method are validated through extensive simulations and comparisons.〈/p〉〈/div〉 〈/div〉

Description: 〈p〉Publication date: Available online 26 June 2019〈/p〉 〈p〉〈b〉Source:〈/b〉 Aerospace Science and Technology〈/p〉 〈p〉Author(s): B.S. Yu, Z. Huang, L.L. Geng, D.P. Jin〈/p〉 〈h5〉Abstract〈/h5〉 〈div〉〈p〉This paper studies the spinning stability of a triangular tethered satellite formation that flies on a low earth orbit. The spinning around the center of mass at a constant angular rate gives rise to a periodic motion in the non-inertial orbital frame. Floquet theory is used to analyze the stability of the periodic motion. A dynamic similarity between the on-orbit dynamics and ground experimental models is built to construct an equivalent ground experiment to verify the stability analysis. The analytical and experimental results show that a stable periodic motion can be guaranteed if the spinning angular rate of the system exceeds a critical value.〈/p〉〈/div〉

Description: 〈p〉Publication date: Available online 26 June 2019〈/p〉 〈p〉〈b〉Source:〈/b〉 Aerospace Science and Technology〈/p〉 〈p〉Author(s): Chunyan Ling, Zhenzhou Lu, Kaixuan Feng〈/p〉 〈h5〉Abstract〈/h5〉 〈div〉〈p〉The failure credibility can be used to measure the safety level of the structure under the fuzzy inputs, but the computational efficiency for estimating the failure credibility is still a challenge. A novel method by combining the adaptive Kriging with fuzzy simulation (AK-FS) is proposed to efficiently estimate the failure credibility. The proposed method firstly employs the FS to transform the estimation of failure credibility into a classification problem, which can be viewed as a bi-level strategy. In the inner loop, a Kriging model for the actual complicated performance function is actively trained by U-learning function in the sample pool generated by FS until the convergent condition is satisfied, on which the samples are divided into failure group and safety group by the well-trained Kriging model instead of the actual performance function in the outer loop. Finally, the failure credibility is obtained by respectively searching the maximum joint membership degrees of the samples in these two groups. The proposed AK-FS method only evaluates the actual performance function in the process for constructing the Kriging model. Since the U-learning function can iteratively construct a sufficiently accurate Kriging model with model evaluations as less as possible, thus the failure credibility can be efficiently and accurately estimated by the proposed AK-FS method. The advantages of the proposed AK-FS method are demonstrated by several examples.〈/p〉〈/div〉

Description: 〈p〉Publication date: September 2019〈/p〉 〈p〉〈b〉Source:〈/b〉 Aerospace Science and Technology, Volume 92〈/p〉 〈p〉Author(s): Lu-Kai Song, Guang-Chen Bai, Cheng-Wei Fei〈/p〉 〈h5〉Abstract〈/h5〉 〈div〉〈p〉The performance and reliability of aircraft engine are seriously affected by multiple failures induced by multi-physical loads. Multi-failure probabilistic design is an effective measure to estimate the multi-failure response traits and quantify the multi-failure risk for the improvement of component reliability. In this paper, we propose a neural network regression-distributed collaborative strategy (NNR-DCS) based on a developed two-step error control technique, to improve the efficiency and accuracy of multi-failure probabilistic analysis. We firstly mathematically model NNR-DCS and then introduce the corresponding multi-failure probabilistic framework. With respect to various failure modes such as deformation failure, stress failure and strain failure, the multi-failure probabilistic analysis of a turbine bladed disk is conducted to evaluate the proposed method. From this simulation, we gain the probabilistic distribution features, reliability degree and sensitivity degree of each failure mode and overall failure modes on turbine bladed disk, which provides a useful reference for improving the reliability and performance of aircraft engine. The comparison of methods (Monte Carlo method, RSM, DCRSM, DCFRM, NNR and NNR-DCS) shows that the proposed NNR-DCS holds high efficiency and accuracy for multi-failure probabilistic analysis. The efforts of this study offer an effective way for multi-failure evaluation from a probabilistic perspective and shed light on the multi-objective reliability-based design optimization of complex structures besides turbine bladed disk.〈/p〉〈/div〉

Description: 〈p〉Publication date: June 2019〈/p〉 〈p〉〈b〉Source:〈/b〉 Aerospace Science and Technology, Volume 89〈/p〉 〈p〉Author(s): Jing Chang, Zongyi Guo, Jerome Cieslak, Weisheng Chen〈/p〉 〈h5〉Abstract〈/h5〉 〈div〉〈p〉This paper deals with the development of Integrated Guidance and Control (IGC) law for a class of hypersonic interceptors which are equipped with the infrared seeker. Since in practice the line-of-sight (LOS) angular rates are difficult to measure for this class of hypersonic interceptor, an unknown state observer based on high order sliding mode technique is firstly proposed to estimate the unknown state and disturbance precisely with the partial measurable state. Then, an adaptive Incremental Backstepping (IBS) scheme that relies on estimates of the LOS angular rates, angular acceleration measurements of the current control deflections is proposed to achieve robust tracking of a maneuvering target. The stability analysis of the closed-loop system is also conducted. At last, a series of simulation results are presented to show the great potential of the proposed IGC method in interception accuracy, even if the LOS angle rates are unmeasurable. It is shown that the proposed scheme is not only more robust to uncertainties in the hypersonic interceptor dynamics compared to conventional IGC, but also has better trajectory characteristic with less control effort. The hit-to-kill interception is achieved in the simulation results.〈/p〉〈/div〉

Description: 〈p〉Publication date: June 2019〈/p〉 〈p〉〈b〉Source:〈/b〉 Aerospace Science and Technology, Volume 89〈/p〉 〈p〉Author(s): M. Bouziane, A.E.M. Bertoldi, P. Milova, P. Hendrick, M. Lefebvre〈/p〉 〈h5〉Abstract〈/h5〉 〈div〉 〈p〉To investigate the effects of oxidizer injection on the performance of hybrid rocket motors (HRMs), we have designed, manufactured and tested four types of injector: showerhead (SH), hollow-cone (HC), pressure-swirl (PSW) and vortex (VOR). This study is motivated by the fact that the experimental measurements of N〈sub〉2〈/sub〉O/paraffin firings are poorly presented in the open literature. Besides few publications are dedicated to the characterization of novel types of injectors in hybrid rocket propulsion application such as HC, PSW, and VOR. It is advantageous that the study was conducted in the same motor configuration, with the advantage that it allowed to compare the performance of different types of injectors.〈/p〉 〈p〉This paper analyzes the influence of the oxidizer injector design on the main performance parameters, such as fuel regression rate, specific impulse and combustion efficiency. First, in order to observe injector spray qualities, a series of cold tests using liquid water and liquid nitrous oxide are carried out, providing a good understanding of the spray profiles. Then, the motor performance data is obtained by a series of firing tests using N〈sub〉2〈/sub〉O as oxidizer and paraffin as fuel. Comparison of the various injectors data is made with the same average oxidizer mass flux and feeding pressure. The showerhead is used as a benchmark in this study. During this experimental analysis, the VOR injector exhibits the highest regression rate, followed by HC and SH. Because the assessment of the regression rate was not enough to explain all the effects of the injectors on the motor performance, firing tests of small-scale hybrid motor have been carried out. In terms of spray properties, the PSW exhibits significant differences compared with the other injectors; it generates the smallest Sauter Mean Diameter (SMD) in the formed spray and achieves good atomization. In spite of the fact that the PSW injector leads to the lowest regression rate, it provides good specific impulse, increases the oxidizer-to-fuel ratio (〈math xmlns:mml="http://www.w3.org/1998/Math/MathML" altimg="si1.gif" overflow="scroll"〉〈mi〉O〈/mi〉〈mo stretchy="false"〉/〈/mo〉〈mi〉F〈/mi〉〈/math〉), as well as a uniform and smooth consumption of the paraffin fuel grain. VOR leads to the highest specific impulse. In terms of stability, VOR, HC and SH injectors exhibit lower oscillations in the chamber pressure. Some observations are made on exhaust plume intensity developed during combustion, and in firing tests with SH a blow-out phenomenon occur often. Similarly to liquid engines, it is possible in hybrid motors to increase the global propulsive performance using alternative designs of the injection system.〈/p〉 〈/div〉

Description: 〈p〉Publication date: June 2019〈/p〉 〈p〉〈b〉Source:〈/b〉 Aerospace Science and Technology, Volume 89〈/p〉 〈p〉Author(s): Yuxi Luo, Fengbo Wen, Shuai Wang, Shibo Zhang, Songtao Wang, Zhongqi Wang〈/p〉 〈h5〉Abstract〈/h5〉 〈div〉〈p〉In this study, the structure of a harbor seal vibrissa, or whisker, is applied to the trailing edge of a high subsonic turbine blade to reduce wake loss. A method of generating biomimetic geometry is provided. In order to study the mechanism of wake loss reduction for the biomimetic trailing edge, four trailing edge shapes are studied numerically: with circular (prototype blade), elliptic, sinusoidal and biomimetic. The delayed detached-eddy simulation (DDES) approach is used to predict the flow details around the blades. Results show that a substantial reduction of mixing loss is obtained for the blade with a biomimetic trailing edge. The coherent structures of Karman vortices are completely suppressed and transformed into smaller vortices with distinct three-dimensional characteristics owing to the influence of biomimetic trailing edge. Such small vortices dissipate quickly during the mixing with the mainstream, so the wake area for biomimetic trailing edge case is much narrower, and the velocity in the wake recovers more quickly compared with the situation in the other three cases. The vortices shedding from the suction side and pressure side move from the geometric peaks to the valleys when moving downstream.〈/p〉〈/div〉

Description: 〈p〉Publication date: June 2019〈/p〉 〈p〉〈b〉Source:〈/b〉 Aerospace Science and Technology, Volume 89〈/p〉 〈p〉Author(s): Zhenhua Li, Bing He, Minghao Wang, Haoshen Lin, Xibin An〈/p〉 〈div xml:lang="en"〉 〈h5〉Abstract〈/h5〉 〈div〉〈p〉Considering the requirements of saturation attack, a time-coordination entry guidance algorithm for multi-hypersonic vehicles is proposed in this paper, which is composed of three parts: generation of coordination time, predictive lateral guidance based rough adjustment of reentry flight time and longitudinal predictor-corrector guidance based time coordination. Firstly, the coordination flight time is generated by analyzing the time range of entry flight for each hypersonic vehicle. Next, in the lateral profile, the predictive lateral guidance is employed to roughly adjust the reentry flight time of the vehicle, and provide feasible time range of coordinative flight for longitudinal predictor-corrector guidance. In the longitudinal profile, a quadratic bank angle magnitude profile is designed and adjusted by the Newton iteration scheme to meet the range and time constraints, and realize the time-coordination entry guidance for multi-hypersonic vehicles. As the impacts of the lateral motion and longitudinal motion on entry flight time are both considered, the proposed algorithm greatly extends the feasible time range of coordinative flight, which is more suitable for engineering practice. Finally, the feasibility and robustness of the proposed time-coordination entry guidance algorithm are tested and verified by the multi-mission and multi-vehicle guidance simulations.〈/p〉〈/div〉 〈/div〉

Description: 〈p〉Publication date: June 2019〈/p〉 〈p〉〈b〉Source:〈/b〉 Aerospace Science and Technology, Volume 89〈/p〉 〈p〉Author(s): Hasan Keshavarzian, Kamran Daneshjou〈/p〉 〈h5〉Abstract〈/h5〉 〈div〉〈p〉In ground effect conditions, the wake of Quad-copter rotors interacting with the ground causes significant perturbation to the flow near the rotor blades, as well as to the body. These interactions have serious effects on the handling qualities and cause instability in flight. Most studies in the ground effect are focused on hover and landing capabilities of rotorcraft. In this paper, a comprehensive non-linear model of quadrotor is provided in the state-space, which is suitable for all flight types near the ground. First of all, it is tried to add a ground surface quality coefficient in modified ground effect relation by experimental investigations. The experiments have been done in outdoor during calm days with no wind. In the next step, a robust non-linear control strategy is designed based on the modified model. The effects of ground effect and the increase of the advance ratio are then investigated on controller performance. The simulation results show that this model is closer to the actual flight conditions and the stability and the trajectory tracking are significantly improved by implementation of this control strategy. So, this strategy provides an appropriate platform to run other methods without the need for iterative learning techniques and high battery energy consumption.〈/p〉〈/div〉