The precision forming of thin-walled components has been urgently needed in aviation and aerospace field. However, the wrinkling induced by the compressive instability is one of the major defects in thin-walled part forming. The initiation and growth of the wrinkles are interactively affected by many factors such as stress states, mechanical properties of the material, geometry of the workpiece and boundary conditions. Especially when the forming process involves complicated boundary conditions such as multi-dies constrains, the perturbation of clearances between workpiece and dies and the contact conditions changing in time and space, etc., the predication of the wrinkling is further complicated. In this paper, the current prediction methods were summarized including the static equilibrium method, the energy method, the initial imperfection method,the eigenvalue buckling analysis method, the static-implicit finite element method and the dynamicexplicit finite element method. Then, a systematical comparison and summary of these methods in terms of their advantages and limitations are presented. By using a combination of explicit FE method, initial imperfection and energy conservation, a hybrid method is recommended to predict plastic wrinkling in thin-walled part forming. Finally, considering the urgent requirements of complex thin-walled structures’ part in aviation and aerospace field, the trends and challenges in wrinkling prediction under complicated boundary conditions are presented.
Continuously rotating detonation engine(CRDE) is a focus for concern in the field of aerospace propulsion. It has several advantages, including one-initiation, high thermal efficiency and simple structure. Due to these characteristics, it is expected to bring revolutionary advancements to aviation and aerospace propulsion systems and now has drawn much attention throughout the world. In this paper, an overview of the development of CRDE is given from several aspects:basic concepts, applications, experimental studies, numerical simulations, and so on. Representative results and outstanding contributions are summarized and the unresolved issues for further engineering applications of CRDE are provided.
How to control the microstructure of titanium alloy bars is important to fabricating high-performance aerial forgings. This paper gives a thorough survey of the manufacturing methods and microstructure control techniques for titanium alloy bars. It summarizes the effects of processing parameters on the mechanisms and laws of microstructure evolution during b working and(a + b) working, including the kinetics and grains size of dynamic recrystallization(DRX) during b deformation and the kinetics and grains size of spheroidization during(a + b) deformation. The trends in microstructure control techniques are presented for fabricating titanium alloy bars with high efficiency, low cost, and high quality by means of b/(a + b) working, and the puzzles and challenges in the future are also pointed out.
A new approach for the prediction of lift, drag, and moment coefficients is presented.This approach is based on the support vector machines(SVMs) methodology and an optimization meta-heuristic algorithm called extended great deluge(EGD). The novelty of this approach is the hybridization between the SVM and the EGD algorithm. The EGD is used to optimize the SVM parameters. The training and validation of this new identification approach is realized using the aerodynamic coefficients of an ATR-42 wing model. The aerodynamic coefficients data are obtained with the XFoil software and experimental tests using the Price–Pa?¨doussis wind tunnel.The predicted results with our approach are compared with those from the XFoil software and experimental results for different flight cases of angles of attack and Mach numbers. The main purpose of this methodology is to rapidly predict aircraft aerodynamic coefficients.
Applications of a novel curve-fitting technique are presented to efficiently predict the motion of the vortex filament, which is trailed from a rigid body such as wings and rotors. The governing equations of the motion, when a Lagrangian approach with the present curve-fitting method is applied, can be transformed into an easily solvable form of the system of nonlinear ordinary differential equations. The applicability of Be′zier curves, B-spline, and Lagrange interpolating polynomials is investigated. Local Lagrange interpolating polynomials with a shift operator are proposed as the best selection for applications, since it provides superior system characteristics with minimum computing time, compared to other methods. In addition, the Gauss quadrature formula with local refinement strategy has been developed for an accurate prediction of the induced velocity computed with the line integration of the Biot–Savart law. Rotary-wing problems including a vortex ring problem are analyzed to show the efficiency, accuracy, and flexibility in the applications of the proposed method.
This paper describes a simplified transition model based on the recently developed correlation-based c Rehttransition model. The transport equation of transition momentum thickness Reynolds number is eliminated for simplicity, and new transition length function and critical Reynolds number correlation are proposed. The new model is implemented into an in-house computational fluid dynamics(CFD) code and validated for low and high-speed flow cases, including the zero pressure flat plate, airfoils, hypersonic flat plate and double wedge. Comparisons between the simulation results and experimental data show that the boundary-layer transition phenomena can be reasonably illustrated by the new model, which gives rise to significant improvements over the fully laminar and fully turbulent results. Moreover, the new model has comparable features of accuracy and applicability when compared with the original c Rehtmodel. In the meantime, the newly proposed model takes only one transport equation of intermittency factor and requires fewer correlations, which simplifies the original model greatly. Further studies, especially on separationinduced transition flows, are required for the improvement of the new model.
Gust load alleviation(GLA) tests are widely conducted to study the effectiveness of the control laws and methods. The physical parameters of models in these tests are aeroelastic scaled,while the scaling of GLA control system is always unreached. This paper concentrates on studying the scaling laws of GLA control system. Through theoretical demonstration, the scaling criterion of a classical PID control system has been come up and a scaling methodology is provided and verified. By adopting the scaling laws in this paper, gust response of the scaled model could be directly related to the full-scale aircraft theoretically under both open-loop and closed-loop conditions.Also, the influences of different scaling choices of an important non-dimensional parameter, the Froude number, have been studied in this paper. Furthermore for practical application, a compensating method is given when the theoretical scaled actuators or sensors cannot be obtained. Also,the scaling laws of some non-linear elements in control system such as the rate and amplitude saturations in actuator have been studied and examined by a numerical simulation.
A theoretical nonlinear aeroelastic response analysis for a flexible high-aspect ratio wing excited by harmonic gust load is presented along with a companion wind tunnel test. A multidisciplinary coupled numerical calculation is developed to simulate the flexible model wing undergoing gust load in the time domain via discrete nonlinear finite element structural dynamic analysis and nonplanar unsteady vortex lattice aerodynamic computation. A dynamic perturbation analysis about a nonlinear static equilibrium is also used to determine the small perturbation flutter boundary. A novel noncontact 3-D camera measurement analysis system is firstly used in the wind tunnel test to obtain the spatial large deformation and responses. The responses of the flexible wing under different static equilibrium states and frequency gust loads are discussed. The fair to good quantitative agreements between the theoretical and experimental results demonstrate that the presented analysis method is an acceptable way to predict the geometrically nonlinear gust response for flexible wings.
The effects of blade lean and vortex design on the aerodynamics of a turbine entry nozzle guide vane(NGV) are considered using computational fluid dynamics. The aim of the work is to address some of the uncertainties which have arisen from previous studies where conflicting results have been reported for the effect on the NGV. The configuration was initially based on the energy efficient engine turbine which also served as the validation case for the computational method. A total of 17 NGV configurations were evaluated to study the effects of lean and vortex design on row efficiency and secondary kinetic energy. The distribution of mass flow ratio is introduced as an additional factor in the assessment of blade lean effects. The results show that in the turbine entry NGV, the secondary flow strength is not a dominant factor that determines NGV losses and therefore the changes of loading distribution due to blade lean and the associated loss mechanisms should be regarded as a key factor. Radial mass flow redistribution under different NGV lean and twist is demonstrated as an addition key factor influencing row efficiency.
A numerical study was performed to explore the unsteady interaction between the upstream propeller and the downstream swirl recovery vane(SRV) by transient simulations. Much larger fluctuations of thrust coefficient were observed on the vane, which indicates that the variations of the total efficiency depend mainly on the working performance of the stator. The harmonic loads of the decomposed unsteady blade-surface pressures show that the stator experiences about ten times higher of unsteadiness compared with the rotor. Notable changes appear at the vane leading edge due to the potential disturbance as well as the sweeping effects from the wake of the upstream propeller, whereas more significant unsteadiness occurs at the stator tip region as a result of the interaction between the rotor/stator tip vortices. The visualization of vortex structures addresses that the rotor tip vortex has a dominant effect on the stator tip vortex since the latter one starts right at the impingement location on the vane top in this configuration. Furthermore,a longer and a shorter SRV were investigated based on the original case to explore different interaction patterns for the rotor/stator tip vortices. Weaker effects have been observed as expected.
Numerical simulations are performed to study the aeroelastic responses of an elastically suspended airfoil in transonic buffet flow, by coupling the unsteady Reynolds-averaged NavierStokes(RANS) equations and structural motion equation. The current work focuses on the characteristic analysis of the lock-in phenomenon. Great attentions are paid to studying the frequency range of lock-in and the effects of the three parameters, namely the structural natural frequency,mass ratio and structural damping, on lock-in characteristic of the elastic system in detail. It is found that when the structural natural frequency is close to the buffet frequency, the coupling frequency of the elastic system is no longer equal to the buffet frequency, but keeps the same value as the structural natural frequency. The frequency lock-in occurs and stays present until the structural nature frequency is near the double buffet frequency. It means that the lock-in presents within a broad range, of which the lower threshold is near the buffet frequency, while the upper threshold is near the double buffet frequency. Moreover, the frequency range of lock-in is affected by mass ratio and structural damping. The lower the mass ratio and structural damping are, the wider the range of lock-in will be. The upper threshold of lock-in grows with the mass ratio and structural damping decreasing, but the lower threshold always keeps the same.
The flutter characteristics of folding control fins with freeplay are investigated by numerical simulation and flutter wind tunnel tests. Based on the characteristics of the structures, fins with different freeplay angles are designed. For a 0° angle of attack, wind tunnel tests of these fins are conducted, and vibration is observed by accelerometers and a high-speed camera. By the expansion of the connected relationships, the governing equations of fit for the nonlinear aeroelastic analysis are established by the free-interface component mode synthesis method. Based on the results of the wind tunnel tests, the flutter characteristics of fins with different freeplay angles are analyzed. The results show that the vibration divergent speed is increased, and the divergent speed is higher than the flutter speed of the nominal linear system. The vibration divergent speed is increased along with an increase in the freeplay angle. The developed free-interface component mode synthesis method could be used to establish governing equations and to analyze the characteristics of nonlinear aeroelastic systems. The results of the numerical simulations and the wind tunnel tests indicate the same trends and critical velocities.
To balance the contradiction between comprehensiveness of system-of-systems(So S)description and cost of modeling and simulation, a non-uniform hybrid strategy(NUHYS) is proposed. NUHYS groups elements of an So S operation into system community or relatively independent system based on contributors complexity and focus relationship according to the focus of So S problem. Meanwhile, modeling methods are categorized based on details attention rate and dynamic attention rate, seeking for matching contributors. Taking helicopter rescue in earthquake relief as an example, the procedure of applying NUHYS and its effectiveness are verified.
In the present article, the linear and the nonlinear deformation behaviour of functionally graded(FG) spherical shell panel are examined under thermomechanical load. The temperaturedependent effective material properties of FG shell panel are evaluated using Voigt’s micro-mechanical rule in conjunction with power-law distribution. The nonlinear mathematical model of the FG shell panel is developed based on higher-order shear deformation theory and Green-Lagrange type geometrical nonlinearity. The desired nonlinear governing equation of the FG shell panel is computed using the variational principle. The model is discretised through suitable nonlinear finite element steps and solved using direct iterative method. The convergence and the validation behaviour of the present numerical model are performed to show the efficacy of the model.The effect of different parameters on the nonlinear deformation behaviour of FG spherical shell panel is highlighted by solving numerous examples.
A hierarchic optimization strategy based on the offline path planning process and online trajectory planning process is presented to solve the trajectory optimization problem of multiple quad-rotor unmanned aerial vehicles in the collaborative assembling task. Firstly, the path planning process is solved by a novel parallel intelligent optimization algorithm, the central force optimization-genetic algorithm(CFO-GA), which combines the central force optimization(CFO)algorithm with the genetic algorithm(GA). Because of the immaturity of the CFO, the convergence analysis of the CFO is completed by the stability theory of the linear time-variant discrete-time systems. The results show that the parallel CFO-GA algorithm converges faster than the parallel CFO and the central force optimization-sequential quadratic programming(CFO-SQP) algorithm. Then,the trajectory planning problem is established based on the path planning results. In order to limit the range of the attitude angle and guarantee the flight stability, the optimized object is changed from the ordinary six-degree-of-freedom rigid-body dynamic model to the dynamic model with an inner-loop attitude controller. The results show that the trajectory planning process can be solved by the mature SQP algorithm easily. Finally, the discussion and analysis of the real-time performance of the hierarchic optimization strategy are presented around the group number of the waypoints and the equal interval time.
For the terminal guidance problem of missiles intercepting maneuvering targets in the three-dimensional space, the design of guidance laws for non-decoupling three-dimensional engagement geometry is studied. Firstly, by introducing a finite time integral sliding mode manifold, a novel guidance law based on the integral sliding mode control is presented with the target acceleration as a known bounded external disturbance. Then, an improved adaptive guidance law based on the integral sliding mode control without the information of the upper bound on the target acceleration is developed, where the upper bound of the target acceleration is estimated online by a designed adaptive law. The both presented guidance laws can make sure that the elevation angular rate of the line-of-sight and the azimuth angular rate of the line-of-sight converge to zero in finite time. In the end, the results of the guidance performance for the proposed guidance laws are presented by numerical simulations. Although the designed guidance laws are developed for the constant speed missiles, the simulation results for the time-varying speed missiles are also shown to further confirm the designed guidance laws.
To solve the receding horizon control(RHC) problem in an online manner, a novel numerical method called the indirect Radau pseudospectral method(IRPM) is proposed in this paper. Based on calculus of variations and the first-order necessary optimality condition, the RHC problem for linear time-varying(LTV) system is transformed into the two-point boundary value problem(TPBVP). The Radau pseudospectral approximation is employed to discretize the TPBVP into well-posed linear algebraic equations. The resulting linear algebraic equations are solved via a matrix partitioning approach afterwards to obtain the optimal feedback control law.For the nonlinear system, the linearization method or the quasi linearization method is employed to approximate the RHC problem with successive linear approximations. Subsequently, each linear problem is solved via the similar method which is used to solve the RHC problem for LTV system.Simulation results of three examples show that the IRPM is of high accuracy and of high computation efficiency to solve the RHC problem and the stability of closed-loop systems is guaranteed.
A novel biased proportional navigation guidance(BPNG) law is proposed for the close approach phase, which aims to make the spacecraft rendezvous with the target in specific relative range and direction. Firstly, in order to describe the special guidance requirements, the concept of zero effort miss vector is proposed and the dangerous area where there exists collision risk for safety consideration is defined. Secondly, the BPNG, which decouples the range control and direction control, is designed in the line-of-sight(LOS) rotation coordinate system. The theoretical analysis proves that BPNG meets guidance requirements quite well. Thirdly, for the consideration of fuel consumption, the optimal biased proportional navigation guidance(OBPNG) law is derived by solving the Schwartz inequality. Finally, simulation results show that BPNG is effective for the close approach with the ability of evading the dangerous area and OBPNG consumes less fuel compared with BPNG.
As powerful torque amplification actuators, control moment gyros(CMGs) are often used in the attitude control of many state-of-the-art high resolution satellites. However, the disturbance generated by the CMGs can not only reduce the attitude stability of a satellite but also deteriorate the performance of optic payloads. Currently, CMG vibration isolators are widely used to target this problem. The isolators can affect the singularity of the CMG system as they are placed between the CMGs and the satellite bus and provide additional freedoms to the CMG system due to their flexibility. The formulation of the output torque of a CMG is studied first considering the dynamic imbalance of its spin rotor and then the deformation angle as a result of the isolator’s flexibility is calculated. With the additional freedoms, the influence of isolator on the singularity problem is studied and a new steering logic to escape from the singular states is proposed.
Considering defects of current single celestial-body positioning methods such as discontinuity and long period, a new sun positioning algorithm is herein put forward. Instead of traditional astronomical spherical trigonometry and celestial coordinate system, the proposed new positioning algorithm is built by theory of mechanisms. Based on previously derived solar vector equations(from a C1R2P2 series mechanism), a further global positioning method is developed by inverse kinematics. The longitude and latitude coordinates expressed by Greenwich mean time(GMT) and solar vector in local coordinate system are formulated. Meanwhile, elimination method of multiple solutions, errors of longitude and latitude calculation are given. In addition, this algorithm has been integrated successfully into a mobile phone application to visualize sun positioning process. Results of theoretical verification and smart phone’s test demonstrate the validity of presented coordinate’s expressions. Precision is shown as equivalent to current works and is acceptable to civil aviation requirement. This new method solves long-period problem in sun sight running fixing and improves applicability of sun positioning. Its methodology can inspire development of new sun positioning device. It would be more applicable to be combined with inertial navigation systems for overcoming discontinuity of celestial navigation systems and accumulative errors of inertial navigation systems.
Out of phase(OP) thermal mechanical fatigue(TMF) behavior of a directionally solidified(DS) superalloy DZ125 was experimentally and numerically studied. Two different temperature conditions, which are 500–1000 °C and 400–900 °C, were considered in the present research.Stress and strain responses as well as fatigue life results were presented and discussed. Scanning electron microscope(SEM) and metallographic analysis were used to study the damage mechanism. An oxidation assisted crack initiation and propagation phenomenon were found to explain the shorted life under TMF cycles. In order to characterize the stress and strain deformations under TMF loadings, a modified Chaboche’s constitutive model was applied. Additionally, the TMF life of the material was modeled and predicted by Neu–Sehitoglu damage law with high accuracy.
It is well-known that the application of ultrasound during liquid to solid transitions for alloys can refine the solidification microstructure and thus improves the mechanical properties. However, most published work focuses on single phase dendritic growth, whereas little has been conducted on the multiphase alloys with complicated phase transformations during solidification. In this work, the solidification process of ternary Cu40Sn45Sb15 alloy was realized within intensive ultrasonic field with a resonant frequency of 20 k Hz and ultrasound power from 0 W to 1000 W. The ultrasound refines the size of the primary e(Cu3Sn) intermetallic compound by two orders of magnitudes. If the ultrasound power increases to 1000 W, g(Cu6Sn5) phase nucleates and grows directly from parent liquid phase without the occurrence of peri-eutectic reaction on the top of the alloy sample where the ultrasound intensity is sufficiently high. These microstructural variations lead to the enhancement of compressive strength and elasticity modulus of ternary Cu40Sn45Sb15 alloy.
In aircraft assembly, interlayer burr formation in dry drilling of stacked metal materials is a common problem. Traditional manual deburring operation seriously affects the assembly quality and assembly efficiency, is time-consuming and costly, and is not conducive to aircraft automatic assembly based on industrial robot. In this paper, the formation of drilling exit burr and the influence of interlayer gap on interlayer burr formation were studied, and the mechanism of interlayer gap formation in drilling stacked aluminum alloy plates was investigated, a simplified mathematical model of interlayer gap based on the theory of plates and shells and finite element method was established. The relationship between interlayer gap and interlayer burr, as well as the effect of feed rate and pressing force on interlayer burr height and interlayer gap was discussed. The result shows that theoretical interlayer gap has a positive correlation with interlayer burr height and preloading pressing force is an effective method to control interlayer burr formation.
Sealing clearance is a key factor for a metal rubber seal’s sealability. The expansion coefficient and expansion deformation in the radial direction of metal rubber have been obtained through a thermal expansion experiment of metal rubber. The influence of the elastic modulus to the sealing clearance has been analyzed theoretically. By combining the temperature and elasticity factors of metal rubber with the elastic mechanics theory, the calculation formula of the sealing clearance has been derived, and the values of the sealing clearance and the leakage rate in certain working conditions have been calculated. Experimental results are consistent with calculation results in a high degree. The calculation formula of the sealing clearance can explain the influences of the temperature and elastic modulus factors of metal rubber on the sealing clearance. It can provide guidance for the study of sealing mechanism of metal rubber seals.