Journal of Aerospace Engineering

In this work, the modeling of triaxially braided composites was explored through a subcell approach using an improved semianalytical discretization scheme. The unit cell of the braided composite was divided into four unique subcells, each approximated by a mosaic stacking of unidirectional composite plies and modeled through the use of layered-shell elements within the finite-element model. Two subcell discretization schemes were investigated: a model explicitly capturing pure matrix regions, and a model which absorbed pure matrix pockets into neighboring tow plies. Differences in the mesostructure between single-ply and multi-ply braid coupons were addressed through modifications to the subcell discretization. The absorbed matrix model simulated the unique out-of-plane deformations observed experimentally in single-ply tensile tests with acceptable moduli predictions. An investigation of single-shell versus multi-shell coupons for the analysis of multi-ply braids revealed the through-thickness modeling approach was found to have a significant effect on the apparent transverse modulus. Improved moduli predictions in both the axial and transverse directions were obtained by explicitly modeling braided plies with individual layers of shell elements.

A popular application of propellant containing of aluminum powder is as fuel for solid rocket motors (SRMs). The combustion gas contains large amounts of alumina particles, which will worsen the SRM thermal protection system by inducing severe mechanical erosion on the surface of the insulation material exposed directly to the high-temperature and high-velocity gas and solid two-phase flow. To protect the structures of SRMs from being damaged, a thermal protection system is required, and it must be optimally designed to be as thin as possible, but thick enough to survive the SRM operating time. To optimize the thickness of insulation material, the particle erosion model should be set up. But most existing particle erosion models are the empirical expressions obtained by experiment or direct reference to pipeline particle erosion models, which limit the application of the model. The ablation process and the flow are coupled with one another. So, to optimize thickness, numerical calculation of dynamic ablation of the insulation material in SRMs should be carried out. In this work, the ethylene propylene diene monomer (EPDM) ablative composite used as heat shield was analyzed. The physical, chemical, and mechanical processes of ablation were expounded. Models for insulation thermal decomposition and thermochemical ablation were constructed, in which pyrolysis and carbonization temperature, pyrolysis rate, latent heat and pyrolysis products, and the main reaction chemical equation of the thermal chemical ablation were researched and confirmed. The coupling of physics, chemistry, and mechanics is through the charring layer in the process of ablation of the insulation material. The charring layer was treated as a porous medium. The ablation models and its numerical method were applied to calculate the ablation of insulation material in an experimental SRM. The computed result is in good agreement with that of the experiment, which shows that the model and numerical method have good accuracy.

A simple, easy-to-use method to compute vehicle drag at low and moderate supersonic speeds is presented in this study. A novel version of the Newtonian sine-squared law is proposed and used to compute supersonic pressure or wave drag for various geometries. The new version implicitly includes shockwave contribution not present in the original law. The results obtained are compared with the those found in the literature and discussed in order to estimate the applicability and validity of the method. The method’s suitability for optimization purposes is also briefly considered.

Numerical investigation of flow phenomena in a scramjet combustor has been performed for different geometric and operating parameters. The present investigation aims to find the optimal geometric parameters for a better fuel injection system with maximum combustion efficiency. Two-dimensional unsteady equations governing the compressible, turbulent reacting flow are solved using a commercial CFD solver. A combination of eddy dissipation (ED) and finite rate chemistry (FRC) models are used to model the combustion. The effect of divergence angles and scaling on the performance of a scramjet combustor is reported here. The effect of shocks generated by the strut and inlet conditions of a scramjet combustor on combustion efficiency has been reported. Present results show that the divergence angle and inlet conditions of combustors have significant effects on performance. A multiple-strut combustor has shown higher efficiency than a single-strut combustor. The modifications that improved the combustion efficiency of a DLR combustor are reported. The present simulation results match well with the experimental results available from the literature.

This paper considers integrated guidance and control system with input saturation for homing missiles based on the three channels independence design idea. An auxiliary system is first introduced into the integrated guidance and control model for the pitch channel. Then, a smooth function and a Nussbaum type function are employed in the dynamic surface control design to deal with the problem of input saturation constraint. The proposed control scheme guarantees the desired interception performance and the stability of the missile dynamics. The similar control scheme is further applied to the yaw channel and roll channel respectively. Nonlinear 6-DOF simulations are carried out to demonstrate the effect of the approach.

Advanced performances in nanospacecraft are typically associated with high onboard power availability, requiring deployed solar arrays. Nanospacecraft geometrical and weight constraints do not allow the use of reversible dynamics and lock systems used in conventional spacecraft. Most nanospacecraft deployed solar panel systems are therefore based on compact and reliable miniaturized mechanisms, such as simple passively deployed preloaded springs. The use of these kinds of deployed solar arrays may interfere with the performance of the attitude control system. Agile nanospacecraft, designed to perform rapid attitude maneuvers, may potentially be affected by the solar array dynamical interaction with the main satellite body motion. In addition, the presence of inherently nonlinear elements, such as bounds on allowed rotations and preloaded springs, deserve a dedicated analysis. The analysis developed in this paper is based on the performance of state-of-the art agile nanosatellite attitude control systems, specifically developed for delivering high torque and fast attitude maneuvering, such as miniaturized control moment gyros. By numerical simulation, it is shown that, despite the simplicity of the deployment system and the absence of locking mechanisms, the attitude dynamics remains substantially unaffected by the solar array dynamics. This is obtained straightforwardly by properly sizing the spring stiffness and by a sufficient level of preload.

The paper deals with the effect of moisture and temperature on the buckling response of a laminated composite plate subjected to hygrothermomechanical loadings. Mechanical loading consists of uniaxial, biaxial, and shear. The distribution of temperature and moisture on the surface is considered to be uniform. The degradation in material properties due to moisture and temperature is taken into account using a micromechanical model. The mathematical formulation is based on higher order shear deformation theory and von Karman’s nonlinear kinematics. A $C^0$ finite-element method based on higher-order shear deformation plate theory is used for deriving the standard Eigenvalue problem. A Taylor series based mean-centered first-order perturbation technique is used to find out the second-order statistics of the hygrothermal buckling loads. The effects of temperature rise, moisture concentration, fiber-volume fraction, and plate parameters on buckling response of the plate are presented.

Understanding the mechanical behavior of lunar regolith is of great importance for future missions to the Moon and other similar extraterrestrial environments. However, due to the scarcity of lunar regolith a number of simulants have been developed to facilitate experimental testing. This paper presents the geotechnical properties of an inexpensive lunar regolith simulant named Black Point 1 (BP-1). An experimental program including particle-size distribution, microscopy observations using scanning electron microscope (SEM) images, density measurements, compressibility, and shear strength was performed. Additionally, BP-1 was compared with regolith recovered from lunar missions and a number of its simulants. The physical properties of BP-1 were found to be similar to other existing simulants and to the natural lunar regolith.

The modified Picard-Chebyshev method, due to its formulation, has the potential for computationally efficient orbit propagation in a parallel computing environment. However, it has not found extensive use in commercial software applications, primarily due to its instability over long orbit propagation times. In this paper, the method was evaluated based on the propagation of several test orbits with a high-precision force model available through NASA’s Java Astrodynamics toolkit. The goal was to analyze the stability properties of the method, with emphasis on its dependence on the initial guess and the order of the approximating Chebyshev polynomial. Simulation results indicated a dependence of the method on the precision of the computation as well. Additionally, the performance of the method when implemented in parallel was evaluated by propagating the orbits on a multicore processing unit and comparing run times to well-known sequential numerical propagators. The preliminary results show that the run-time performance of the MPCM improves with the complexity in the computation, but further investigations are necessary to establish the capabilities of the method.

In aerospace engine industry, large casted components are, because of sustainability considerations, being replaced by smaller parts that are welded together. This reduces weight because some parts can be made of lighter material. It also opens up for use of platforms. The division of a large component into smaller parts is called form division. The form division affects the geometrical robustness in the weld splitlines between the parts and thereby the weldability. By optimizing the robustness in weld splitlines, conditions for welding can be improved. A greedy algorithm for weld splitline division is described and exemplified on aerospace case studies.

A novel sun-seeking eternal flight solar-powered airplane (SESPA) for Venus exploration is proposed with consideration of the high atmospheric density and long rotation period of Venus. The flight track is in accordance with the latitude circle of Venus. The flight speed equals the resultant velocity of the zonal wind speed and the rotation speed of Venus, and the flight direction is opposite to the resultant velocity. The SESPA can fly consistently at daytime under the local midday, thus increasing the utilization ratio of solar cells and eliminating the need for a battery. This feature reduces the system complexity and mission costs of the SESPA. This paper introduces the atmospheric characteristics of Venus and analyzes the feasibility of a SESPA for Venus exploration. A set of constraints of realizing sun-seeking eternal flight is derived. The designed SESPA is feasible for the atmospheric flight only when all constraints have been satisfied. Finally, the favorable flight condition for the SESPA is analyzed. The result shows the feasible flight altitude of the SESPA ranges from 71.5–75 km.

This paper describes the design, fabrication, and test of a scaled-down unmanned personal air vehicle (PAV) with dual mode drive and flight capabilities. Considerations of the operational requirements of such a vehicle led to the concept of a quad-tilt-prop platform that has nacelle tilt capability with a multiple rotor configuration. Motors for both propeller propulsion and the driving mechanism were integrated into a single nacelle, and they were then implemented with a nacelle tilt mechanism for conversion between the drive and flight modes. The primary design parameters were examined through the conceptual design process, as follows: (1) overall dimensions, (2) empty weight, (3) propeller diameter, and (4) propulsion motor were determined based on design requirements. Functional tests were then performed, with test platforms designated for each development step. Comprehensive drive and flight tests demonstrated the operational capabilities of the finalized prototype. The fabricated vehicle has a takeoff weight of about 17 kg, including a 2-kg payload, an overall length of 1,000 mm, an overall width of 750 mm, and 12 min of flight endurance.

In this paper, finite-time output feedback coordinated attitude control for spacecraft formation subject to input saturation is investigated. More specifically, a finite-time observer is first designed to remove the requirement of measureable angular velocity. Rigorous proof shows the observer can achieve finite-time stability within given attractive region, which implies that the estimated values can be taken into use instead of the real angular velocity after a certain convergence time. Then, a bounded finite-time control law for the spacecraft formation is designed to render the attitude of each spacecraft to converge to reference attitude provided by the virtual leader in finite time. Finally, numerical simulation results show that the proposed schemes can guarantee the spacecraft formation achieve coordinated attitude consensus and tracking, with high accuracy and fast convergence.

A successful launch is the first step in a successful lunar exploration (LE) mission. Because spacecraft are very expensive, it is difficult to conduct sufficient flight tests to assess the launch success ratio (LSR) for the LE. Instead, the LSR needs to be comprehensively assessed using simulation tests with several large data samples and flight tests with small data samples. However, the data distribution in these different tests may not come from the same population, and the large-sample simulation test information may dominate the small-sample flight test information. Aimed at these problems, this paper proposes an LSR assessment model for the LE. The model introduces reliability measures to more accurately reflect the differences between the data populations, uses a goodness-of-fit test to check the compatibility of the simulation test data and the flight test data and to calculate the corresponding reliability, and combines the prior simulation test data distribution and the flight test data to estimate the final LSR using Bayesian statistical inference. A theoretical analysis and simulation tests illustrate how the proposed model can overcome the problem of reconciling different data sources with different populations, successfully estimate the LSR for an LE, and solve the problem of large-sample data dominating small-sample data.

In this paper, heavy rain effects on the aerodynamic efficiency of National Advisory Committee for Aeronautics (NACA) 23015 airfoil cruise and landing configurations and NACA 64210 high lift configuration have been studied. For this study, preprocessing software Gridgen has been used for creation of geometry and mesh, and fluent software is used as a solver. Discrete phase modeling (DPM) in a Langrangian frame of reference has been used to simulate the rain particles dispersed in continuous phase using two-phase flow approach. The coupling between the two phases and its impact on both phases has been included. In discrete phase model (DPM), the wall film is modeled by the injected rain particles. In a simulated rain environment all the cruise, landing, and high lift configurations of airfoils showed a significant decrease in lift and increase in drag in heavy rain environment. In this study, it is found that the heavy rain causes premature boundary-layer transition at low angle of attack (AOA) and separation at high AOA. The water film layer formed on the surface of the airfoil is thought to alter the airfoil geometry and increase the mass effectively. In the simulation for NACA 23105 the increase in drag is less contrary to simulations done for NACA 64210 high lift configuration airfoil, for which the aerodynamic efficiency degradation is much higher. The relative differences appeared to be related to the susceptibility of each airfoil to premature boundary-layer transition. Postprocessing software like MATLAB, Tec plot, and Origin are used to see the effects of the heavy rain, and the results obtained are compared with the experimental results. It is strongly believed that this study will be useful for the aviation engineers and scientists to design the airplanes and UAVs capable of flying in severe weather conditions and to train the pilots to control the airplane in heavy rain conditions well.

Spacecraft attitude and momentum management aims to slow down the accumulation of control momentum gyro (CMG) momentum through active attitude adjustments. This paper is a systematical discussion of the momentum management problem in the inertial space. The paper can be mainly divided into two parts, as follows: (1) dynamical/constraints analyses, and (2) optimal controller design. In the analysis part, the stable momentum equilibrium attitudes and the constraints for successful momentum management are presented. The analyses show that inertial oriented spacecraft with momentum management consideration is required to keep one of its principal axes of inertia perpendicular to the orbital plane. Due to the uncertainty of the inertia property, the requirement of absolute perpendicularity is difficult to obtain. To solve this problem, the second part of the paper presents an optimal momentum management controller, in which CMG momentum is feedback controlled to avoid momentum accumulation. Theoretical analyses and simulation results show that the CMG momentum oscillates periodically under the impact of gravitational and aerodynamic torque with no significant growth in magnitude when managed in the inertial space.

In this paper a novel model predictive control (MPC) approach is proposed based on mixed-$H_2/H_{\infty }$ control for space teleoperation systems with unknown, large time-varying delays and input constraints. This novel approach only measures the online delay; it does not presume the delay bounds. In addition, $H_{\infty }$ control has been incorporated into the original MPC, which compensates for large time-varying delays, handles input constraints, and provides the desired tracking performance. First, the space teleoperation system is built based on a type of control architecture and then converted into a discrete state-space equation. Next, a state-feedback controller is designed based on linear matrix inequality (LMI). Meanwhile, the corresponding sufficient conditions are derived for the controller, enabling the closed-loop system to be asymptotically stable and guaranteeing the prescribed mixed-$H_2/H_{\infty }$ performance under input constraints. The parameters of the controller are calculated online by updating the mixed-$H_2/H_{\infty }$ optimizing problem during each interval. Finally, comparative simulations reveal that the proposed approach can achieve better performance in terms of tracking ability than that of traditional MPC. Moreover, it can ensure the input constraints in the unknown, large time-varying delay scenario whereas the original $H_{\infty }$ control approach cannot guarantee performance.

An analytical study is conducted to present thermal radiation, Dufour, and Soret effects on unsteady viscous flow over a contracting cylinder. The coupled nonlinear partial differential equations are transformed into a system of coupled nonlinear ordinary differential equations by using a suitable similarity transformation. The homotopy analysis method (HAM) and HAM with a nonhomogeneous term are employed to obtain analytical solutions for the system of coupled nonlinear ordinary differential equations. A significant reduction in the averaged square residual error is obtained when the nonhomogeneous term is introduced. A comparison between analytical and numerical solutions is presented for validation. The effects of various emerging parameters on flow variables are discussed. It is found that the temperature distribution increases with an increase in Dufour number, but decreases with an increase in Soret number. The concentration distribution decreases for a given increase in the Dufour number, but increases with an increase in Soret number.

Physically based, multilayered zigzag models have the merit of providing accurate stress predictions with few fixed unknowns, but unfortunately they involve derivatives of functional degrees of freedom (e.g., $u_{,x}$, $w_{,x}$, $w_{,xxx}$, and so on). Here, a variable kinematics zigzag model with up to third-order derivatives always giving accurate predictions is considered in order to apply a technique that enables the conversion of derivatives of the unknowns. Such technique relies on energy updating concepts for finding closed form, modified expressions of displacements constituting the equivalent $C^0$ model. From this model, a simple quadrilateral plate element with standard $C^0$ interpolation functions is developed and tested. The first objective is finding a priori modified displacement fields that make the $C^0$ zigzag model equivalent from the energy standpoint to its initial counterpart containing derivatives. Then a finite element is developed, whose structural model is equivalent from the energy standpoint to its counterpart with derivatives as variables. Comparisons with three-dimensional (3D) elasticity solutions of benchmark test cases and to an experiment show that the present element obtains accurate results.

This paper presents a sliding mode control scheme on for autonomous spacecraft formation flying via a virtual leader state trajectory. The configuration space for a spacecraft is the Lie group SE(3), which is the set of positions and orientations of the rigid spacecraft in three-dimensional (3D) Euclidean space. A virtual leader trajectory, in the form of natural attitude and translational (orbital) motion of a satellite, is generated offline. Each spacecraft tracks a desired relative configuration with respect to the virtual leader in a decentralized and autonomous manner to achieve the desired formation. The relative configuration between each spacecraft and the virtual leader is described in terms of exponential coordinates on SE(3). A new feedback control scheme is proposed for coupled translational and rotational maneuver using a new sliding surface, which is defined as the exponential coordinates and velocity tracking errors such that the sliding surface converges to zero without explicit reference states, unlike conventional control schemes where explicit reference states are used. The asymptotic convergence of the sliding surface and tracking errors are guaranteed, in the presence of unknown external disturbances and uncertain system parameters, and an auxiliary signal is generated to deal with control saturation. The convergence and stability of the closed-loop system are proven within a Lyapunov framework. Finally, numerical simulations are presented to validate the developed control scheme.

This paper considers the finite-time control problem of 6DOF spacecraft formation flying (SFF) systems in the presence of external disturbances. First of all, based on nonsingular terminal sliding mode (NTSM) method, a local finite-time control law is proposed. Then a finite-time disturbance observer (FTDOB) is introduced to estimate the disturbances, and a composite control law which consists of a feedback control law based on NTSM method and a feed-forward compensation term based on FTDOB technique is constructed such that the follower spacecraft can track the desired trajectory in finite time. The major merit of the proposed method is that chattering is substantially reduced because the switching gain of the discontinuous control law is only required to be greater than the bound of disturbance estimation error rather than that of the disturbances. Furthermore, based on switching control method, continuous NTSM and FTDOB technique, a global finite-time control law is developed for 6DOF SFF systems. Finally, the effectiveness of the proposed methods are validated by numerical simulations.

The mixing efficiency of the fuel and the supersonic freestream has a great impact on the operation of the scramjet engine because of the very short residence time. In the current study, the cantilevered ramp injector was used to promote mixing of the fuel and the supersonic flow, and the injector’s configuration was optimized by the surrogate-assisted evolutionary algorithm—namely, the nondominated sorting genetic algorithm coupled with the Kriging surrogate model. In the optimization process, the swept angle, the ramp angle, the compression angle, the length of the step, and the diameter of the injector were considered as the design variables, and the mixing efficiency, the total pressure recovery efficiency, and the drag force were taken as the objective functions. Finally, the Pareto front for the multiobjective design optimization results was obtained, and the relationship between the objective functions was explored. The results show that optimized drag force increases with increased mixing efficiency and total pressure recovery efficiency, and that there must be a compromise between objective functions. The larger swept angle promotes the mixing of the fuel and the supersonic airstream, and the smaller diameter of the injector is beneficial for reducing the drag force and improving both mixing efficiency and total pressure recovery efficiency.

Predictive computational modeling of the response of armor systems to dynamic threats such as blast and impact requires understanding and quantification of the behavior of the armor materials. This paper describes the mechanical characterization of Dyneema HB26. The in-plane tensile, compressive, and shear stress-strain behavior and strength of the laminate at low rates has been determined experimentally. The tensile behavior of the Dyneema SK76 fibers, which comprise 83% of the laminate has been determined, including the effect of temperature and rate.

Wear is defined as “the removal of material volume through some mechanical process between two surfaces.” This paper focuses on the steps necessary to evaluate wear due to dry sliding between two surfaces at high velocity. Currently, there is a need to identify and compare materials that can endure sliding wear under severe conditions such as high velocities. The high costs associated with the field experimentation of systems subject to high-speed sliding have prevented the collection of the necessary data required to characterize this phenomenon. Simulating wear through finite elements (FEs) would enable its prediction and would reduce experimentation. In the aerospace, automotive, and weapon industries, such a model can aid in material selection, design, and/or testing of systems subjected to wear.

Arbitrary shapes of engineering objects demand spatial discretization in design analysis. Therein, an accurate estimation of energy-like integrals, within each volume subset, becomes essential to formulate a variety of approximate solution procedures. Higher numerical accuracy in finite element methods, especially in evaluations of system matrices, is incessantly sought to make commercial computer programs more dependable. Three-dimensional thermo-viscoplastic stress analyses motivated this research. In addition, applications in optimization can be cited. The divergence theorem of vector calculus provided the algorithm that leads to better accuracy as compared with conventional numerical quadratures.

In this paper, a continuum theory capable of describing the deformation of a cracked body based on Castigliano’s second theorem is presented. The additional deformation due to the crack presence is described by the concept of stress intensity factor (SIF) of linear elastic fracture mechanics. Both the crack opening distance and the body deformation can be easily computed for any arbitrary loading on any boundary. As a demonstration, the proposed theory is applied to cases of edge-cracked infinite plane and edge-cracked beam, and the results show high accuracy when compared to those predicted by the numerical finite-element method.