Journal of Aerospace Engineering

Due to the importance of the ballistic trajectory computation including the estimation of launch point and impact point problems, a fast and accurate model is needed to increase the weapon precession. A discrete time transfer matrix method is developed and presented in this paper to compute the trajectory for both fin- and spin-stabilized artillery projectiles. This method is based on discretizing the projectile continuous nonlinear system into a set of linear equations using a second-order Taylor series. These equations are collected into one transfer matrix which can be used in trajectory prediction with high computation speed by evaluating a set of algebraic equations instead of integrating differential equations. Results demonstrate that the presented formulation agrees very well with those obtained numerically using the well-known nonlinear 6-degree-of-freedom model.

The geotechnical properties of Martian soils are critical parameters in predicting and simulating soil behavior with regard to vehicle performance on Mars. In preparation for manned or robotic missions to Mars, surface vehicles must be tested on terrains that represent the mechanical characteristics of the Martian ground. This paper presents the development of a lightweight simulant and its preparation method to emulate the mechanical properties of Martian soil for high sinkage mobility tests. A geotechnical testing program was developed to measure specific gravity, particle size distribution, bulk density, compression indices and shear strength. The simulant can achieve the typical Martian regolith density range, which is approximately 38% of that on earth. This is of particular importance because strength parameters of granular materials, which characterize the plastic behavior of soil samples in sinkage tests, are controlled by the effective confining pressure, which itself is induced by gravity.

This paper presents a numerical model for investigation of aeroelastic behavior of an arbitrary-shaped panel with arbitrary boundary conditions under the effect of supersonic external flow and internal liquid sloshing simultaneously. A finite-element model is used to describe the structural dynamics along with the linear piston theory for describing the supersonic external flow effects. Also a potential flow model discretized by the boundary element method is used for describing the internal flow field. Using modal analysis technique a coupled fluid-solid interaction reduced-order model is developed for the system. The developed model is used for investigation of dynamic stability boundaries of the system for different conditions of internal fluid, and new dynamic behaviors is observed in interaction of structure with internal and external flow simultaneously.

The horizontal stabilizer is primarily assembled by edges and ribs. Edges and ribs are the typical thin-walled deformable aluminum components of the horizontal stabilizer with different manufacturing distortions. Part-to-part assembly of edges and ribs regularly cause difficulties associated with dimensional variations. It is time-consuming and impossible to construct one finite element analysis (FEA) model to find the propagation relationship from the edges’ and ribs’ distortion to the trailing edge’s deformation. The paper proposes a method to analyze deformation in horizontal stabilizer assembly using FEA and multilevel analysis (MLA) method. FEA method is employed to analyze the deformation induced by edges’ and ribs’ distortion, respectively, because of their different structural features.MLA connects edge’s and rib’s FEA models with orthogonal design of experiments, and accurate analysis in the interactive influence between edge’s distortion and rib’s distortion is obtained. It provides the bases for achieving high-precision calculation results with sequential small computational models. It will enhance the understanding of the compliant components deformation in assembly, and help systematically improve the precision control efficiency in civil aircraft assembly.

This study examines the series solution for magnetohydrodynamics (MHD) three-dimensional flow past a porous shrinking sheet. The thermal radiation effects are also taken into account. The model used for the nanofluid incorporated the effect of Brownian motion and the thermophoreses. The involved equations are solved analytically by homotopy analysis method (HAM). Graphical results for the dimensionless velocity, temperature, and nanoparticle fraction are reported and examined for various physical parameters showing the interesting aspects of the obtained solutions. The numerical table for local Nusselt and Sherwood numbers are also computed.

The minimum-time coplanar interception problem with an upper-bounded tangent impulse for a nonmaneuvering target is studied. This problem requires the same flight time for two spacecraft and a transfer orbit tangent to the initial orbit. For a given impulse point, the transfer time for any conic orbit is a function only of the true anomaly of the target orbit; then the minimum-time orbit is determined by comparing all feasible solutions that are obtained by the secant method in the range of solution existence. Moreover, the true-anomaly range of the target orbit is derived for an upper bound on the magnitude of the tangent impulse. Finally, the global minimum-time solution with initial coasting time is obtained by numerical optimization algorithms. An example is given to apply the proposed technique for all the minimum revolution-number solutions at different impulse points and the global minimum-time solution.

Based on the nonlinear stable inversion approach, an exact output tracking control law is designed for flexible air-breathing hypersonic vehicles (FAHVs). This problem is challenging because of the inherent couplings between the propulsion system, the airframe dynamics, and the presence of strong flexibility effects. Because of the complex nonlinear dynamics of the vehicles, a control-oriented model, which can retain the dominant features of the higher-fidelity model, is adopted for the control design. With respect to velocity and altitude as the regulated outputs, a partially linearized model of FAHVs is obtained through input/output linearization. Then the internal dynamics of FAHVs model are constructed and the stability of the zero dynamics are discussed. Based on the noncausal stable inversion approach, the ideal internal dynamics of FAHVs are computed. A state tracking model is derived from the output tracking problem, and an optimal controller is designed. The stability of the closed-loop system, including the unstable internal dynamics, is guaranteed by the designed control law. Finally, simulations are given to show the effectiveness of the proposed control method.

A dynamic model showing the relationship between backpressure and shock-train leading edge (STLE) is necessary in the development of a STLE automatic control algorithm. In this research, a nonlinear dynamic model based on equilibrium manifold expansion is presented to locate the STLE position in a large range of operation conditions. The initially started flow field of a hypersonic inlet at different throttling ratios was experimentally studied at freestream Mach numbers 4.5–6.0. The nominal STLE position was defined by the pressure ratio method. The variation of STLE position with backpressure and freestream Mach number was then discussed. Testing with the experimental data, past models could only be used locally, whereas for large range of operation conditions they lose effectiveness. Therefore, the nonlinear dynamic modeling method based on equilibrium manifold is introduced to solve this problem, the simulation results, and comparisons at large range of operation conditions show its validity.

A new method is presented to improve the safety and reliability of an aeroengine rotor system (AERS) in early fault recognition and health monitoring, addressing the problems that fault samples are not sufficient and that early weak faults are not easy to recognize. First, the stochastic resonance system is used to refine the early weak feature signal so as to amplify the fault information. Second, the early fault features are extracted by multiresolution performance of wavelet packet analysis, and the fault characteristic vector can be constructed. Finally, the extracted eigenvectors import the support vector machine (SVM) classifier to carry on the fault recognition and then make use of intelligent monitor module to monitor early faults in AERS. Experimental results have shown that this method can not only early recognize faults in AERS but also monitor the fault online. It provides a new way to increase the safety of AERS and predict the sudden fault.

This study considers a satellite orbit design when a given set of target sites needs to be visited within an assigned period. Optimization of the orbit design takes into account target properties such as priority and minimum duration of observation time. Equations describing satellite orbit are used to compute the satellite’s field of view and ground track. Scheduling of visits to the targets is carried out by maximizing the observation time over priority-weighted targets and minimizing the distance between the ground track and the target position while satisfying the constraints with respect to the minimum duration of observation time and revisit requirement. A self-adaptive differential evolution (SA-DE) algorithm with the ability to adjust its search parameters is used to find the best solution. The proposed design methodology computes the optimal values of design variables and also the start and end observation time points of each target. Regional and global case studies demonstrate that the SA-DE/best/1/bin variant is found to be the most suitable optimizer for orbit design.

Aerothermodynamics analysis of the Orion crew exploration vehicle (CEV) reentry vehicle at high altitude has been studied numerically. At high altitude in the rarefied flow regime noncontinuum technique such as direct simulation Monte Carlo (DSMC) method is used to solve the Boltzmann equation of kinetic theory. Results are reported at different altitudes in the rarefied regime for ideal gas and real gas model. The effects of nose radius and free stream velocity are presented. The estimation of bow shock strength and peak surface heat flux is higher for ideal gas compared with real gas model. The flow field characteristics have shown significant change with decrease in altitude and increase in free stream velocity. Present results are matching well with the results available in the literature.

Coning motion has been detected in flight experiments of many spinning projectiles for decades and the stability analysis of this motion has been extensively studied recently. In this paper, a novel optimal control algorithm based on stability analyses is developed to suppress the coning motion. A Lagrangian functional is built to compute the optimal control that minimizes the coning motion for a given control cost. Then a linearized approach, which produces accurate enough control within the range of parameters considered, is proposed to reduce the computational cost. It is analytically presented that the optimal solution of this linearized optimization is unique at any given values of control cost. The analytical analysis also indicates that there exists a direct solution of the control that approaches the linearized optimal control when the control cost is small enough and that this direct solution can be obtained much more efficiently than the linearized optimal control. The controllability problem associated with this optimal control algorithm is also discussed and the results shed lights on the choice of numerical parameters in the optimization. These optimal control algorithms are implemented to calculate the optimal control that minimizes the coning motion induced by initial disturbances to a wrap-around-fin missile.

To develop the high-performance high-reliability of aeroengine, the probabilistic design of high-pressure turbine (HPT) blade-tip radial running clearance (BTRRC) with three objects (disk, blade, and casing) and two disciplines (heat and mechanical loads) was completed based on distributed collaborative response surface method (DCRSM) from a probabilistic perspective considering dynamic loads and nonlinear material properties. The mathematical model of DCRSM was established based on the quadratic response surface function. The basic idea of BTRRC probabilistic design based on DCRSM was introduced. The BTRRC probabilistic analysis results, consisting of the probabilistic distribution characteristics of input-output variables, the failure probability, and reliability of BTRRC under different static blade-tip clearances $\delta$ and the major factors affecting the BTRRC, reveal that the static blade-tip clearance $\delta =1.865\times {10}^{-3}\mathrm{m}$ is an optimally acceptable option for BTRRC design. The comparison of methods shows that the DCRSM has high accuracy and high efficiency in the BTRRC probabilistic analysis; meanwhile, the strengths become more obvious with the increasing times of simulations. The present research offers an effective way for HPT BTRRC probabilistic design, as well as provides a promising method for the further probabilistic optimal design of complex mechanical system.

Helicopter blades are known to experience complicated unsteady flow conditions during high-speed forward flight. Transonic shock takes place along the advancing side, which decreases the performance of the blades and results in impulsive noise. Adaptive airfoils seem promising for increasing performance at considerably variable flow conditions. In this paper, a blade with a passively morphing layer is presented for transonic shock/boundary layer control at forward flight conditions. The morphing layer embedded in the upper surface of the blade can form an adaptive bump, which in turn can control transonic shock/boundary layer interaction. A fluid-structure-interaction (FSI) enabled simulator is employed to study the performance of the morphing layer at two different forward conditions. The results demonstrate that the adaptive bump formed by the morphing layer could efficiently weaken the strong normal shock over the layer by producing a λ-structural shock. Moreover, the morphing layer retreats to its undeformed state to prevent detrimental effects at weaker shock or shock-free flow. To further increase the performance of the morphing layer, a morphing layer that is inclined with respect to the spanwise direction is used to replace the layer that lies parallel to the spanwise direction. The results indicate that the inclined morphing layer performs better while the shock position takes on a less uniform distribution.

Optimization studies for the next generation of aircraft and engine families are typically associated with the addition of technological advances to improve fuel efficiency, noise levels, and overall aircraft economics. In the conceptual study phase, the optimization process must address the uncertainties related to the anticipation of market requirements and potential technological advances at the time of entry into service, along with the fidelity of the modeling applied in the analysis. The current paper presents a framework using a robust design optimization methodology to mitigate these uncertainties when an airframe and its engines are designed jointly. The optimization of a 70 passenger aircraft illustrates the approach with a formulation seeking to optimize the design while attenuating its sensitivity to the potentially adverse effects of uncertainties.

Flight dynamics deals principally with the natural response of the aircraft to longitudinal perturbations that typically consist of two underdamped oscillatory modes having rather different time scales. One of the modes has a relatively short period and is usually quite heavily damped. This is called the short period mode. The other mode has a much longer period and is rather lightly damped. This is called the phugoid mode. The required degree of dynamic stability is specified by the time it takes the motion to damp to half of its initial amplitude. The dynamic characteristics of an airplane is important in assessing its handling or flying qualities and for designing autopilots. This paper investigates the dynamic characteristics of short take off and landing (STOL) airplanes employing root locus technique and state space approach. The nonlinear differential equation of motion of an airplane developed from Newton’s second law of motion was linearized using the small disturbance theory. Different phugoid approximations for longitudinal modes were developed and compared with the exact solution formulated using statistical software. Further research work about active control technology to improve aerodynamic efficiency which results in potential fuel savings should be carried out.