This paper investigates the effects of nondimensional parameters on the characteristics of synthetic jets. Flow inside the synthetic jet cavity and orifice, the flow outside are simulated together using large-eddy simulations (LES). A comparison of the present results and those of the experiment shows that simulating the flow both inside and outside the jet cavity is essential for accurately estimating the velocity and velocity fluctuations of the synthetic jet. LES results under various flow conditions show that strong three-dimensional vortices are generated when the Reynolds number is large, but finer vortex structures form because of stronger vortex interaction as the Strouhal number increases.

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This paper presents dynamic and static instability characteristics of a free-free multi-stepped Timoshenko beam subjected to a follower force, where each step has a point mass with transverse and rotary inertia. A finite element code is developed to model two-dimensional multi-stepped Timoshenko beams using the extended Hamilton’s principle. The effects of each section parameters on the type of instability and critical follower force have been studied for different models of multi-stepped beams. It is concluded that the assumption of a uniform beam to study most aerospace structures such as launch vehicles could lead to an unrealistic instability analysis, where any small change in the characteristics parameters for each step could change the instability behavior significantly. Furthermore, this paper introduces a new parametric approach concept for analyzing the characteristics effects of a multi-step aerospace structure.

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An unsteady panel method based on potential flow theory is developed for unsteady aerodynamic analysis of helicopter rotors. The panel method uses a piecewise constant source and doublet singularities in the solution. This potential-based panel method is based on the Dirichlet boundary condition coupled with a time-stepping method. The proposed method uses a time-stepping loop to simulate the unsteady motion of helicopter rotors. A free wake model is used for the wake simulation of a helicopter rotor. The proposed method is validated by comparing our results with experimental data. The incompressible results of the proposed method for helicopter rotors are in good agreement with the experimental data for hovering and forward flight. The proposed method can be used for practical applications such as aerodynamic designs and the analysis of various helicopter rotor configurations.

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In this paper, we propose a sliding-mode controller for a mini unmanned aerial vehicle (UAV) with propellers to follow the predetermined attitude trajectory. First the vehicle attitude dynamic model is established for angular displacements and for angular velocities, respectively. Next a sliding-mode controller with a switching surface is designed to eliminate uncertainties and disturbances. Then the attitude tracking control of a mini-UAV using the proposed control strategy is illustrated while flight. Finally, we employ the Lyapunov stability theory to fulfill the stability analysis of the proposed controller for the overall nonlinear control system. Extensive simulation results are gained to validate the effectiveness of the proposed sliding-mode controller.

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Spectroscopic measurement is conducted using a free-piston double-diaphragm shock tube to investigate the nonequilibrium phenomena in the shock layer. The shock velocity and spectrum position correlated to the shock front are determined using a double-laser schlieren measurement system for precise localization. Emission spectra of N₂(1+), N₂(2+) and N₂⁺(1−) band systems are obtained by means of time-frozen imaging spectroscopy. A spectrum fitting method is used to determine the rotational and vibrational temperatures from the measured spectra, and temperature distribution correlated to the shock front is finally obtained. The measured rotational temperatures are in high nonequilibrium with the translational temperature expected from the numerical prediction of the two-temperature model. The measured rotational temperature for N₂(2+) is lower than those for N₂(1+) and N₂⁺(1−) immediately behind a shock wave. Hence, rotational relaxation for the N₂C state looks slower than those for the N₂B and N₂⁺ B states. On the contrary, the measured vibrational temperatures for N₂(1+), N₂(2+) and N₂⁺(1−) are close to each other, and agree well with the numerical prediction of the two-temperature model. The experiment and numerical analysis suggest that the electronic excitation temperature is in nonequilibrium with the vibrational temperature.

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In this paper, we discover the existence of a sort of special periodic orbit around the equilibrium points of irregularly shaped asteroids and present an algorithm to accurately determine this novel orbit. Since the environment of orbital dynamics around the equilibrium points of an asteroid is similar to that in the circular restricted three-body problem (CRTBP), this article introduces the research approach used in the research of halo orbits in the CRTBP to determine the special orbits. Firstly, the dynamics of particles orbiting on irregularly shaped spinning asteroid are analyzed and then the formula to compute the positions of asteroid’s equilibrium points can be deduced from the equations obtained from the analysis. Secondly, third-order approximate analytical solutions for the periodic orbit around the equilibrium points were obtained using the Lindstedt-Poincaré method. Finally, in the high-order gravitational field model, differential corrections were applied to amend the initial orbit value obtained from third-order analytical solutions in order to obtain accurate numerical solutions of this periodic orbit, which are called aureole orbits in this paper due to their similarity to the halo orbits around the Lagrangian point in the CRTBP.

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This paper presents a propulsive performance evaluation for a H₂/Air pulse detonation engine (PDE) equipped with a converging-diverging exhaust nozzle. The evaluation was conducted using system-level modeling and multi-cycle numerical simulations. This study discusses two-dimensional and axisymmetric compressible Euler equations with a detailed chemical reaction model. First, the single-shot propulsive performance of a simplified-PDE (i.e., no exhaust nozzle) is evaluated to show the validity of the numerical and performance evaluation method. The influences of the initial conditions, ignition energy, grid resolution and scale effects on the propulsive performance are studied using multi-cycle simulations. The present results are compared with the results calculated by Ma et al. and Harris et al. The differences between their results and the present simulations are approximately 2–3% because their chemical reactions use a one-step model with a single-γ model. The effects of the specific heat ratio should be estimated for various nozzle configurations and flight conditions.

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Shock layer radiation is observed in the wavelength range from vacuum-ultraviolet (VUV) to near-infrared (NIR) using a free-piston double-diaphragm shock tube. In the VUV region, atomic lines of N and C are predominant and no molecular band spectra are seen. In the ultraviolet (UV) region, the molecular bands of N₂(2+) and N₂⁺(1−) systems are predominant. In the visible (VIS) and NIR regions, atomic lines of N are intense, and weak molecular bands of the N₂(1+) system are seen. The spectrum in the VUV region becomes more intense than that in the UV region with increasing shock velocity. On the other hand, the spectrum in the UV region is almost same with increasing shock velocity. The measured spectrum is analyzed using the radiation analysis code SPRADIAN 2. The analysis shows that the numerical spectrum can reproduce the measured one when the electronic excitation temperature is higher than the vibrational temperature. The electronic excitation temperature is evaluated from the atomic lines of nitrogen in a wide wavelength range and the spatial distribution is obtained. The result shows the electronic excitation temperature is higher than the vibrational temperature immediately behind a shock wave, consistent with the result obtained from the spectrum analysis.

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A guidance law tailored for an interceptor missile steered by aerodynamic lift and divert-thruster and with its attitude oriented by forward attitude-thruster (dual-thruster control) is proposed. The bang-bang strategy has the maximal capture zone. However, it leads to control chattering, resulting in excessive control effort. The continuous strategies have smaller capture zones, but can require a smaller control effort. The combination of bang-bang and continuous strategies may lead to a cooperative effect of the capture zone and control effort. In this paper, the guidance law is derived using a linear-quadratic differential game formulation and then modified for bounded controls. For the derivation, the interceptor closed-loop dynamics is represented by a first-order biproper transfer function for the attitude-thruster control system and a strictly proper one for the divert-thruster control system, respectively. Analysis based on the proposed guidance law shows that the divert-thruster control system has higher efficiency due to its small time constant and the direct forces that mainly affect the short-term behavior of the missiles. The parameters’ effects and the control command distribution are also studied.

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An output-feedback H_∞ preview controller is proposed. In the flight control operations performed in this study, the disturbance input is estimated by the observer, and the disturbance responses are predicted and restricted using the H_∞ preview controller. The proposed system was applied to the longitudinal dynamics of an experimental unmanned aerial vehicle (UAV), which was developed to study the flight control of rocket planes. The accuracy of the estimated disturbance and the performance of the preview controller were validated through performing flight testing operations. From the results of the flight testing operations, the proposed system based on the disturbance model of the pitching moment can be used for the flight control of the experimental UAV.

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