Journal of Aerospace Engineering

This paper considers an aeroelastic two-dimensional wing-flap system with uncertainty and external disturbance. Nonlinear stiffness in pitch is expressed as the second-order polynomial form. In the uncontrolled aeroelastic model, limit cycle oscillation (LCO) appears at postflutter speed. By utilizing leading-edge and trailing-edge control surfaces as the control input, a sliding mode control scheme using varying boundary layers based on backstepping design is developed to suppress aeroelastic vibration. The proposed control strategy combines the advantages of both sliding mode control with varying boundary layers and backstepping technique. In contrast to the conventional sliding mode control design, the proposed method not only avoids chattering but also obtains stronger robustness. The results of numerical simulation clearly demonstrate the effectiveness of the proposed control strategy for aeroelastic vibration suppression under saturated controller input.

This paper focused on improving a small-scaled helicopter rotor blade prototype with a flap-driving mechanism termed the Seoul National University Flap (SNUF). The design of the SNUF included realizing vibratory load reduction. First, a multibody structural dynamics analysis was performed to determine the influence of the flap dimension and location within the rotor blade with respect to hub vibratory load reduction. This process was followed by selecting a specific blade configuration that maximized vibration reduction capability. Following this, a numerical optimization technique was applied to improve the cross sectional design of the SNUF blade. The design optimization procedure obtained improved blade sectional design with decreased first torsional frequency and reduced blade weight, and it satisfied sufficient structural integrity. Three-dimensional nonlinear static structural analysis also was performed for the optimized SNUF design. The von Mises stress distribution on the blade and components was predicted by considering external aerodynamic loads, centrifugal loads because of rotation, and contact among the internal components. Future studies will include fabricating the prototype blade on the basis of the optimized design and performing a whirl tower test.

In this paper, first the mathematical models of four major environmental disturbance torque components and corresponding bound equations are presented. Then the effect of the inertia matrix uncertainty on rigid satellite’s attitude dynamics is defined as an external torque in the derived state equation, which together with environmental torques forms the so-called lumped disturbance torque. After obtaining the complete equation for the model uncertainty-induced torque vector, its exact bound equation is derived by using matrix-vector norm relations. To validate the significance of these preliminary results for use in robust attitude controller design, a new modification of the classical sliding mode attitude controller present in literature is proposed, which is the primary contribution of this paper. The new design that is based on comprehensive knowledge of the lumped disturbance’s bounded variation leads to a decision rule on the switching control gain that is not excessively conservative. After verifying the accuracy of the bound equations in a simulation under no control, a second simulation is carried out with control input from the designed sliding mode controller to show that the proposed design works. The superiority of the new design is discussed in comparison with another design from literature that does not exploit the complete model of the inertia matrix uncertainty-induced torque through a comparative simulation’s result. The conclusion is that the modified controller design results in an attitude control system that has guaranteed robust stability in addition to reasonable conservativeness thanks to the newly obtained comprehensive decision rule on the switching control gain.

Lunar exploration, which slowed in the 21st century after the Apollo program, has seen more activity recently with the participation of Asian countries such as Japan, China, and India. Because lunar modules and rovers cannot be tested directly on lunar soil, these countries have developed lunar simulants. Simulating lunar soil is difficult and expensive because its formation mechanism and geotechnical behavior are comprehensively different from those of the terrestrial soil. Johnson Space Center Number One (JSC-1) and Johnson Space Center Number One A (JSC-1A), developed by the National Aeronautics and Space Administration, are the most widely used simulants. Korea has yet to succeed in developing a lunar simulant that meets international standards. The authors perform basic research on lunar simulant development based on basalt samples having similar chemical and mechanical properties to those of lunar soil, with reference to lunar soil data reported under the Apollo program. The resulting prototype is named Korea Lunar Stimulant—Type 1 (KLS-1). Compared with other lunar simulants [JSC-1 and Fuji Japanese Simulant 1 (FJS-1)], KLS-1 shows promise in terms of affordability and similarity to real lunar soil. As such, it is expected to find a wide variety of applications not only in space development projects but also in international research.

In this paper, damage monitoring and fatigue-life prediction of cross-ply silicon carbide fiber–reinforced calcium aluminosilicate (SiC/CAS) ceramic-matrix composites (CMCs) at room temperature and elevated temperatures of 700 and 850°C in an air atmosphere have been investigated. Considering the fatigue-damage mechanisms of interface wear, interface oxidation, and fiber oxidation, the relationships among fatigue hysteresis-loop evolution, interface shear-stress degradation, broken fibers, and fatigue lifetime have been established. The evolution of damage parameters, i.e., fatigue hysteresis dissipated-energy changing rate, interface shear-stress degradation rate, broken-fiber fraction, and fatigue-life stress-cycle (S-N) curves have been analyzed. The interface shear-stress degradation rate is affected by the maximum stress level and testing environment; the fatigue hysteresis dissipated energy decreases with the applied increment cycles and experimental temperatures. The fatigue limit stresses are 50, 36, and 30% of tensile strength at room temperature and 700 and 850°C, respectively.

The definition of a set of precursory safety metrics is critical to detect when an airspace is degrading in terms of safety and thus undesired effects are becoming more likely. Furthermore, safety metrics are paramount in the measurement of the impact of new operational procedures or technical improvements in the air traffic control system. The study presented in this paper introduces three safety metrics (reaction time performance indicator, time to closest point of approach performance indicator, and time to closest point of approach critical limit ratio) derived from a given airspace and a sizable, assorted traffic sample extracted from traffic surveillance track data. The metrics are used to characterize the airspace as a function of the safety outcome, which can be continuously overseen. The final goal of the safety metrics is to be used as an airspace safety warning system, where precursory metrics would signal the need to act to maintain the air traffic control system safety target in the face of operational, organizational, technical, or legal changes.