Journal of Aerospace Mechanics
Year:2019 | Volume:15 | Issue:4 (58)|Papers Count:8

In the investigation, the vibrations of a rotating functionally graded cylindrical shell under axial and internal pressure with ring and stringer stiffened and simply supported boundary condition based on Love's shell theory is studied. The cylindrical shell with ring and stringer stiffened widely used in structures such as rockets, submarines and fuel tank of aircraft. The material properties of the rotating functionally graded (FG) cylindrical shell vary continuously across the thickness according to the power law distribution. In this study considers functionally graded material composed of Nickel and stainless steel, in which FG cylindrical shell has Nickel on its inner surface and stainless steel on its outer surface. The governing equations of a cylindrical shell are derived using Hamilton's principle and energy method and the effect of various parameters such as angular speed, ring and stringer stiffened, axial load, internal pressure and the functionally graded material are investigated. The validity of the results by comparing them with the results of previous research is investigated, in which there is a very good agreement between the results of the present work and previous studies.

Condition Assessment is one of the most significant techniques of the equipment health management. PHM cycle is a developed form of Condition Based Maintenance (CBM). Condition Assessment is the most important step of this cycle. In this paper, a process is presented in which by introducing a new feature for vibration, simulating and predicting it using a autoregressive Markov regime switching model, and using a new approach for combining condition monitoring sensor information based on fuzzy c-means clustering and Dempster-Shafer theory, the degradation state is determined and remaining useful life is estimated. In order to evaluate the model, sensor data for PHM Data Challenge 2012 have been used to forecast the bearing's Remaining Useful Life and the results of the study have been compared with the winning results. The results obtained from the comparison show the competitiveness of the proposed model with the winner of challenge.

In an aircraft system that is in the rectilinear steady flight state ready for a longitudinal maneuver, instability, model uncertainty, and disturbances could make the maneuver impossible without the help of a controller. Considering that the dynamical equations of the aircraft are nonlinear, using a linear controller designated based on linearization at the equilibrium of the point may not be entirely efficient. Therefore, using non-linear controllers is of utmost importance. The present study aimed to design a controller using the gain scheduling method. Designing the gain scheduling-based controller was performed in two stages. Initially, linear controllers were developed, which were the controlling components of gain scheduling. The second stage involved planning in order to determine the criteria based on which one of the linear controllers was in the closed loop. The required linear controllers were selected through optimal control methods. In order to verify the number and location of the linearized points and criteria for using the linear controllers, the concepts of stability margin and v-gap metric were applied. Furthermore, longitudinal maneuver, which is requested as a command or a base signal in the rectilinear steady flight state, was considered as the rotational constant value for the Pitch angle. According to the results, combining the nonlinear gain scheduling controller and a nonlinear system in a closed loop could contribute to the desired longitudinal maneuver.

Modeling and control of Unmanned Aerial Vehicle is undoubtedly one of the most active areas in the field of control engineering. The characteristics of aerial vehicles such as nonlinear dynamics, time variability and the structural and parametric uncertainties make flight control issue relatively a complex and important subject in their design. One of the widely used methods in this area is sliding mode control technique. This paper proposes the design, the simulation, and the analysis of two controllers namely PID and sliding mode in order to optimally control the pitch angle of an unmanned aircraft. Initially, in order to dynamically model the system, an appropriate mathematical model is presented to describe the longitudinal motion of aircraft. Subsequently, a robust sliding mode controller is designed using particle swarm optimization (PSO) algorithm to control the pitch angle of an aircraft against the uncertainties and the external disturbances. The optimal parameters are determined by minimizing both the control signal and tracking error. Then, the performance quality of the proposed method was compared with the results of the optimal PID controller in terms of transient response characteristics. Finally, simulation results indicated that the proposed sliding mode controller can reduce the overshoot of system response and yield more robust performance than PID for the control of pitch angle.

In this paper, design and mathematical simulation of a hybrid piezoelectric capacitive-resistive nano accelerometer sensor has been investigated using Comsol multiphysics finite element software. The governing equation of motion in the transverse direction for a cantilever Euler-Bernoulli beam has been used considering electrostatic and van der Waals-Casimir forces. The effect of existence of piezoelectric layer, piezoresistive and capacitive patches in the bending behavior of the sensor is considered applying the physical properties of piezopatches and then simulated and solved using mathematical modules of the Comsol software. The obtained results are included the values of ∆ R/R, ∆ V/V and ∆ C/C for piezo patches with respect to the variation of applied acceleration of the sensor, stress distribution along the length of the beam and maximum value of von Mises stress to compare with the yield stress of the sensor, deflction of the beam with respect to the variation of applied acceleration of the sensor, fatigue analysis for the sensor and the sensor's mode shapes and natural frequencies to determine the performance of the sensor. The effects of changes of different values including nano g acceleration and material properties of piezoresistive patch on the behavior of the sensor are also considered in the analyses. It can be observed that nansensor has been satisfctory well designed from measuring sensitivity, stress level and fatigue life point of views.

Shock tube is an equipment in which by creating a pressure difference between driver and driven section via the bursting membrane has the ability to generate shock wave with very short rise time. One of the important parameters in the shock tube is the planar shock wave and the distance of its formation along the driven section. In this study, the shock wave pressure has measured at different sections along the shock tube as well as at different radial distances, using three piezoresistive pressure sensors. Experiments were repeated with three different thickness of diaphragms 0. 1, 0. 2 and 0. 3 mm. Diaphragms were made of Mylar. The results of the experiments were extracted using TRAww software, which is a software for signal processing of the pressure sensors; and the distance of the planar shock wave for different diaphragms was obtained. The results show that by increasing the diaphragm thickness and thus increasing the explosion pressure (pressure of the driver area), the shock wave pressure increased and the planar shock wave propagates further away in the driven section. The uniform duration of the shock wave using a diaphragm with a thickness of 0. 1 mm is smaller than the other two diaphragms, and the planar shock wave is not stable until the end of the shock tube. Also, the pressure drop in driven section after rupture of the diaphragm increases with increasing diaphragm thickness.

At the present paper, thermodynamic simulation of a twin spool separate exhaust turbofan engine, at the on design and off design conditions is performed. At this engine, a part of air flow is extracted from the high pressure compressor in order to cool the turbine blades and provide cockpit air. Also, power is extracted from the fan to provide the power for accessories. For the validation of developed thermodynamic model, the results are compared with CFM56-7B engine experimental outputs including: corrected low pressure spool speed, engine thrust, fuel flow rate, high pressure spool speed, and exhaust gas temperature of engine. Combination of engine revolution equations with thermodynamic cycle equations and preparing the results in the frame work of generalized thrust curves, which are the innovations of the current study, demonstrate that the model has 12% error at most and can predict the off design behavior of the engine correctly. It is obtained that at the cruise condition, the bypass duct exit and core stream exit are chocked. The results show that in the case of inlet Mach number increasing, the bypass duct chocking sensitivity is more than core stream chocking sensitivity. Moreover, positive ISA condition decreases the thrust and specific fuel consumption and negative ISA condition increases the thrust and specific fuel consumption.

In this paper, a LEO to GEO transfer orbit is presented using the invariant dynamics of the Restricted Three Body Problem (RTBP) and the Lagrangian equilibrium points of the Earth-Sun system in the presence of disturbances. After launching the spacecraft to the LEO, its incidence angle is at least equal to the latitude of the launch site, so to transfer it to the GEO, it should be able to bring the spacecraft circle to zero with a large amount of energy. Here, with the help of the dynamics of the Lagrangian points, a method is proposed that minimizes the energy generated by the change in the incidence angle. For that purpose, the problem is divided into three parts. In the first stage, the appropriate initial conditions are obtained in GEO, which reaches the L1 point after the set time. The second phase is similar to the first one, with the difference that the initial conditions in the LEO are obtained in the form of forward transfer from L1. Finally, the Hallo orbit around L1 is used to connect two trajectories to achieve the integrated transfer orbit. In the following, the disturbances model is also added to the motion equations of the problem of three finite objects, in order to evaluate its effect on the results. The simulation results show that the impulse required by this method for inclination orbit greater than 29 degrees, even in the presence of disturbances, is better than the Hohmann transfer method.