JOURNAL OF AERONAUTICAL ENGINEERING (JOAE)
Year:2014 | Volume:16 | Issue:1|Papers Count:7

In this paper, a dynamic model of Proton Exchange Membrane Fuel Cells (PEMFC) with hydrogen storage for partial electrical generation of C-130 Hercules has been presented. A three-phase two-level inverter is used to connect the fuel cell DC output to the AC system of the Hercules. Another solution would be the direct connection of the output DC terminal of the proposed system to the DC essential bus. The proposed method performance is simulated using MATLAB/SIMULINK. The study results indicate that the voltage level is fixed to an acceptable level and following the electrical load is properly achieved. The proposed method employs an electrolyser to generate the hydrogen and oxygen. By using the stored hydrogen in the proposed structure, the need for batteries for storage purposes would cease to exist. In addition, the oxygen produced can be used in the aircraft in case it’s required.

In this paper the problem of multidisciplinary and multi-objective conceptual design optimization of an air launched projectile (ALP) is investigated. The proposed task is performed using a three degree of freedom (DOF) flight dynamics simulation model and taking into account all the constraints involved in the optimization process. To maximize the payload weight as well as the range of ALP, the vehicle weight and balance, aerodynamics and stability disciplines are selected in the optimization process using a model with moderate levels of fidelity for each subject. The ALP design optimization problem contains 14 design variables and 2 target functions that include the payload weight and the vehicle range. Finally, a performance based comparison of results between the optimized ALP and its non-optimum initial configuration has been made. In addition, a Monte Carlo analysis is performed over the optimal ALP design to see the effects of launching uncertainties in meeting the mission requirements.

In this paper a novel optimal and robust proportional navigation (PN) guidance law for three dimensional missile-target system is proposed based on extended Kalman filter (EKF). New guidance law is a combination of PN, sliding mode and back stepping guidance laws. In this guidance law there are some coefficients that must be adjusted. Genetic algorithm (GA) is used to achieve this goal. Also, system states are estimated using extended Kalman-Bucy filter (EKBF). Finally, proposed guidance law (PN Back stepping (PNB)) is compared with augmented proportional navigation guidance (APNG) law. Simulation results indicate that PNB law has far better performance compared to APNG law.

The purpose of this article is to develop a fast algorithm applicable in trajectory planning for UAVs. The proposed new algorithm plans a smooth path with continuous curvature through the predefined waypoints in 2-D space. This algorithm forms a G2 continuous trajectory Dubins-based path that replaces the Dubins line with Bezier curve with length close to an optimum. Continuous curvature property in this new trajectory, unlike the Dubins path, prevents the sudden change in lateral acceleration in an airplane that is tracking this trajectory. Obtained algorithm is based on analytical solutions so it could have real time applications and hence, it could be capable of re-planning the trajectory during flight for the guidance of UAVs in any kind of mission.

One of the difficulties toward analyzing complex structures is loading phenomena and determining the constraints due to the boundary conditions. Although in aerodynamic loading, applying distributed load is an appropriate approach, in case of being simultaneous with inertia loading, the simulation process can be extremely difficult and near impossible. This article is an attempt to introduce a new method for simultaneous aerodynamic and inertia loading on aircraft finite element model using genetic and mapping approach. To this end, the equivalent aerodynamic and inertia load have been regarded as an unspecified pressure distribution. Furthermore, the surfaces under study were divided into bands in which the amount of pressure was considered as the optimizing variable. The objective function has been chosen in a manner so that it could satisfy the required lift and drag forces as well as the torque resulted from these forces. One of the merits of this method is that the pressure would be distributed on the surface geometry, and even in case of change in the elements, there is no need to reapply the loading. To ensure the validity of this method, we conducted it on the finite element of the wing and the fuselage of a sample aircraft which led to satisfactory results. The results confirm the accuracy and efficiency of this method.

The Plasma Actuators (Dielectric Barrier Discharges) are strong tools for active flow and boundary layer control that cause to postpone the location of separation point on control surfaces which recently have been a focus point for aerospace scientists and engineers. In first stages, standard structure of DBD (include one-particle encapsulated electrode) was used in researches, but recently a new structure which include a three particles encapsulated electrode is employed. Encapsulated three-particle electrodes cause to recover the body force of plasma actuator and increase the induced velocity, significantly. In this paper by Electro-Static Lumped-element numerical method is applied and validated. Then, the three-particle and one-particle plasma actuators are used for NACA0012 and NACA0015 airfoils. It is numerically shown that the separation point displaces back and the pressure distribution is changed favorably using DBD. It is also found that the effectiveness of three-particle structure is considerably higher than the standard structure. This research was performed at Rec=1.0´106 and the plasma actuator simulated for high voltage AC power supply up to 14kVp-p with 10 kHz frequency in steady state condition.

An accurate calibration of inertial sensors, loses its validity over time. In this paper, a practical method is proposed to compensate the modeling error of a strap down inertial navigation system (INS) in order to resolve the requirement for re-calibration. Since strap down INS needs a specific turntable for its calibration, this method improves the availability of flying object that equipped with strap down INS. In this process, using the Earth’s gravity and rotation rate, deviation of calibration coefficients will be compensated. To verify effectiveness of suggested method, three different tests are performed, and it is proved that the proposed technique decreases the part of navigation error that is related to calibration deviation.