JOURNAL OF AERONAUTICAL ENGINEERING (JOAE)
Year:2015 | Volume:17 | Issue:1|Papers Count:7

Rapid growth in development of Micro Electro-Mechanical Sensors (MEMS) as well as wide utilization of advanced estimation algorithms have increased application of MEMS in various fields. However, MEMS-based navigation is not efficient without an appropriate sensor calibration due to large amount of MEMS errors. Therefore, error identification and sensor calibration of micro electro-mechanical inertial sensors of gyroscopes and accelerometers have been investigated in the present paper. In this regard, Allan variance as a simple and understandable method has been adopted to identify stochastic errors of the inertial sensors. Then, all error parameters including scale factors, biases, and misalignments have been determined utilizing weighted least square method. Finally, the presentation of error sphere has shown the performance and viability of the inertial MEMS calibration process.

The flow fields over a non-slender lambda shaped wing including vortex formation and the effect of angle of attack on structure, strength and location of the vortices were investigated in a closed circuit low speed wind tunnel by pressure measurement. The tests were conducted at velocity of 20 m/s and different angles of attack from 5 to 20 degrees. The Reynolds number of the model was about 2×105 according to the root chord. The results showed three vortical flows were formed over the wing surface. The first vortex was the apex vortex; the second one was generated before the leading edge crank and strengthened after the crank and the third one was formed after the leading edge crank. By increasing the angle of attack the vortices became larger and stronger and closer to the wing root. At a specific angle of attack the structure of the vortices was changed rapidly and the vortex break down was occurred. The location of vortex break down moved toward the wing apex by increasing the angle of attack.

In this paper, a new method based on optimal combination of inertial sensors; in process of initial alignment for strap down navigation system; is proposed. The equations of initial alignment are usually based on accelerometer outputs and/or gyroscope outputs, depending on sensor’s accuracy. Our initial alignment algorithm leads to linear combination of output vectors. Although the error of this method is independent of sensor’s biases, unfortunately the coefficient of this combination is unknown. Knowledge of designer or sensor’s accuracy is a normal solution, but that obviously will not lead to the best estimation. The proposed idea is utilizing genetic algorithm to achieve optimal combination of sensors. In this regard, the optimal transformation matrix must be estimated, and the performance index is a function of alignment error. Final result of optimization problem is the best coefficient to combine outputs. The simulation results show excellent performance of proposed algorithm.

Discovering the cause of the aircraft accident is the major part of the aviation industry which has been act as the driving force of the research and development and is a reference to design & manufacturing deficiencies. In this research, considering the necessity of upgrading and updating the military aircraft accidents investigation model, the current model has been studied and its shortcomings are addressed. Afterwards, 12 efficient methods which were presented to investigate the accidents are introduced. It should be noted that there is no universal model to investigate the military aircraft accidents and it is based on the countries laws, regulations and joint military pacts. Therefore, considering the characteristics of the introduced efficient methods, a new model has been presented. In this new model, while eliminating defects and bugs of the current military aircraft accidents investigation model which only includes the accident cause discovery, investigations are focused on exploring and discovering how and why an accident occurs. Therefore, by finding the real root cause of the accident, similar accidents would be prevented in the future.

Flight Management Systems (FMS) in today’s airplanes help the pilot in guidance, control and handling flight plans based on standard tables and performance monitoring in order to reduce the pilot work load during flight. Despite great advances, modern avionics, regular checks and using multiple redundancies in vital systems, the possibility of fault or failure occurrence still exist. Considering the changes in airplane in emergency conditions and pilot challenges for fast decision making and choosing a safe landing site, the FMS must enhance to help the pilot in emergency conditions. Accordingly, the emergency FMS is introduced in this paper to cover the pilots’ challenges in emergency, effective communication between the pilot and the automatic system and therefore to improve the flight safety. The proposed FMS include different parts and subsystems like identification, structural health monitoring, controller, trajectory generation, environment database, flight envelope, pilot interface; etc. in addition, a novel landing site selection algorithm and a pilot interface are introduced as an inevitable subsystems in emergency flight management system.

The aerial vehicle performance capabilities are usually ignored during the guidance laws design. This may result in saturation of commands and undesirable behavior of integrated guidance and control system. In the current research, based on the Augmented Proportional Navigation (APN) law and considering the performance and maneuvering capabilities of a flying vehicle through different flight conditions, the parameters of APN are optimized. In contrast with the conventional APN, in the proposed guidance law, the guidance parameters are not fixed but are functions of environmental conditions and engagement scenarios. In this regard, utilizing different flight conditions and engagement scenarios, a multi-objective optimization problem is formed. The performance and maneuvering capabilities are considered as the optimization constraints. Multi-objective problem includes three objectives of miss distance, acceleration command and flight time. The NSGA-II algorithm is utilized here. The limits of the acceleration commands are not fixed and vary depending on the flight conditions. The simulation of several scenarios, reveals a superior behavior for the proposed guidance law in comparison with the conventional laws.

The present research aim is to evaluate effective design parameters on reliability of a light aircraft’s flap mechanism which equipped by electrical actuators. The flap mechanism configuration and parts were explained and the associated fault tree for the mechanism were drawn to investigate the influence of any individual part’s failure on system reliability. Also, reliability block diagram for the flap mechanism and supporting structure were depicted and associated to this diagram, system reliability were demonstrated according to each element. It has been investigated that without periodical check, the reliability drops significantly. Moreover, preflight checks effect on flap mechanism’s reliability were studied and shown the effective character of this check on the flight reliability.