Abstract: Transition from tertiary level education to employment is one of the challenges that many fresh university graduates face after graduation. The transition period or the waiting time to obtain the first employment varies with the socio-economic factors and the general characteristics of a graduate. Compared to other fields of study, Arts graduates in Sri Lanka, have to wait a long time to find their first employment. The objective of this study is to identify the determinants of the transition from higher education to employment of these graduates using survival models. The study is based on a survey that was conducted in the year 2016 on a stratified random sample of Arts graduates from Sri Lankan universities who had graduated in 2012. Among the 469 responses, 36 (8%) waiting times were interval censored and 13 (3%) were right censored. Waiting time for the first employment varied between zero to 51 months. Initially, the log-rank and the Gehan-Wilcoxon tests were performed to identify the significant factors. Gender, ethnicity, GCE Advanced level English grade, civil status, university, class received, degree type, sector of first employment, type of first employment and the educational qualifications required for the first employment were significant at 10%. The Cox proportional hazards model was fitted to model the waiting time for first employment with these significant factors. All factors, except ethnicity and type of employment were significant at 5%. However, since the proportional hazard assumption was violated, the lognormal Accelerated failure time (AFT) model was fitted to model the waiting time for the first employment. The same factors were significant in the AFT model as in Cox proportional model.
Abstract: The introduction of tilt-rotor aircraft into the existing civilian air transportation system will provide beneficial effects due to tilt-rotor capability to combine the characteristics of a helicopter and a fixed-wing aircraft into one vehicle. The disposability of reliable tilt-rotor simulation models supports the development of such vehicle. Indeed, simulation models are required to design automatic control systems that increase safety, reduce pilot's workload and stress, and ensure the optimal aircraft configuration with respect to flight envelope limits, especially during the most critical flight phases such as conversion from helicopter to aircraft mode and vice versa. This article presents a process to build a simplified tilt-rotor simulation model, derived from the analysis of flight data. The model aims to reproduce the complex dynamics of tilt-rotor during the in-flight conversion phase. It uses a set of scheduled linear transfer functions to relate the autopilot reference inputs to the most relevant rigid body state variables. The model also computes information about the rotor flapping dynamics, which are useful to evaluate the aircraft control margin in terms of rotor collective and cyclic commands. The rotor flapping model is derived through a mixed theoretical-empirical approach, which includes physical analytical equations (applicable to helicopter configuration) and parametric corrective functions. The latter are introduced to best fit the actual rotor behavior and balance the differences existing between helicopter and tilt-rotor during flight. Time-domain system identification from flight data is exploited to optimize the model structure and to estimate the model parameters. The presented model-building process was applied to simulated flight data of the ERICA Tilt-Rotor, generated by using a high fidelity simulation model implemented in FlightLab environment. The validation of the obtained model was very satisfying, confirming the validity of the proposed approach.
Abstract: Historically, actuators’ redundancy was used to deal
with faults occurring suddenly in flight systems. This technique was
generally expensive, time consuming and involves increased weight
and space in the system. Therefore, nowadays, the on-line fault
diagnosis of actuators and accommodation plays a major role in the
design of avionic systems. These approaches, known as Fault
Tolerant Flight Control systems (FTFCs) are able to adapt to such
sudden faults while keeping avionics systems lighter and less
expensive. In this paper, a (FTFC) system based on the Geometric
Approach and a Reconfigurable Flight Control (RFC) are presented.
The Geometric approach is used for cosmic ray fault reconstruction,
while Sliding Mode Control (SMC) based on Lyapunov stability
theory is designed for the reconfiguration of the controller in order to
compensate the fault effect. Matlab®/Simulink® simulations are
performed to illustrate the effectiveness and robustness of the
proposed flight control system against actuators’ faulty signal caused
by cosmic rays. The results demonstrate the successful real-time
implementation of the proposed FTFC system on a non-linear 6 DOF
aircraft model.
Abstract: This paper presents a Neural Network (NN) identification of icing parameters in an A340 aircraft and a reconfiguration technique to keep the A/C performance close to the performance prior to icing. Five aircraft parameters are assumed to be considerably affected by icing. The off-line training for identifying the clear and iced dynamics is based on the Levenberg-Marquard Backpropagation algorithm. The icing parameters are located in the system matrix. The physical locations of the icing are assumed at the right and left wings. The reconfiguration is based on the technique known as the control mixer approach or pseudo inverse technique. This technique generates the new control input vector such that the A/C dynamics is not much affected by icing. In the simulations, the longitudinal and lateral dynamics of an Airbus A340 aircraft model are considered, and the stability derivatives affected by icing are identified. The simulation results show the successful NN identification of the icing parameters and the reconfigured flight dynamics having the similar performance before the icing. In other words, the destabilizing icing affect is compensated.
Abstract: Simplified coupled engine block-crankshaft models
based on beam theory provide an efficient substitute to engine
simulation in the design process. These models require accurate
definition of the main bearing stiffness. In this paper, an investigation
of this stiffness is presented. The clearance effect is studied using a
smooth bearing model. It is manifested for low shaft displacement.
The hydrodynamic assessment model shows that the oil film has no
stiffness for low loads and it is infinitely rigid for important loads.
The deformation stiffness is determined using a suitable finite
elements model based on real CADs. As a result, a main bearing
behaviour law is proposed. This behaviour law takes into account the
clearance, the hydrodynamic sustention and the deformation stiffness.
It ensures properly the transition from the configuration low rigidity
to the configuration high rigidity.
Abstract: The measurement of aerodynamic forces and moments
acting on an aircraft model is important for the development of wind
tunnel measurement technology to predict the performance of the full
scale vehicle. The potentials of an aircraft model with and without
winglet and aerodynamic characteristics with NACA wing No. 65-3-
218 have been studied using subsonic wind tunnel of 1 m × 1 m
rectangular test section and 2.5 m long of Aerodynamics Laboratory
Faculty of Engineering (University Putra Malaysia). Focusing on
analyzing the aerodynamic characteristics of the aircraft model, two
main issues are studied in this paper. First, a six component wind
tunnel external balance is used for measuring lift, drag and pitching
moment. Secondly, Tests are conducted on the aircraft model with
and without winglet of two configurations at Reynolds numbers
1.7×105, 2.1×105, and 2.5×105 for different angle of attacks. Fuzzy
logic approach is found as efficient for the representation,
manipulation and utilization of aerodynamic characteristics.
Therefore, the primary purpose of this work was to investigate the
relationship between lift and drag coefficients, with free-stream
velocities and angle of attacks, and to illustrate how fuzzy logic
might play an important role in study of lift aerodynamic
characteristics of an aircraft model with the addition of certain
winglet configurations. Results of the developed fuzzy logic were
compared with the experimental results. For lift coefficient analysis,
the mean of actual and predicted values were 0.62 and 0.60
respectively. The coreelation between actual and predicted values
(from FLS model) of lift coefficient in different angle of attack was
found as 0.99. The mean relative error of actual and predicted valus
was found as 5.18% for the velocity of 26.36 m/s which was found to
be less than the acceptable limits (10%). The goodness of fit of
prediction value was 0.95 which was close to 1.0.
Abstract: In the present era of aviation technology, autonomous navigation and control have emerged as a prime area of active research. Owing to the tremendous developments in the field, autonomous controls have led today’s engineers to claim that future of aerospace vehicle is unmanned. Development of guidance and navigation algorithms for an unmanned aerial vehicle (UAV) is an extremely challenging task, which requires efforts to meet strict, and at times, conflicting goals of guidance and control. In this paper, aircraft altitude and heading controllers and an efficient algorithm for self-governing navigation using MATLAB® mapping toolbox is presented which also enables loitering of a fixed wing UAV over a specified area. For this purpose, a nonlinear mathematical model of a UAV is used. The nonlinear model is linearized around a stable trim point and decoupled for controller design. The linear controllers are tested on the nonlinear aircraft model and navigation algorithm is subsequently developed for for autonomous flight of the UAV. The results are presented for trajectory controllers and waypoint based navigation. Our investigation reveals that MATLAB® mapping toolbox can be exploited to successfully deliver an efficient algorithm for autonomous aerial navigation for a UAV.
Abstract: Article presents the geometry and structure
reconstruction procedure of the aircraft model for flatter research
(based on the I22-IRYDA aircraft). For reconstruction the Reverse
Engineering techniques and advanced surface modeling CAD tools
are used. Authors discuss all stages of data acquisition process,
computation and analysis of measured data. For acquisition the three
dimensional structured light scanner was used. In the further sections,
details of reconstruction process are present. Geometry
reconstruction procedure transform measured input data (points
cloud) into the three dimensional parametric computer model
(NURBS solid model) which is compatible with CAD systems.
Parallel to the geometry of the aircraft, the internal structure
(structural model) are extracted and modeled. In last chapter the
evaluation of obtained models are discussed.