AAA provides a powerful framework to support the iterative and non-unique process of aircraft preliminary design. The AAA program allows students and preliminary design engineers to take an aircraft configuration from early weight sizing through open loop and closed loop dynamic stability and sensitivity analysis, while working within regulatory and cost constraints.
AAA is used for preliminary and Class II design and stability and control analysis of new and existing aircraft. Class II design incorporates detailed weight & balance, aerodynamics, stability & control calculations including trim analysis and flying qualities used in conjunction with the preliminary design sequence. Class II design accounts for power plant installation, landing gear disposition and component locations on the airplane. Class II uses more sophisticated methods than Class I and requires more detailed information of the airplane to be known. The accuracy of Class II methods is therefore greater than Class I methods.
The purpose of the Analysis submodule is to provide the user with Class II analysis methods for predicting the performance characteristics of an aircraft. The methodology used to analyze the performance characteristics can be found in Chapter 5 of Airplane Design Part VII and Airplane Aerodynamics and Performance.
The purpose of the Control module is to help the user analyze single and double loop feedback control systems of the aircraft. If the open loop dynamic characteristics of the aircraft are known, root locus analyses may be performed in the S-plane. The control analysis submodule can also be used to analyze a system open loop transfer function in the frequency domain (Bode diagram). The methodology used to analyze feedback control systems can be found in Airplane Flight Dynamics Part II.
The transfer functions can be selected from the standard aircraft transfer functions or defined by the user. If the longitudinal and lateral-directional stability derivatives of the aircraft are known, the user may use the Dynamics module prior to using the Control analysis module to generate the longitudinal and lateral-directional dynamic transfer functions of the aircraft. These transfer functions are transferred into the Control analysis module and can only be generated in the Dynamics module.
This research is an effort to understand current method and to propose an advanced method for Damage Tolerance Analysis (DTA) for the purpose of monitoring the aircraft service life. As one of tasks in the DTA, the current indirect Individual Aircraft Tracking (IAT) method for the F-16C/D Block 32 does not properly represent changes in flight usage severity affecting structural fatigue life. Therefore, an advanced aircraft service life monitoring method based on flight-by-flight load spectra is proposed and recommended for IAT program to track consumed fatigue life as an alternative to the current method which is based on the crack severity index (CSI) value. Damage Tolerance is one of aircraft design philosophies to ensure that aging aircrafts satisfy structural reliability in terms of fatigue failures throughout their service periods. IAT program, one of the most important tasks of DTA, is able to track potential structural crack growth at critical areas in the major airframe structural components of individual aircraft. The F-16C/D aircraft is equipped with a flight data recorder to monitor flight usage and provide the data to support structural load analysis. However, limited memory of flight data recorder allows user to monitor individual aircraft fatigue usage in terms of only the vertical inertia (NzW) data for calculating Crack Severity Index (CSI) value which defines the relative maneuver severity. Current IAT method for the F-16C/D Block 32 based on CSI value calculated from NzW is shown to be not accurate enough to monitor individual aircraft fatigue usage due to several problems. The proposed advanced aircraft service life monitoring method based on flight-by-flight load spectra is recommended as an improved method for the F-16C/D Block 32 aircraft. Flight-by-flight load spectra was generated from downloaded Crash Survival Flight Data Recorder (CSFDR) data by calculating loads for each time hack in selected flight data utilizing loads equations. From the comparison of interpolated fatigue life using CSI value and fatigue test results, it is obvious that proposed advanced IAT method via flight-by-flight load spectra is more reliable and accurate than current IAT method. Therefore, the advanced aircraft service life monitoring method based on flight-by-flight load spectra not only monitors the individual aircraft consumed fatigue life for inspection but also ensures the structural reliability of aging aircrafts throughout their service periods.
One of the problems in the aviation world is that accidents are increasing. The increase in incidents and accidents can be caused by several factors including human error, weather conditions, and failure of the structure / component of the aircraft. Accidents that occur from failure factors, one of which is caused by cracks that occur in the structure due to continuous loading which does not become more attention until the failure. In this final project an analysis of damage tolerance is carried out at the Wing Centre, Lower Surface and Skin Access Hole locations. The choice of location is because aircraft wings are the most important component in aircraft control while operating. Damage tolerance analysis is carried out as an effort to minimize failures in aircraft structures due to cracks that arise. Damage tolerance analysis is done by giving load in the form of an aircraft wing, then an analysis of the growth of crack length that occurs due to loading. The result of damage tolerance analysis will be obtained crack length, aircraft flight life cycle, and critical crack value. And in this study will also bring up a recommendation that is the right time to do the inspection interval.
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First, rerun the service analysis using the gross moment of inertia. To force the service analysis to use the gross moment of inertia, the service modifier in the advanced analysis criteria should be larger than 1 divided by the bending cracked factor assigned in RAM Modeler. The program will cap the stiffness at 1.0 times the gross moment of inertia in the analysis. If the maximum service moment in the wall considering 2nd order effects is below 2/3 the cracking moment, then the wall remains uncracked for the service combination per ACI slender wall. The assumption to use the gross moment of inertia was valid.
RAM Concrete Shear Wall has been renamed RAM Concrete Wall in v17.00. The program was enhanced to allow reinforcement to be placed asymmetrically and consider reveals. In addition, consideration of the advanced analysis strength combinations and tilt-up reinforcement cover have been added.
Structural damage reduces the lifespan and reliability of engineering structures such as aircraft, buildings, and bridges, and can lead to serious fatalities and economic losses. The monitoring of structural damage is essential to improve the lifetime safety, maintainability, and reliability of structures. Therefore, structural health monitoring (SHM) techniques have been developed to monitor the extraction of damage-sensitive features such as strain, acoustic emission (AE), vibration signals, and electromechanical impedance, recorded using various sensors installed on the structure and determine the current state of structural health through statistical analysis of the extracted features1,2,3,4,5,6. Recently, with the development of high-performance graphics processing units and parallel computing, convolutional neural networks (CNN)-based approach using the 2D images for detecting damage have been proposed7,8,9,10.
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