Manned water-powered aerial vehicles have been implemented into specialized missions around water bodies, such as firefighting and rescue. However, the dual requirement of vehicle motion control and performing tasks challenges operators. Moreover, in the presence of a low visibility, dense smoke, and extreme temperature, they always face potential risks. Motivated by these difficulties, this paper proposed an unmanned water-powered aerial vehicle using a nozzle rotation mechanism. This mechanism allows the vehicle to have a wide range of forces and torques in multiple directions under constant mass flowrate condition. A simple controller was designed to investigate the fundamental flight motions and verify dynamic properties of the vehicle in practical testing. To come up with the control law, the following steps were taken. Firstly, a mathematical model was derived to reflect the vehicle’s dynamic characteristics. Secondly, a well-known proportional-derivative-integral controller incorporating gravity compensation was deployed to regulate the 3-degree-of-freedom motion system. Thirdly, experiments were conducted to confirm the flight ability of the proposed vehicle. Results demonstrated that the control system preserved stability and the vehicle could fly following the desired altitude.
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A pod mounted on an aircraft external must be installed on an aircraft after its structural safety is verified under flight conditions. This paper presents methods of flight load and test load generation. Evaluation of test result data and standards for failure mode are also presented. First of all, to verify the static structural stability, flight loads for the aircraft maneuvering conditions were calculated. Finite element analysis was then performed with flight loads. As a result of the analysis, structures were verified to have a margin of safety for a given design requirement. In addition, it was confirmed that the launcher tube had enough rigidity to support the missile. Thus, the role of stinger such as longeron and hardback was insignificant. Finally, based on results of tests and analysis, the static structural stability of pod was substantiated and the reliability and effectiveness of the analysis model were obtained. These results and dynamic stability verification results suggest that an optimal design is necessary.
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In this study, a novel input signal transformer is introduced to remove an overshoot and steady-state error that hinder stable tracking and landing of an unmanned aerial vehicle (UAV) based on image recognition, and the performance is verified. The Input Signal Transformer (IST) is designed in the shape of a sigmoid function to attenuate or amplify the input signal to resolve the aforementioned problems. For the verification of the UAV control system based on the IST, the UAV, target, and sensors were implemented in a virtual environment using the robot operating system (ROS). And data exchange structure and control system were built by the ROS-based message communication. A simulation was performed to confirm the elimination of the overshoot and steady-state error when the UAV to which the above control system was applied tracks and lands a fixed target and a moving target. As a result of the simulation, when the IST was not used, the UAV performed an unstable because of the overshoot and could not land on the target by the steady-state error. Conversely, in the case of using IST, it was confirmed that the flight was stable and landed successfully.
The main function of aircraft ejection system is that it separates the store from the aircraft. The ejection force is important for the safety of the aircraft when the store is ejected, because the store can be lifted by air flow affected by the aircraft’s speed. If the ejection force is low, the aircraft can be damaged by the floating store. The ejection force of the suspension system should be designed in order to release the store safely. In this study, the ejection force of the pyrotechnic suspension using the cartridge to eject the store was researched. This research was performed, based on the precedent study about the over-center linkage mechanism and the pressure drop by the orifice. The ejection force was calculated, after analyzing mathematical fundamentals about the pressure in the system of the suspension and analyzed through AutoDyn and ADAMS software. Finally, the theoretical results were compared with the ejection test results of the suspension system.
An airborne black box should preserve the recorded data under the extreme environmental conditions such an aircraft crash. Through the recorded information from the black box, the cause of the aircraft crash can be analyzed. This requires the black box to safely protect its memory board inside from the external forces or heat generated by the aircraft crash. This paper is a case study conducted to develop the black box to satisfy the crash survival requirements of ED-112A, the standard of performance to develop the black box, and optimize its weight and size. Through the sequential design, analysis, and test process, the black box successfully demonstrated the optimized design, analysis and test subjected to crash survival requirements of ED-112A. The prototype of the black box was fabricated and tested for the verification of each analysis within the requirements of ED-112A.
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If fatigue failure occurs during aircraft operation, it can cause catastrophic injuries. So, it is necessary to study fatigue failure at the design stage. Frequency domain fatigue analysis is used to predict fatigue failure. During frequency domain fatigue analysis, results can be calibrated by PSD analysis. In this study, fatigue failure is predicted by the Dirlik method, Lalanne method and Steinberg method. Regarding results, life determined by the Dirlik method, Lalanne method and Steinberg method were 8.737, 8.314, and 7.901 times the standard life, respectively. The Steinberg method is the most conservative but the difference with other methods was approximately 10%. In the cycle histogram, the Dirlik method had more counts than the Lalanne method in lower stress range. However, it does not affect the life of material used in this study. However, if material has a lower fatigue limit or stronger PSD data is used, life difference will occur.
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Major aerospace developers continue to push for new structural composite applications to reduce the environmental impact of greenhouse gas emissions, improve both aircraft performance and costs. In this study, the parts that carry the load in the regions where mechanical joints are applied, require whole processing to tighten and identify stress concentration points. In addition, failure modes caused by bearing and by-pass loads were set as the main design factors. Optimum sizing was performed through the application of factors taken into account in the buckling failure mode and production using the preliminary design analysis model of the composite wing structure. In the area where the fuselage is joined with the fuselage, bearing and bypass load were considered important design factors.
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