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A periodic wave having a frequency that is an integral multiple of the fundamental power line frequency component is harmonics. They are the byproducts of modern electronics devices so it is necessary to mitigate the harmonics and offer techniques to mitigation of harmonics.This paper provides an explanation of the various harmonic mitigation techniques available to solve harmonic problems in three phase power systems. Included are the advantages and disadvantages of each method, their normal circuit connection as well as typical performance that can be expected when each method is properly employed.

Thrust vectoring is a requirement for fifth generation fighters, giving them super-maneuverability capabilities, allowing them to execute tactical maneuvers that are not possible using conventional aerodynamic mechanisms. The most widely used and successful method for achieving this is by using gimbaled engines or nozzles. The complexities involved in this method, have encouraged future engine designers to explore different avenues for achieving thrust vectoring, one of which is fluidic thrust vectoring. In fluidic thrust vectoring, jet deflection is achieved by fluid injection at various locations on the nozzle. During thrust vectoring operations, the engine performance is affected. This is primarily due to the change in effective nozzle area. When a nozzle is gimbaled, as is the method used in currently operational thrust vectored engines, or during fluidic thrust vectoring operations, there is a change in effective nozzle area. This impacts the engine mass flow rate, thus affecting the engine operation. The change in performance is similar to that of an engine fitted with a variable area nozzle. In this study, we attempted to retrofit a thrust vectoring nozzle to an existing engine designed for a fourth-generation fighter aircraft, in order to give it fifth-generation fighter aircraft capabilities. A Twin spool mixed flow turbofan engine with a convergent nozzle is selected and its performance is simulated using Gasturb 13. The baseline engine consists of a low pressure spool, high pressure spool, combustion chamber and convergent-divergent nozzle. For the sake of simplicity, the convergent-divergent nozzle is replaced with a convergent nozzle, with no loss in thrust at design point. The design point is arrived at based on engine data available in open literature. Following this, offdesign performance is simulated, for studying the effect of thrust vectoring operations, which are modeled as a nozzle area change. Suitably scaled generic maps provided in Gasturb are used for off-design simulations. The effect of nozzle area change on engine performance is studied at sea level static conditions. The nozzle area is decreased by a maximum of 15%, in steps of 1%. During thrust vectoring operations, there is a significant change in bypass ratio and fan surge margin, with the other performance parameters being relatively constant. Following this, simulations are conducted at different flight conditions to understand the effect of nozzle area change for different flight regimes. A total of seven different flight conditions are selected to understand the operational envelope of thrust vectoring operation. It is found that at all flight conditions, thrust vectoring has a significant influence on bypass ratio and fan surge margin. While for most conditions, there is an improvement in fan surge margin, there are two conditions where fan surge margin decreases substantially.

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During the static testing of a single spool turbojet engine (400 N Titan Engine) running on Jet A-1 as fuel, the starting sequence was unsuccessful, however, upon substituting the fuel with Diesel the engine started successfully. Till that point, the engine operation/performance was normal, and a series of tests were conducted for a total time of 5 hours. These 5 hours of testing spanned over five months during which ambient temperature varied from 5 °C to 37 °C, and relative humidity varied from 40% to 90%. The starting sequence was unsuccessful at the maximum ambient temperature (37 °C) and moderate relative humidity condition (50%). During the test, various parameters including pressure, temperature, rotor speed and mass flow rate were measured and recorded using a computer based data acquisition. For the failed start-up case, during the starting sequence, a sudden decrease in rotor speed was observed as the rotor reached ∼80% of the calibration value followed by flameout in combustion chamber. This sudden drop in rotor speed occurred due to mild-surge in compressor which continued up to flameout. The cause of mild-surge is attributed to the momentary blockage at the exit of the compressor caused by the combustion process. These pressure fluctuations propagated downstream leading to flame out. The combination of high ambient temperature (35 °C) and moderate humidity (∼50%) resulted in higher combustion chamber temperature which resulted in the momentary blockage at compressor exit during the transient operation. The substitution of Diesel as fuel resulted in lower combustion chamber temperature, and resolved the issue. It is important to note that, the mild-surge observed in this case is caused by downstream condition, and would occur only at some special situations (inlet conditions). Hence, it may be concluded, that the compressor surge in a jet engine may be caused by downstream components (under some special conditions), and design modifications to these downstream components may resolve the issue in those cases.

Turbines remain one of the most efficient devices for extracting energy from a flowing fluid. In a gas turbine engine, axial flow turbines are used to extract energy from the working fluid and drive the compressor, to which they are mechanically connected. To maximize the performance of the axial flow turbine, it is necessary to carry out a design optimization of the components while suitably accounting for losses generated by secondary flows. An axial flow turbine rig is designed, fabricated, and installed to better understand and improve upon secondary flow models used in design procedures. The rig is driven by a blower operating at a constant speed, capable of delivering a maximum airflow rate of 0.4 kg/s and a maximum pressure rise of 500 mbar across the device. The axial flow turbine is mechanically connected to a dynamometer capable of operating at a full load capacity of 5 kW and a maximum rotational speed of 10,000 RPM. The axial flow turbine, housed between the blower and dynamometer, consists of nozzle guide vanes followed by a rotor. The design pressure ratio is chosen as 1.04, based on the blower delivery conditions and dynamometer specifications. For an initial design, a low-pressure ratio low rotor speed design was selected, allowing for easy installation and testing of the rotating components. The design space for the axial flow turbine was generated by varying flow and geometrical parameters in suitable steps, using a program written in MATLAB 2020a. Using the input variables and applying free vortex theory for three-dimensional blade design, the aerodynamic design of the axial flow turbine was carried out. The axial flow turbine design is experimentally tested with suitable pressure measurements at every station. Experiments are conducted for four different air mass flow rates. At each air mass flow, the rotor speed is varied by increasing/decreasing the dynamometer load. The data is recorded and compared with the design point. The difference between the design and measured performance parameters is observed to be within acceptable limits.