Volume 40, Issue 2

Machining operations on titanium and its alloys have always been a challenge in the design of compressor blade in the jet engine which undergo profound stress during impact of debris and affects the life span of both static and rotary blades. Titanium grade 5 alloy is used to resist the creep and fatigue due to the thermal expansion and sudden impact including bird strike. However, machining them is difficult. An attempt was made to mitigate the temperature during drilling of Ti-6Al-4V by adopting cryogenic coolants LN 2 , CO 2 and a novel approach of coconut oil under varying cutting speed and feed rates. LN 2 provided 17 and 55% reduction in temperature over CO 2 and coconut oil condition respectively. Employing the optimum machining procedure for Titanium alloy can reduce the weight of the aircraft by mounting the titanium alloy at the leading edge of the carbon fiber reinforced plastic composite blades.

This paper applies a body force model developed recently to investigate the interaction between total temperature distortion and a multistage fan. The off-design performance of the fan shows the reasonable predicting accuracy and supports the present model is applicable for high-speed multistage machines. The transfer behaviors of 90° steady-state circumferential total temperature distortion as well as combined total pressure and total temperature distortion in the multistage environment are captured successfully by the model. The mechanism of the phase shift of the high temperature sector is discussed by the model to advance the understanding of the total temperature distortion problem. The results reveal that the large-scale flow feature of total temperature distortion in the multistage environment can be capably quantified by the present body force model with the acceptable computational consumption.

Temperature probes of different designs were widely used in aero gas turbine engines for measurement of air and gas temperatures at various locations starting from inlet of fan to exhaust gas from the nozzle. Exhaust Gas Temperature (EGT) downstream of low pressure turbine is one of the key parameters in performance evaluation and digital engine control. The paper presents a holistic approach towards life assessment of a high temperature probe housing thermocouple sensors designed to measure EGT in an aero gas turbine engine. Stress and vibration analysis were carried out from mechanical integrity point of view and the same was evaluated in rig and on the engine. Application of 500 g load concept to clear the probe design was evolved. The design showed strength margin of more than 20% in terms of stress and vibratory loads. Coffin Manson criteria, Larsen Miller Parameter (LMP) were used to assess the Low Cycle Fatigue (LCF) and creep life while Goodman criteria was used to assess High Cycle Fatigue (HCF) margin. LCF and HCF are fatigue related damage from high frequency vibrations of engine components and from ground-air-ground engine cycles (zero-max-zero) respectively and both are of critical importance for ensuring structural integrity of engine components. The life estimation showed LCF life of more than 4000 mission reference cycles, infinite HCF life and well above 2000 h of creep life. This work had become an integral part of the health monitoring, performance evaluation as well as control system of the aero gas turbine engine.

Reducing the carbon footprints of aerial transportation became a major target for both industries and academia. Various solutions have been proposed to develop cleaner alternative methods for green transportation. Full electric, hybrid electric, and turboelectric propulsion system architectures intend to reduce greenhouse gas emissions and fuel consumption of today’s aero gas turbine engines. In this study, a turboelectric propulsion system, which is considered as the most promising technology for future aviation is selected for modelling and simulation. As the main power supply, a high fidelity mathematical model of GE T700 turboshaft engine is constructed in MATLAB/Simulink to emulate the technology of today. Selected aero gas turbine’s mathematical model is combined with NASA’s Baseline electrical power distribution architecture which is firstly designed for N-3X turboelectric aircraft. MATLAB/Simulink model is utilized to analyses a single-engine operational scenario of twin-engine aircraft which is a major design consideration due to single-engine failure. Power requirements, distribution percentages, preliminary power assessment for power electronic systems and nominal power capacities of each electrical unit of a turboelectric propulsion system are obtained using GE T700 as the main power supply.

In this paper, an active fault-tolerant control (FTC) scheme for turbofan engines subject to simultaneous multiplicative and additive actuator faults under disturbances is proposed. First, a state error feedback controller is designed based on interval observer as the nominal controller in order to achieve the model reference rotary speed tracking control for the fault-free turbofan engine under disturbances. Subsequently, a virtual actuator based reconfiguration block is developed aiming at preserving the consistent performance in spite of the occurrence of the simultaneous multiplicative and additive actuator faults. Moreover, to improve the performance of the FTC system, the interval observer is slightly modified without reconstruction of the state error feedback controller. And a theoretical sufficiency criterion is provided to ensure the stability of the proposed active FTC system. Simulation results on a turbofan engine indicate that the proposed active FCT scheme is effective despite of the existence of actuator faults and disturbances.

The over-under turbine based combined cycle (TBCC) propulsion system dynamic model is established, including the engine model and inlet model. A dual-channel inlet is designed to match the engine and the inlet performance parameters are obtained by computational fluid dynamics (CFD) calculation, which provides original data for the integrated Inlet/Engine matching model. Control plan is designed to ensure that the engine is not over-temperature, over-rotation or surging based on correlation analysis. The simulation results show that the designed control plan can keep the TBCC propulsion system continuous in thrust during the mode transition, and the thrust fluctuation is about 10%.

Micro-propulsion is considered to be the emerging technology for the propulsion of micro and micro aerospace vehicles as it is preferred over mesoscale thruster due to lower overall life-cycle cost and launching costs. Hence this paper investigates the influence of critical parameters like the Nozzle Pressure Ratio (NPR) and Reynolds number (Re) on the operational characteristics of the micronozzle. A conical nozzle with throat diameter 710 µm and exit/throat area ratio ∼2.14 has been designed and is analyzed numerically by using a model based on pressure-based coupled implicit for various NPR, the backpressure with three Res namely, 1000, 1500, and 2000. The performance of this micronozzle has been characterized in terms of thrust, thrust coefficient, and specific impulse for all three Re cases. A subsequent analysis of the subsonic layer reveals that the nozzle is subjected to high viscous losses at low NPRs, which are independent of Re.

The use of Computational Fluid Dynamics (CFD) is now central to the design process of aero-engine combustors, enabling optimal, safe and stable operation, increased efficiencies, and the reduction of pollutant emission. To benefit maximally from the use of CFD it is essential to account for the relevant physical phenomena, in particular the fuel spray breakup and its evaporation. Different strategies for modelling the injection of fuel spray are applied - in the simplest approach the fuel is assumed to be gaseous upon injection, in the most advanced approach the fuel is modelled, using a Lagrangian-Eulerian approach, as a liquid spray which breaks up, evaporates and eventually burns inside the combustion chamber. The effects of the various modelling strategies on the flow, temperature, and compositional fields are investigated. The radial distribution of the simulated temperature field is compared to experimental data, demonstrating that acceptable accuracy is only achieved when the fuel is modelled as a liquid spray and a two-way momentum coupling between the spray and the gas-phase is accounted for.

The article describes the stability of gaseous fuel combustion in gas turbine low-emission combustion chambers with the plasma-chemical assistance. The mathematical model of unsteady processes in a low-emission combustion chamber with a plasma-chemical stabilizer that takes into consideration the impact of low-temperature plasma on aerodynamics flow in a combustion chamber and the characteristics of heat release is developed. A methodology of a numerical experiment concerning the stability of gaseous fuel combustion in a combustion chamber with plasma assistance using computational fluid dynamics, which enhances the efficiency of designing and adjustment, is proposed. Practical recommendations for improvement of stability of a gas turbine combustion chamber with partially premixed lean fuel–air mixtures, working on gaseous fuels, are developed. They allow to reduce pressure fluctuations inside the flame tube by 10–35%, to decrease spectral power of static pressure in the flame tube in 1.5–2.0 times, to reduce nitrogen oxide emission up to 33.6 ppm in the exit section while retaining a carbon monoxide emission level, that corresponds modern international ecological standards.

Experiments and simulations were carried on C-D nozzles with four different exit geometry aspect ratios to investigate the impact of supersonic decay characteristics. Rectangular and elliptical exit geometries were considered for the study with various aspect ratios. Numerical simulations and Schlieren image study were studied and found the agreeable logical physics of decay and spread characteristics. The supersonic core decay was found to be of different length for different exit geometry aspect ratio, though the throat to exit area ratio was kept constant to maintain the same exit Mach number. The impact of nozzle exit aspect ratio geometry was responsible to enhance the mixing of primary flow with ambient air, without requiring a secondary method to increase the mixing characteristics. The higher aspect ratio resulted in better mixing when compared to lower aspect ratio exit geometry, which led to reduction in supersonic core length. The behavior of core length reduction gives the identical signature for both under-expanded and over-expanded cases. The results revealed that higher aspect ratio of the exit geometry produced smaller supersonic core length. The aspect ratio of cross section in divergent section of the nozzle was maintained constant from throat to exit to reduce flow losses.

The present study investigates the effect of chevron with tab (Tabbed chevron) fixed at the exit of the co-flowing primary nozzle on the mixing characteristics in subsonic and sonic jets. The experiment was conducted for the jet Mach number 0.6 and 0.8. The centerline Mach number decay was calculated for all these jet Mach numbers for the co-flow baseline nozzle, nozzle with chevrons and tabbed chevron (primary flow) nozzle respectively. It was found that tabbed chevron was effective in reducing the potential core length by 88.23% as compared to the chevron nozzle with baseline nozzle. The radial profile also shows that the mixing enhancement by tabbed chevron is better than chevron nozzle.

We propose a new type of effervescent atomizer with bushings installed in the liquid channel perpendicular to the channel’s axis. Bushings have holes through which air is injected to create bubbles. The air is released into the gap between the channel and the bushing. This investigation evaluates the bubbles’ atomization quality. Atomizer tests were conducted at multiple water and air flow rates, under different configurations, without an exit nozzle and with a 2 mm nozzle diameter. The atomizer’s design enables a homogenous bubble flow with small air bubbles. At an ALR = 0.012–0.036 and water flow rates of 1.67 and 2.17 L/min without an exit nozzle, bubble diameters of 0.2–0.4 mm comprised 40–50% of the total number of bubbles. The number of the bubbles with diameters of 0.8–1.0 mm does not exceed 5%. After increasing the injection parameter ε twice, the average diameter of the bubbles remained constant. Upon testing, an atomizer with one bushing, 2 mm-diameter outlet nozzle, and a water flow rate of 1.67 L/min produced particle diameters of SMD = 32–100 μm at ALR values of 0.02–0.12.