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Because of the appreciable application of heat recovery systems for the increment of overall efficiency of micro gas turbines, promising evaluation and optimization are crucial. This paper aims to propose a multi-factor theoretical methodology for analysis, optimization and comparison of potential plate-fin recuperators incorporated into micro gas turbines. Energetic, exergetic, economic and environmental factors are covered.

To demonstrate applicability and reliability of the methodology, detailed thermo-hydraulic analysis, sensitivity analysis and optimization are conducted on the recuperators with louver and offset-strip fins using a genetic algorithm. To assess the relationship between investment cost and profit for the recuperated systems, payback period (PBP), which incorporates all the factors is used as the universal objective function. To compare the performance of the recuperated and non-recuperated systems, exergy efficiency, exergy destruction and corresponding cost rate, fuel consumption and environmental damage cost rates, capital and operational cost rates and acquired profit rates are determined.

Based on the results, optimal PBP of the louvered-fin recuperator (147 days) is slightly lower than that with offset-strip fins (153 days). The highest profit rate is acquired by reduction of exergy destruction cost rate and corresponding decrements for louver and offset-strip fins are 2.3 and 3.9 times compared to simple cycle, respectively.

This mathematical study, for the first time, focuses on introducing a reliable methodology, which covers energetic, exergetic, economic and environmental points of view beneficial for design and selection of efficient plate-fin recuperators for micro gas turbine applications.

Diesel has traditionally been considered the best-suited and most widely used fuel in various sectors, including manufacturing industries, power production, automobiles and transportation. However, with the ongoing crisis of fossil fuel inadequacy, the search for alternative fuels and their application in these sectors has become increasingly important. One particularly interesting and beneficial alternative fuel is biodiesel derived from bio sources.

In this research, an attempt was made to use biodiesel in an unconventional micro gas turbine engine. It will remove the concentric use of diesel engines for power production by improving fuel efficiency as well as increasing the power production rate. Before the fuel is used enormously, it has to be checked in many ways such as performance, emission and combustion analysis experimentally.

In this paper, a detailed experimental study was made for the use of Spirulina microalgae biodiesel in a micro gas turbine. A small-scale setup with the primary micro gas turbine and secondary instruments such as a data acquisition system and AVL gas analyser. The reason for selecting the third-generation microalgae is due to its high lipid and biodiesel production rate. For the conduction of experimental tests, certain conditions were followed in addition that the engine rotating rpm was varied from 4,000, 5,000 and 6,000 rpm. The favourable and predicted results were obtained with the use of microalgae biodiesel.

The performance and combustion results were not exactly equal or greater for biodiesel blends but close to the values of pure diesel; however, the reduction in the emission of CO was at the appreciable level for the used spirulina microalgae biodiesel. The emission of nitrogen oxides and carbon dioxide was a little higher than the use of pure diesel. This experimental analysis results proved that the use of spirulina microalgae biodiesel is both economical and effective replacement for fossil fuel.

The lean blowout (LBO) limit of the combustor is one of the important performance parameters for any gas turbine combustor design. This study aims to predict the LBO limits of an in-house designed swirl stabilized 3kW can-type micro gas turbine combustor.

The experimental prediction of LBO limits was performed on 3kW swirl stabilized combustor fueled with methane for the combustor inlet velocity ranging from 1.70 m/s to 6.80 m/s. The numerical prediction of LBO limits of combustor was performed on two-dimensional axisymmetric model. The blowout limits of combustor were predicted through calculated average exit gas temperature (AEGT) method and compared with experimental predictions.

The results show that the predicted LBO equivalence ratio decreases gradually with an increase in combustor inlet velocity.

This LBO limits predictions will use to fix the operating boundary conditions of 3kW can-type micro gas turbine combustor. This methodology will be used in design stage as well as in the testing stage of the combustor.

This is a first effort to predict the LBO limits on micro gas turbine combustor through AEGT method. The maximum uncertainty in LBO limit prediction with AEGT is 6 % in comparison with experimental results.

This paper aims to predict the effect of combustor inlet area ratio (CIAR) on the lean blowout limit (LBO) of a swirl stabilized can-type micro gas turbine combustor having a thermal capacity of 3 kW.

The blowout limits of the combustor were predicted predominantly from numerical simulations by using the average exit gas temperature (AEGT) method. In this method, the blowout limit is determined from characteristics of the average exit gas temperature of the combustion products for varying equivalence. The CIAR value considered in this study ranges from 0.2 to 0.4 and combustor inlet velocities range from 1.70 to 6.80 m/s.

The LBO equivalence ratio decreases gradually with an increase in inlet velocity. On the other hand, the LBO equivalence ratio decreases significantly especially at low inlet velocities with a decrease in CIAR. These results were backed by experimental results for a case of CIAR equal to 0.2.

Gas turbine combustors are vulnerable to operate on lean equivalence ratios at cruise flight to avoid high thermal stresses. A flame blowout is the main issue faced in lean operations. Based on literature and studies, the combustor lean blowout performance significantly depends on the primary zone mass flow rate. By incorporating variable area snout in the combustor will alter the primary zone mass flow rates by which the combustor will experience extended lean blowout limit characteristics.

This is a first effort to predict the lean blowout performance on the variation of combustor inlet area ratio on gas turbine combustor. This would help to extend the flame stability region for the gas turbine combustor.

This paper seeks to evaluate the potential of heat exchanged aeroengines for future Unmanned Aerial Vehicle (UAV), helicopter, and aircraft propulsion, with emphasis placed on reduced emissions, lower fuel burn, and less noise.

Aeroengine performance analyses were carried out covering a wide range of parameters for more complex thermodynamic cycles. This led to the identification of major component features and the establishing of preconceptual aeroengine layout concepts for various types of recuperated and ICR variants.

Novel aeroengine architectures were identified for heat exchanged turboshaft, turboprop, and turbofan variants covering a wide range of applications. While conceptual in nature, the results of the analyses and design studies generally concluded that heat exchanged engines represent a viable solution to meet demanding defence and commercial aeropropulsion needs in the 2015‐2020 timeframe, but they would require extensive development.

As highlighted in Parts I and II, early development work was focused on the use of recuperation, but this is only practical with compressor pressure ratios up to about 10. For today's aeroengines with pressure ratios up to about 50, improvement in SFC can only be realised by incorporating intercooling and recuperation. The new aeroengine concepts presented are clearly in an embryonic stage, but these should enable gas turbine and heat exchanger specialists to advance the technology by conducting more in‐depth analytical and design studies to establish higher efficiency and “greener” gas turbine aviation propulsion systems.

It is recognised that meeting future environmental and economic requirements will have a profound effect on aeroengine design and operation, and near‐term efforts will be focused on improving conventional simple‐cycle engines. This paper has addressed the longer‐term potential of heat exchanged aeroengines and has discussed novel design concepts. A deployment strategy, aimed at gaining confidence with emphasis placed on assuring engine reliability, has been suggested, with the initial development and flight worthiness test of a small recuperated turboprop engine for UAVs, followed by a larger recuperated turboshaft engine for a military helicopter, and then advancement to a larger and far more complex ICR turbofan engine.

The purpose of this study is to predict the performance and emission characteristics of micro gas turbine engines powered by alternate fuels. The micro gas turbine engine performance, combustion and emission characteristics are analyzed for the jet fuel with different additives.

The experimental investigation was carried out with Jet A-1 fuel on the gas turbine engines at different load conditions. The primary blends of the Jet A-1 fuels are from canola and solid waste pyrolysis oil. Then the ultrasonication of highly concentrated multiwall carbon nanotubes is carried with the primary blends of canola (Jet-A fuel 70%, canola 20% and 10% ethanol) and P20E (Jet-A 70% fuel, 20% PO and 10% ethanol).

The consumption of the fuel is appreciable with the blends at a very high static thrust. The 39% reduction in thrust specific fuel consumption associated with a 32% enhance in static thrust with P20E blend among different fuel blends. Moreover, due to the increase in ethanol concentration in the blends PO20E and C20E lead to a 22% rise in thermal efficiency and a 9% increase in higher oxygen content is observed.

The gas turbine engine emits very low emission of gases such as CO, CO₂ and NOx by using the fuel blends, which typically reduces the fossil fuel usage limits with reduced pollutants.

The emission of the gas turbine engines is further optimized with the addition of hydrogen in Jet-A fuel. That is leading to high specific fuel exergy and owing to the lower carbon content in the hydrogen fuel when compared with that of the fossil fuels used in gas turbine engines. Therefore, the usage of hydrogen with nanofluids was so promising based on the results obtained for replacing fossil fuels.

This paper aims to investigate the effect of additives in Jet-A fuel blends, especially on performance, combustion and emission characteristics.

Jet-A fuel was formed by using Kay’s and Gruenberg–Nissan mixing rules by adding additive glycerol with TiO₂. While measuring the combustion performance, the amount of oxygen content present in fuel and atomization are the key factors to consider. As such, the Jet-A fuel was created by adding additives at different proportion. A small gas turbine engine was used for conducting tests. All tests were carried out at different load conditions for all the fuel blends such as neat Jet-A fuel, G10T (glycerol 10% with 50 ppm TiO₂ and Jet-A 90%), G20T (glycerol 10% with 50 ppm TiO₂ and Jet-A 90%) and G30T (glycerol 10% with 50 ppm TiO₂ and Jet-A 90%).

From tests, the G20T and G10T produced better results than other blends. The thermal efficiency of the blends of G20T and G10T are 22% and 14% higher than neat Jet-A fuel. Further, the improved static thrust with less fuel consumption was noticed in G20T fuel blend.

The G20T blends showed better performance because of the increased oxygenated compounds in the fuel blends. Moreover, the emission rate of environmentally harmful gases such as NOx, CO and HC was lower than the neat Jet-A fuel. From the results, it is clear that the rate of exergy destruction is more in the combustion chamber than the other components of fuel.

The purpose of this paper is to analyze the performance of the gas turbine cycle integrated with solid oxide fuel cell technology. In the present work, intermediate temperature solid oxide fuel cell has been considered, as it is economical, can attain an activation temperature in a quick time, and also have a longer life compared to a high-temperature solid oxide fuel cell, which helps in the commercialization and can generate two ways of electricity as a hybrid configuration.

The conceptualized cycle has been analyzed with the help of computer code developed in MATLAB with the help of governing equations. In this work, the focus is on the performance investigation of a Gas turbine intermediate temperature solid oxide fuel cell hybrid cycle. The work also analyzes the performance behavior of the proposed cycle with various design and operating parameters.

It is found that the power generation efficiency of the IT-SOFC-GT hybrid system reaches up to 60% (LHV) for specific design and operating conditions. The cycle calculations of an IT-SOFC-GT hybrid system and its conceptual design have been presented in this work.

The unique feature of this work is that IT-SOFC has been adopted for integration instead of HT-SOFC, and this work also provides the performance behavior of the hybrid system with varying design and operating parameters, which is the novelty of this work. This work has significant scientific merit, as the cost involved for the commercialization of IT-SOFC is comparatively lower than HT-SOFC and provides a good option to energy manufacturers for generating clean energy at a low cost.

This paper aims to develop a dynamic performance model of three-shaft gas turbine for electricity generation and to study a multi-loop control strategy of three-shaft gas turbine for electricity generation.

In this paper, the dynamic performance model of the three-shaft gas turbine is established and developed. A novel approach, variable partial differential coefficient deviation linearization method is used to simulate the dynamic performance of the three-shaft gas turbine. Single-loop control system, feed-forward feedback control system and cascade system are assessed to control the engine during transient operation.

A novel approach, variable partial differential coefficient deviation linearization method is used to simulate the dynamic performance of the three-shaft gas turbine. According to the results shown, the cascade control system is most satisfactory due to its fastest response and the best stability and robustness.

The method of variable partial linearization is adopted to make the dynamic simulation of the model achieve higher precision, better steady state and less computation time. This paper provides a theoretical study for the multi-loop control system of a marine three-shaft gas turbine.

This study aims to investigate the prevalence of shock boundary layer interaction (SBLI) in air-breathing intake system is highly undesirable since this leads to high pressure gradients, typical stream mutilation and pressure drop. A novel flow control mechanism is incorporated in this research holding an array configuration of passive flow control device (micro ramps [MR]) that is adapted to improve the boundary layer stability.

Two geometric variants of the MR, namely, MR40 and MR80 is considered which reduce the pressure drop during SBLI. The incidence oblique shock wave angle of 34° is considered for the modelling. Large eddy simulation (LES) turbulence model was used with subgrid models of Wall modelled LES, Smagorinsky–Lilly to compute the unsteady effects of SBLI control using micro vortex generators. The unsteady results are compared with steady Reynold’s average Naviers–Stoke’s equation for calibrating the turbulence models.

The array configuration of MR80 reduces the pressure drop by 22% as compared with no ramp configuration and also reduces the flow distortion in hypersonic inlet. The most affected region of the MR is in the vicinity of center-line. Quantitative results prove that the upstream influence of the shock waves has been largely reduces by MR80 array configuration as compared to single MR80 pattern configuration. Different vortex structures found in the experiments was exclusively predicted using LES.

This paper substantiates the requirement of MR array configuration for transferring the momentum from free stream to the boundary layer and thereby energizing the boundary layer. This process of energization delays the flow separation in hypersonic flow.