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Suggests that the next civil supersonic passenger aircraft project will pose a number of challenges. The propulsion system for this aircraft will have to achieve economic operation for both supersonic and subsonic cruise modes. In addition, the current and intended noise and pollutant emissions legislation will have to be met. Suggests that, while there are a number of proposed engines for the next generation civil supersonic aircraft, they all exhibit difficulties in meeting the compromises inherent in the engine duty. Offers a novel solution based on a unique design. Discusses the underlying issues and presents the design based on retractable fans driven by a single stage double pass tip turbine.

The purpose of this paper is to study the condensation flow of wet steam in the last stage of a steam turbine and to obtain the distribution of condensation parameters such as nucleation rate, Mach number and wetness.

Because of the sensitivity of the condensation parameter distribution, a double fluid numerical model and a realizable k-ε-k_d turbulence model were applied in this study, and the numerical solution for the non-equilibrium condensation flow is provided.

The simulation results are consistent with the experimental results of the Bakhtar test. The calculation results indicate that the degree of departure from saturation has a significant impact on the wet steam transonic condensation flow. When the inlet steam deviates from the saturation state, shock wave interference and vortex mixing also have a great influence on the distribution of water droplets.

The research results can provide reference for steam turbine wetness losses evaluation and flow passage structure optimization design.

This paper aims to present the designing and investigating various types of impulse blade profiles to find the optimal profile that has better performance than the first or original blade. The studied model is a turbine with an output power below 1 MW and a large pressure ratio up to 20, which is used to gain relatively high specific work output. As a result of its low mass flow rate, the turbine is used under partial-admission conditions. The turbine’s stator is a group of convergence–divergence nozzles that provide supersonic flow.

More than 10 types of two-dimensional blade profiles were designed using the developed preliminary design calculations and numerical analysis. The numerical results are validated using the existing experimental results. Finally, the case with improved performance is introduced as the final optimum case.

It was found that the performance parameters such as efficiency, power and torque are increased by more than 8% in the selected best model, in comparison with the original model. Moreover, the total pressure loss is 12% decreased for the selected model. Finally, the selected profile with superior performance is proposed.

Simultaneous numerical tests are conducted to examine the interaction of different supersonic blade profiles with the partially injected flow to the rotor.

This study aims to presenting an empirical model for partially admitted turbine efficiency. When the design mass-flow rate is too small that a normal full-admission design would give very-small blade height, it may be advantageous to use partial admission. The losses due to partial admission with long blades may be less than the losses due to leakage and low Reynolds-number of the full-admission turbines with short blades. The turbine efficiency is highly dependent on the degree of partial admission. The empirical model of turbine efficiency is necessary for simulation and analysis of dynamic performances of the turbine system. In this work, appropriate empirical loss correlations are introduced and a proper model is proposed for turbine efficiency.

Experimental and numerical tests are conducted to evaluate the proposed model and the results are compared with the results of existing models. In this work, the effect of nozzles overlapping on the flow pattern is emphasized. Therefore, various models with different degrees of overlapping are simulated and their effects on the turbine efficiency are subsequently evaluated.

A suitable cubic polynomial expression for small axial supersonic turbine efficiency in experiments is suggested. The overlapping nozzles cause change in the flow pattern and the entropy distribution. Therefore, any change in the degree of overlapping of nozzles changes the efficiency of the turbine.

In this work, time-consuming numerous experimental and numerical tests of the turbine are required.

Implication of a proper formula for a partially admitted turbine may result in enhanced prediction and dynamic performance evaluation of the test turbine.

A proper empirical model for a partially admitted supersonic turbine is introduced. This model is suitable for one blocked partially admitted turbine with Mach number between 1.2 and 1.8.

This paper is concerned with improving the flow pattern in the nozzle-rotor axial gap in impulse turbines using a genetic algorithm (GA) and 3D numerical analysis. The paper aims to discuss these issues.

The appropriate model was used to estimate the turbine performance introduced in the beginning of the work. Then, the nozzle design parameters that are effective in the axial gap flow pattern are optimized using a non-linear optimization code. This code works based on the GA theory. Since the GA results are not conclusive, the selected cases were evaluated using 3D numerical analysis. For a detailed comparison of the flow pattern in initial and improved cases, a transient analysis was done. Experimental tests were performed in order to validate the work. For this purpose, the characteristic curves of the turbines were studied and compared with each other.

Improving the nozzle-rotor axial gap flow pattern leading to increase in the total-to-total efficiency of the turbine by more than two points.

Partially injected flow forced to use the full model computational analysis.

Weight reduction in a feeding system.

New loss modeling method presented for partial admission condition.

Originally, this article took the form of the Twenty‐first Brancker Memorial Lecture delivered to a meeting of The Institute of Transport. The author began his lecture by saying how honoured he was by the invitation to present the 1964 Brancker Memorial Lecture and that he felt especially privileged to have the opportunity of surveying a prospect which he believed would have excited Sir Sefton Brancker's most ardent enthusiasm—the prospect of reducing inter‐continental journey times‐by air to the same durations as those universally accepted for inter‐city journeys by rail and road. Previous Brancker Memorial Lectures had summarized the general development of British civil aviation from its earliest days to 1946 and had covered particular aspects of its very rapid expansion since that date. 1946 was a significant year because it marked the resurgence of commercial flying after seven years of wartime restrictions and regulation; it promised a new deal to both operators and travelling public, with the opportunity of usefully applying technical advances achieved during the war period; at the same time it threw into sharp contrast the relative design capabilities of the British and American aircraft manufacturing industries.

OPERATING to increasingly severe noise and NOx emission levels and at the lowest Specific Fuel Consumption (SFC) attainable, current turbofan engines represent considerable advances in all respects on powerplants of only a few years ago. Nevertheless, the demand for even larger engines has meant that the major manufacturers are making considerable efforts to develop operating cycles and improved components to attain optimum performance for the foreseeable future.

In this study, the performance analysis of the split flow turbofan engine with afterburners has been examined using the parametric cycle analysis method. The purpose of this study is to examine how engine performance is impacted by design parameters and flight ambient values and to develop a software in this context.

Software has been developed using the open-source PYTHON programming language to perform performance analysis. Mach number, compressor/fan pressure ratio, bypass ratio and density were used as parameters. The effects of these variables on engine performance parameters were investigated.

Parametric cycle analysis has been calculated for different flight conditions in the range of 0–2 M and 0–15,000 m altitude for turbofan engines. With this study, basic data were obtained to optimize according to targeted flight conditions.

As a result of the performance analysis, the association between the flight conditions and design parameters of engine were determined. A software has been developed that can be used in the design of supersonic gas turbine engines for fast and easy simulation of the design parameters.

The variables used in the literature have been analyzed, and the results of the studies have been incorporated into the developed software, which can be used in innovative engine design. Software is capable to be developed further with the integration of new algorithms and models.

In this two‐part paper, aeroelastic analysis of turbomachinery blade rows and phase‐lagged boundary conditions used for analysis are described. Part I of the paper describes a study of phase‐lagged boundary condition methods used for non‐zero interblade phase angle analysis. The merits of time‐shifted (direct‐store), Fourier decomposition and multiple passage methods are compared. These methods are implemented in a time marching Euler/Navier‐Stokes solver and are applied to a fan for subsonic and supersonic inflow and to a turbine geometry with supersonic exit flow. Results showed good comparisons with published results and measured data. The time‐shifted and Fourier decomposition methods compared favorably in computational costs with respect to multiple passage analysis despite a slower rate of convergence. The Fourier‐decomposition method was found to be better suited for workstation environment as it required significantly less storage, although at the expense of slightly higher computational cost. The time‐shifted method was found to be better suited for computers where fast input‐output devices are available.

The purpose of this paper is to comprehensively evaluate the progress in the development of trapped vortex combustors (TVCs) in the past three decades. The review aims to identify the needs, predict the scope and discuss the challenges of numerical simulations in TVCs applied to gas turbines.

TVC is an emerging combustion technology for achieving low emissions in gas turbine combustors. The overall operation of such TVCs can be on very lean mixture ratio and hence it helps in achieving high combustion efficiency and low overall emission levels. This review introduces the TVC concept and the evolution of this technology in the past three decades. Various geometries that were explored in TVC research are listed and their operating principles are explained. The review then categorically arranges the progress in computational studies applied to TVCs.

Analyzing extensive literature on TVCs the review discusses results of numerical simulations of various TVC geometries. Numerical simulations that were used to optimize TVC geometry and to enhance mixing are discussed. Reactive flow studies to comprehend flame stability and emission characteristics are then listed for different TVC geometries.

To the best of the authors’ knowledge, this review is the first of its kind to discuss extensively the computational progress in TVC development specific to gas turbine engines. Earlier review on TVC covers a wide variety of applications including land-based gas turbines, supersonic Ramjets, incinerators and hence compromise on the depth of analysis given to gas turbine engine applications. This review also comprehensively group the numerical studies based on geometry, flow and operating conditions.