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THE first gas turbine propelled aircraft in this country were the result of Whittle's classic conception using a single‐stage centrifugal compressor. On the other hand the German turbo‐jets had, without exception, multi‐stage axial compressors. The two types are shown diagrammatically in FIG. 1 and the outstanding differences are apparent at a glance. The centrifugal is short and of large diameter and the air flow through the compressor is turned from the axial direction to the radial and then back to the axial. On the other hand, the axial compressor derives its name from the substantially unidirectional flow of the air. It is of relatively small diameter, but much longer because of its many stages, each stage consisting of a large number of moving blades and an equal number of fixed blades. Altogether there may be between one thousand and two thousand individual blades in the compressor. It is from these contrasting features that much argument has arisen.

This paper aims to validate and analyse the NASA35 axial compressor performance based on a numerical approach.

Knowledge about flow property change during compressor operation at high and relatively low speed is still limited. This work provides a numerical approach to address these problems. Validation of numerical methods is proposed to generate confidence the numerical approach adopted, and after that, analysis of compressor performance at different operation conditions is carried out.

The numerical methods proposed are proved capable in predicting compressor performance. Changes of flow property during compressor operation are discussed and explained.

The current numerical work is carried out based on the first stage of the NASA35 axial compressor, where the interactive effects from adjacent stage are not counted in. Furthermore, the steady-state simulation enforces an averaging of flow at rotor-stator interface, where the transient rotor-stator interaction is removed.

This work validates the numerical methods used in the prediction of NASA35 axial compressor performance, and a similar numerical approach can be used for other turbomachinery simulation cases.

This work reinforces the understanding of axial compressor operation and provides reliable results for further investigation of a similar type of compressor. In addition, details of flow field within the NASA35 compressor during operation are given and explained which experiments still have difficult to achieve.

This study aims to use computational fluid dynamics (CFD) to understand and quantify the overall blockage within a transonic axial flow compressor (AFC), and to develop an efficient collaborative design optimization method for compressor aerodynamic performance and stability in conjunction with a surrogate-assisted optimization technique.

A quantification method for the overall blockage is developed to integrate the effect of regional blockages on compressor aerodynamic stability and performance. A well-defined overall blockage factor combined with efficiency drives the optimizer to seek the optimum blade designs with both high efficiency and wide-range stability. An adaptive Kriging-based optimization technique is adopted to efficiently search for Pareto front solutions. Steady and unsteady numerical simulations are used for the performance and flow field analysis of the datum and optimum designs.

The proposed method not only remarkably improves the compressor efficiency but also significantly enhances the compressor operating stability with fewer CFD calls. These achievements are mainly attributed to the improvement of specific flow behaviors oriented by the objectives, including the attenuation of the shock and weakening of the tip leakage flow/shock interaction intensity.

CFD-based design optimization of AFC is inherently time-consuming, which becomes even trickier when optimizing aerodynamic stability since the stall margin relies on a complete simulation of the performance curve. The proposed method could be a good solution to the collaborative design optimization of aerodynamic performance and stability for transonic AFC.

This study’s investigation aims to clarify the effect of an additional geometry, i.e. a fillet radius, to the blades of a single-stage transonic axial compressor, NASA Stage 37, on its aerodynamic and structural performances.

Applying the commercial simulation software and the one-way fluid–structure interaction (FSI) approach, this study first evaluated the simulation results with the experimental data for the aerodynamic performances. Second, this paper compared the structural performances between the models with and without fillets.

This research analyses the aerodynamic results (i.e. total pressure ratio, adiabatic efficiency, stall margin) and the structural outcomes (i.e. equivalent von Mises stress, total deformation) of the single-stage transonic axial compressor NASA Stage 37.

This paper mentions the influence of blade fillets (i.e. both rotor hub fillet and stator shroud fillet) on the compressor performances (i.e. the aerodynamic and structural performances).

This paper aims to understand the aerodynamic blockage related to near casing flow in a transonic axial compressor using numerical simulations and to design an optimum casing groove for stall margin improvement using a surrogate optimisation technique.

A blockage parameter (Ψ) is introduced to quantify blockage across the blade domain. A surrogate optimisation technique is then used to find the optimum casing groove design that minimises blockage at an axial location where the blockage is maximum at near stall conditions.

An optimised casing groove that improves the stall margin by about 1% can be found through optimisation of the blockage parameter (Ψ).

Optimising for stall margin is rather lengthy and computationally expensive, as the stall margin of a compressor will only be known once a complete compressor map is constructed. This study shows that the cost of the optimisation can be reduced by using a suitably defined blockage parameter as the optimising parameter.

A gas turbine comprising a rotor having peripheral blades, a hollow shaft driven by said rotor, a casing around said shaft, at least one combustion chamber furnishing hot gases, means for applying said gases to said blades in a direction parallel to said shaft for driving said rotor, a main compressor driven by said shaft for compressing air to feed said combustion chamber, said rotor and said main compressor being arranged towards opposite ends of said casing, and an axial‐type compressor with multiple rotors and stators inside said casing, said first‐mentioned rotor being aperturcd to admit cold air to said axial‐type compressor, and said axial‐type compressor forcing the air to said main compressor in a direction opposite to the flow of said gases past said blades, in combination with another compressor driven by said shaft and drawing cold air through said hollow shaft, said other compressor feeding air to at least one combustion chamber furnishing hot gases to drive said first‐mentioned rotor.

THE aircraft gas turbine is intriguing in that there were early attempts at its development not only by the established aero engine companies and research establishments in many countries, but also by manufacturers of marine and industrial turbines and — most successfully — by individuals. The aero engine companies failed because in virtually every instance they attempted to produce a power unit of comparable or lower specific fuel consumption to the traditional piston engine. This led to unduly complex designs involving unattainably high component efficiencies and turbine temperatures at far too early a stage in the development of the new prime mover.

IT is well known that war gives a great impetus to development in many fields, not least of which is that of aircraft propulsion. Such was the case in World War II, when great strides were made, but it is interesting to note that the pace has hardly slackened in the years following its conclusion. This is perhaps because of the ‘cold’ war which took its place, or perhaps because the introduction of jet propulsion has stimulated thought and action in realms beyond the dreams of the piston engine era. Whatever the cause, the results are apparent and this is a suitable moment to look back and measure the progress of the past seven or eight years.

THE development of aircraft penumatic equipment has progressed so satisfactorily in recent years that actuation of vital services by means of compressed air is becoming increasingly popular. In this brief review of developments to date an attempt will be made to show how the air which has been made available as a result of compressor improvements is utilized and by way of introduction it may be of interest to recall some of the bold experiments which were made in earlier years.

THIS short article is a further contribution to the scries which is being published in AIRCRAFT ENGINEERING to give a general picture of jet engine design as it is today.