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An axial flow turbojet engine in which the mean direction of flow of working fluid past any moving blade is substantially free from radial components comprising a casing; an air intake in said casing; a low‐pressure axial‐flow compressor mounted in said casing, connected directly to said air intake to receive air through it and having a plurality of rows of moving blades whereof the first row has a hub tip ratio between 0·4 and 0·5; a high‐pressure axial flow compressor mounted in said casing, connected directly to said low‐pressure compressor to receive substantially the whole of the air compressed by said low‐pressure compressor and having a plurality of rows of moving blades; combustion equipment mounted in said casing and connected directly to said high‐pressure compressor to receive substantially the whole of the air compressed by said high‐pressure compressor; a single‐stage axial‐flow high‐pressure turbine mounted in said casing, connected directly to said combustion equipment to receive the products of combustion, and drivingly connected to said high‐pressure compressor, the power developed by said high‐pressure turbine being substantially wholly absorbed by said high‐pressure compressor; and a single‐stage axial‐flow low‐pressure turbine mounted in said casing, connected directly to said high‐pressure turbine to receive the exhaust from it and drivingly connected to said low‐pressure compressor, the power developed by said low‐pressure turbine being substantially wholly absorbed by said low‐pressure compressor; in which engine the ratio of the tip diameter of said low pressure turbine to the tip diameter of said first row of moving blades of said low pressure compressor is between 1 and 1·1; and the ratio between the power absorbed by the high‐pressure compressor and the power absorbed by the low‐pressure compressor is between 2 and 2·5 and the tip diameter of said first row of moving blades of said low pressure compressor is greater than the tip diameter of any other row of moving blades of either of said compressors, and the tip diameter of said low pressure turbine is greater than the tip diameter of said high pressure turbine.

Air is sucked from the boundary layer at the rear of an aircraft surface through apertures connected to an air compressor delivering to a jet propulsion expansion nozzle and may comprise the whole of the airflow through the nozzle or alternatively air from in front of the aircraft may also be passed through the compressor, or through some stages only of a multi‐stage compressor. All or part of the air may be heated by fuel injected through combustion nozzles, and the compressor may be driven by a gas turbine on the same shaft or by a reciprocating engine. The combustion air may be preheated by a heat‐exchanger in the exhaust. The jet nozzle may be the sole propulsive means or an airscrew may be mounted forwardly on the compressor shaft. Slots for intake of boundary layer air may be provided at the points at which the boundary layer tends to break away and may be formed as divergent channels or may comprise merely a scries of holes in the aircraft surface. A scries of such slots may be provided with means by which selected slots only are in operation at any given time, and a pitot tube projecting into the boundary layer in front of a slot may be employed to indicate the condition of the boundary layer before that slot. In landing air may be ejected through supplementary nozzles directed forwardly and downwardly to provide aerodynamic braking. The air may be passed for cooling a radiator mounted in the wing or fuselage. Air taken in forwardly of the aircraft may be compressed by a compressor in the fuselage and then distributed, after heating by fuel injection nozzles, to jet tubes in nacelles in the wing through which also passes the boundary layer air As shown in FIG. 9, air extracted at the trailing edges of the wing of an aircraft passes forwardly through a nacelle and through the first stage CI of a compressor. Part then flows directly to a propulsion nozzle at the rear of the nacelle while part passes through a further stage C2 to a combustion chamber fitted with combustion nozzles ch and through a gas turbine T and heat exchanger E, which heats air bypassing the combustion chamber and turbine, to the propulsion nozzle. The turbine T drives the compressor CI, C2 and an airscrew. In FIG. 11, air is passed from slots 1 in the wings through ducts 2 to a central compressor 4 and thence to a reaction nozzle 6. The compressor 4 is driven by a reciprocating motor 5 which also drives an airscrew 3. In FIG. 12, air from slots 7 in the wings passes through compressors 8 in nacelles in the wings and thence to a reaction nozzle 14. The compressor 8 is driven by a turbine 9 supplied through ducts 12 with air admitted centrally at the nose of the aircraft, compressed by a compressor 10, driven by a reciprocating motor 11, and heated by fuel injection nozzles 13. The exhaust from the turbine 9 is to the jet 14. In FIG. 13, air entering a wing 15 through slots 16, 17 passes through the wing spar 18. to the first stage 19 of a compressor in a nacelle beneath the wing. Part then flows by an annular passage 19a to a propulsion nozzle 20. Part passes through a second stage 21 of the compressor to a combustion chamber 22 fitted with combustion nozzles and through a gas turbine 23 to the propulsion nozzle 20. The turbine 23 drives a propeller 24. Specification 512,064 is referred to.

The geometric parameters of the compressor blade have a noteworthy influence on compressor stability, which should be meticulously designed. However, machining inaccuracies cause the blade geometric parameters to deviate from the ideal design, and the geometric deviation exhibits high randomness. Therefore, the purpose of this study is to quantify the uncertainty and analyze the sensitivity of the impact of blade geometric deviation on compressor stability.

In this work, the influence of blade geometric deviation is analyzed based on a subsonic compressor rotor stage, and three-dimensional numerical simulations are used to compute samples with different geometric features. A method of combining Halton sequence and non-intrusive polynomial chaos is adopted to carry out uncertainty quantitative analysis. Sobol’ index and Spearman correlation coefficient are used to analysis the sensitivity and correlation between compressor stability and blade geometric deviation, respectively.

The results show that the compressor stability is most sensitive to the tip clearance deviation, whereas deviations in the leading edge radius, trailing edge radius and chord length have minimal impact on the compressor stability. And, the effects of various blade geometric deviations on the compressor stability are basically independent and linearly superimposed.

This work provided a new approach for uncertainty quantification in compressor stability analysis. The conclusions obtained in this work provide some reference value for the manufacturing and maintenance of rotor blades.

This paper aims to improve the performance of approximate multiplier in terms of peak signal to noise ratio (PSNR) and quality of the image.

The paper proposes an approximate circuit for 4:2 compressor, which shows a significant amount of improvement in performance metrics than that of the existing designs. This paper also reports a hybrid architecture for the Dadda multiplier, which incorporates proposed 4:2 compressor circuit as a basic building block.

Hybrid Dadda multiplier architecture is used in a median filter for image de-noising application and achieved 20% more PSNR than that of the best available designs.

The proposed 4:2 compressor improves the error metrics of a Hybrid Dadda multiplier.

During the last decade the selection of an appropriate gas compressor for a particular application has become increasingly difficult due to the availability of oil lubricated screw machines which provide a logical alternative to the traditional reciprocating compressors.

This paper aims to present the results of a design optimization study for the impeller of a small mixed‐flow compressor. The objective of the optimization is to obtain an impeller geometry that could minimize a cost function based on the specific thrust and the thrust specific fuel consumption of a small turbojet engine.

The design methodology is based on an optimization process that uses a configurational database for various compressor geometries. The database is constructed using design of experiments and the compressor configurations are generated using one‐dimensional in‐house design codes, as well as various tools and programs of the Agile Engineering Design System^®, which is a commercially available turbomachinery design system developed at Concepts NREC. The cost function variations within the design space are represented through a neural network. The optimum configuration that minimizes the cost function is obtained using a direct search optimization procedure.

The optimization study generated a small 86 mm diameter mixed‐flow impeller with a 50° meridional exit angle. The optimized compressor, as well as the engine that it is designed for, were shown to have improved performance characteristics.

Preliminary performance and flow analysis of the optimized impeller show shock structures and possible shock‐boundary layer interactions within the blade passages indicating further geometrical fine tuning may be required based on more detailed computational studies or experimental tests.

A further study including the effect of diffuser is required to carry the results to a more practical level.

The originality and the value of the paper comes mainly from two different aspects: combining various in‐house and commercial turbomachinery design codes in one robust methodology to obtain an optimum mixed‐flow compressor impeller that will maximize the performance requirements of a small unmanned air vehicle (UAV) turbojet engine under restricted size and power conditions; and investigation of the design optimization and analysis of a mixed‐flow compressor that could have potential applications in small jet engines to be used in high‐performance UAV applications. Design optimization studies on this type of compressor are very limited in the open literature. For many years, these compressors have been disregarded because of their bulky design in large‐scale engines. However, as mentioned above, they present a great potential for small‐scale jet engines by supplying enough pressure rise, as well as high mass flow rate compared to their centrifugal counterparts.

The purpose of this paper is to examine the effect of reciprocating compressor speeds and valve conditions on the roor-mean-square (RMS) value of burst acoustic emission (AE) signals associated with the physical motion of valves. The study attempts to explore the potential of AE signal in the estimation of valve damage under varying compressor speeds.

This study involves the acquisition of AE signal, valve flow rate, pressure and temperature at the suction valve of an air compressor with speed varrying from 450 to 800 rpm. The AE signals correspond to one compressor cycle obtained from two simulated valve damage conditions, namely, the single leak and double leak conditions are compared to those of the normal valve plate. To examine the effects of valve conditions and speeds on AE RMS values, two-way analysis of variance (ANOVA) is conducted. Finally, regression analysis is performed to investigate the relationship of AE RMS with the speed and valve flow rate for different valve conditions.

The results showed that AE RMS values computed from suction valve opening (SVO), suction valve closing (SVC) and discharge valve opening (DVO) events are significantly affected by both valve conditions and speeds. The AE RMS value computed from SVO event showed high linear correlation with speed compared to SVC and DVO events for all valve damage conditions. As this study is conducted at a compressor running at freeload, increasing speed of compressor also results in the increment of flow rate. Thus, the valve flow rate can also be empirically derived from the AE RMS value through the regression method, enabling a better estimation of valve damages.

The experimental test rig of this study is confined to a small pressure ratio range of 1.38–2.03 (free-loading condition). Besides, the air compressor is assumed to be operated at a constant speed.

This study employed the statistical methods namely the ANOVA and regression analysis for valve damage estimation at varying compressor speeds. It can enable a plant personnel to make a better prediction on the loss of compressor efficiency and help them to justify the time for valve replacement in future.

Industrial Marketing.

MBA students and participants of MDPs.

It involves marketing of air compressors in particular and industrial equipment in general. It tries to analyse strategies on the framework of market leader strategies to facilitate growth in a challenging business environment in view of the strengths and weaknesses of the firm. It aims to identify the organizational and business model changes that may be required to be implemented in transforming a firm from a marketer of capital goods to a marketer of projects.

To help students/participants evaluate and select marketing strategies for a market leader under challenging business environments as well as identify important organizational and business model changes involved in transition of any firm from selling products to selling projects.

Teaching notes are available for educators only. Please contact your library to gain login details or email support@emeraldinsight.com to request teaching notes.

CSS 8: Marketing.

For an axial-flow compressor rotor, the upstream inflow conditions will vary as the aircraft faces harsh flight conditions (such as taking off, landing or maneuvering) or the whole compressor operates at off-design conditions. With the increase of upstream boundary layer thickness, the rotor blade tip will be loaded and the increased blade load will deteriorate the shock/boundary layer interaction and tip leakage flows, resulting in high aerodynamic losses in the tip region. The purpose of this paper is to achieve a better flow control for tip secondary flows and provide a probable design strategy for high-load compressors to tolerate complex upstream inflow conditions.

This paper presents an analysis and application of shroud wall optimization to a typical transonic axial-flow compressor rotor by considering the inlet boundary layer (IBL). The design variables are selected to shape the shroud wall profile at the tip region with the purpose of controlling the tip leakage loss and the shock/boundary layer interaction loss. The objectives are to improve the compressor efficiency at the inlet-boundary-layer condition while keeping its aerodynamic performance at the uniform condition.

After the optimization of shroud wall contour, aerodynamic benefits are achieved mainly on two aspects. On the one hand, the shroud wall optimization has reduced the intensity of the tip leakage flow and the interaction between the leakage and main flows, thereby decreasing the leakage loss. On the other hand, the optimized shroud design changes the shock structure and redistributes the shock intensity in the spanwise direction, especially weakening the shock near the tip. In this situation, the shock/boundary layer interaction and the associated flow separations and wakes are also eliminated. On the whole, at the inlet-boundary-layer condition, the compressor with optimized shroud design has achieved a 0.8 per cent improvement of peak efficiency over that with baseline shroud design without sacrificing the total pressure ratio. Moreover, the re-designed compressor also maintains the aerodynamic performance at the uniform condition. The results indicate that the shroud wall profile has significant influences on the rotor tip losses and could be properly designed to enhance the compressor aerodynamic performance against the negative impacts of the IBL.

The originality of this paper lies in developing a shroud wall contour optimization design strategy to control the tip leakage loss and the shock/boundary layer interaction loss in a transonic compressor rotor. The obtained results could be beneficial for transonic compressors to tolerate the complex upstream inflow conditions.

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.