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This paper aims to comprehensively clarify the research status of thermal transport of supercritical aviation kerosene, with particular interests in the effect of cracking on heat transfer.

A brief review of current research on supercritical aviation kerosene is presented in views of the surrogate model of hydrocarbon fuels, chemical cracking mechanism of hydrocarbon fuels, thermo-physical properties of hydrocarbon fuels, turbulence models, flow characteristics and thermal performances, which indicates that more efforts need to be directed into these topics. Therefore, supercritical thermal transport of n-decane is then computationally investigated in the condition of thermal pyrolysis, while the ASPEN HYSYS gives the properties of n-decane and pyrolysis products. In addition, the one-step chemical cracking mechanism and SST k-ω turbulence model are applied with relatively high precision.

The existing surrogate models of aviation kerosene are limited to a specific scope of application and their thermo-physical properties deviate from the experimental data. The turbulence models used to implement numerical simulation should be studied to further improve the prediction accuracy. The thermal-induced acceleration is driven by the drastic density change, which is caused by the production of small molecules. The wall temperature of the combustion chamber can be effectively reduced by this behavior, i.e. the phenomenon of heat transfer deterioration can be attenuated or suppressed by thermal pyrolysis.

The issues in numerical studies of supercritical aviation kerosene are clearly revealed, and the conjugation mechanism between thermal pyrolysis and convective heat transfer is initially presented.

The purpose of this paper is to numerically investigate the heat transfer performance of aviation kerosene flowing in smooth and enhanced tubes with asymmetric fins at supercritical pressures and to reveal the effects of several key parameters, such as mass flow rate, heat flux, pressure and inlet temperature on the heat transfer.

A CFD approach is taken and the strong variations of the thermo-physical properties as the critical point is passed are taken into account. The RNG k-ε model is applied for simulating turbulent flow conditions.

The numerical results reveal that the heat transfer coefficient increases with increasing mass flow rate and inlet temperature. The effect of heat flux on heat transfer is more complicated, while the effect of pressure on heat transfer is insignificant. The considered asymmetric fins have a small effect on the fluid temperature, but the wall temperature is reduced significantly by the asymmetric fins compared to that of the corresponding smooth tube. As a result, the asymmetric finned tube leads to a significant heat transfer enhancement (an increase in the heat transfer coefficient about 23-41 percent). The enhancement might be caused by the re-development of velocity and temperature boundary layers in the enhanced tubes. With the asymmetric fins, the pressure loss in the enhanced tubes is slightly larger than that in the smooth tube. A thermal performance factor is applied for combined evaluation of heat transfer enhancement and pressure loss.

The asymmetric fins also caused an increased pressure loss. A thermal performance factor ? was used for combined evaluation of heat transfer enhancement and pressure loss. Results show that the two enhanced tubes perform better than the smooth tube. The enhanced tube 2 gave better overall heat transfer performance than the enhanced tube 1. It is suggested that the geometric parameters of the asymmetric fins should be optimized to further improve the thermal performance and also various structures need to be investigated.

The asymmetric fins increased the pressure loss. The evaluation of heat transfer enhancement and pressure loss Results showed that the two enhanced tubes perform better than the smooth tube. It is suggested that the geometric parameters of the asymmetric fins should be optimized to further improve the thermal performance and also various structures need to be investigated to make the results more engineering useful.

The paper presents unique solutions for thermal performance of a fluid at near critical state in smooth and enhanced tubes. The findings are of relevance for design and thermal optimization particularly in aerospace applications.

The purpose of this paper is to numerically study the influence of wall conduction on the heat transfer of supercritical n-decane in the active regenerative cooling channels.

A horizontally placed rectangular pipe with a solid zone and another one without a solid zone were used. A drastic variation of thermo-physical properties was emphatically addressed. After the verification of mesh and turbulence models comparing with the experimental results, a mesh number of 4.5 M and the low Reynolds number SST k-ω turbulence model were chosen. The solution of the governing equations and the acquisition of the numerical results were executed by the commercial software FLUENT 2020 R1.

The numerical results indicate that there is a heat transfer deterioration (HTD) potential for the upper wall, lower wall and sidewall with the decrease of mass flux. Due to wall conduction, the distribution of the fluid temperature at spanwise-normal planes becomes uniform and this feature also takes advantage of the relatively uniform transverse velocity. For the streamwise-normal planes, the low fluid temperature appears close to the upper wall at the region near the sidewall and vice versa for the region near the centre. Undoubtedly, the secondary flow at the cross-section plays a crucial role in this process and the relatively cool mainstream is affected by the vortices.

This study warns that the wall conduction must be considered in the practical design and thermal optimization due to the sensibility of thermo-physical properties to the heat flux. The secondary flow caused by the buoyancy force (gravity) plays a significant role in the supercritical heat transfer and mixed convection heat transfer should be further studied.

NASA's role in aeronautics is, by charter, to improve the usefulness, performance, speed, safety, and efficiency of U.S. civil and military aeronautical vehicles and to preserve U.S. leadership in aeronautical science and technology and its applications. To fill that role, NASA has oriented its aeronautics research and technology (R & T) programme to meet the near‐term and far‐term technology needs of the aviation industry, aircraft operators, government regulatory agencies, and the military services. NASA coordinates closely with those organizations in defining the R & T needs and the objectives for its aeronautics programme. The programme objective of potentially greatest interest to attendees of the International Air Safety Seminar is “To generate technology required for safer, more economical, efficient, fuel‐conservative, and environmentally acceptable air transportation systems to satisfy current and projected national needs.” In the spirit of this international meeting, I should note that certain NASA aeronautical research disciplines include cooperative efforts with the government aeronautical research organizations of several foreign countries.

WHEN talking about airbreathing engines it is now generally understood that they are either turbine engines when the maximum flight Mach number is subsonic or moderately supersonic, or ramjets when the Mach number is definitely high. When trying to meet the propulsion requirements from take‐off to a high enough speed the joint use of both engine types has to be considered. In such case most people would think of the ramjet as taking over the propulsion task from the turbine engine when reaching a certain value of the flight Mach number, or more precisely of the air stagnation temperature, above which the turbine engine is no longer able to operate. The most elementary view is that of presenting it as a limitation in the engine structure, with improvements calling for the use of better materials. Bringing thermo‐dynamics into the picture shows that increased air stagnation temperature results in a deterioration in the cycle efficiency of the turbine engine proper and this may result in the specific fuel consumption of the turbojet becoming higher than that of the ramjet. Such a performance limitation can be shifted to higher Mach numbers while using increased turbine intake temperatures. The consideration of the aerodynamics of the internal flow brings out another type of limitation due to the difficulty of keeping the operating line of the turbojet over the flight profile far enough from the surge limit though within the range of good compresser efficiency. Variable geometry in the compressor and turbine stators may produce some improvement in this respect.

The background of missile costs is discussed. Missiles are new and very costly. Developments in this field have been subjected to political vicissitudes which have often upset long‐term developments. Missile technology is on the frontier of science and there is no background of knowledge to draw on; much basic and expensive research is required. Missile engineering models are complex in detail and assembly, and therefore costly, and constant change occurs while making and testing the model. The complexity and functional requirements of missile parts are running a parallel race with the machines and processes being developed to fabricate the materials required. The usually small runs required in missile production again add to costs. Imposed on all these activities is the requirement that reliability of near 100 per cent is needed and in no case can reliability be allowed to be secondary to cost. The inflight life and shelf conditions for a missile are usually fairly well established and 100 per cent reliability for a short operating life with a long shelf life are the real requirements. There is a considerable tendency to overdesign for reliability. Some costly features of design such as finest finish, closest tolerances and highest strength are carried over by habit from aircraft design and are not always required in missiles. Having examined some causes of high costs, a programme for cost reduction is set out. Costs can be reduced by: (i) earlier freezing of designs making changes only in groups of several changes at wider intervals, (ii) making a more realistic approach to reliability designs, (iii) selecting tolerances in a more analytical manner according to individual needs, (iv) selecting materials on the basis of actual design requirements instead of using the very best materials available even when the short life makes them unnecessary, (v) avoiding tool‐room methods in production engineering, (vi) setting work standards on as many operations as possible and enforcing them to the greatest degree possible, (vii) selecting the best type of workers to make the transition from development models to production missiles as smooth as possible, and (viii) setting up rigid systems and parts designation procedures for handling production parts. Finally, methods of organizing research and development and production for bridging the gap between engineering design and production are proposed.

This study aims to introduce exergy analysis of a three-spool turboprop engine during the complete flight.

In this study, a flight scenario of the aircraft is assumed. Operating parameters of the aircraft and its engine are modelled based on the assumed flight scenario with the aid of a genuine code. And then performance analysis of the engine is performed for each flight path point with the aid of exergy.

At the end of the study, major exergy parameters of the engine are calculated during the complete flight of a cargo aircraft three-spool turboprop engine.

Findings of the study may be beneficial for industry and practitioners to improve performance of the evaluated engine.

To the best of authors’ knowledge, this paper presented the exergy analysis of a three-spool turboprop engine during the complete flight for the first time. It was shown how the exergy destruction rate depends on the altitude and manoeuvre.

THE Combustion and Propulsion Panel of A.G.A.R.D. (Advisory Group on Aeronautical Research and Development to N.A.T.O.) held its third Colloquium at Palermo, Sicily, during the period March 17–21, 1958. Nearly 200 delegates attended the Colloquium representing 10 different N.A.T.O. countries. There were 19 delegates from the U.K. and 12 of these participated either as authors, reviewers of papers or to the discussion. During the five‐day period, 18 invited papers were presented with prepared comments from 45 selected reviewers.

Spiral-wound heat exchangers (SWHEs) are widely used in different industries. In special applications, such as cryogenic (HEs), fluid properties may significantly depend on fluid temperature. This paper aims to present an analytical method for design and rating of SWHEs considering variable fluid properties with consistent shell geometry and single-phase fluid.

To consider variations of fluid properties, the HE is divided into identical segments, and the fluid properties are assumed to be constant in each segment. Validation of the analytical method is accomplished by using three-dimensional numerical simulation with shear stress transport k-ω model, and the numerical model is verified by using the experimental data. Moreover, the HE cost is selected as the main criterion in obtaining the proper design, and the most affordable geometry is selected as the proper design.

The accuracy of different heat transfer and pressure drop correlations is investigated by comparing the analytical and numerical results. The average errors in the calculation of effectiveness, shell-side pressure drop and tube-side pressure drop using the analytical method are 2.1%, 13.9% and 13.3%, respectively. Moreover, the effect of five main geometrical parameters on the SWHE cost is investigated. The results indicate that the effect of longitudinal pitch ratio on the SWHE cost can be neglected, whereas other geometrical parameters have a significant impact on the total cost of the SWHE.

This work contains a versatile and low-cost analytical method to design and rating the SWHEs considering the variable fluid property with consistent shell geometry. The previous studies have introduced complex methods and have not considered the consistency of shell geometry.

This paper proposes new methods of fault detection for fuel systems in order to improve system availability. Novel fault systems are required for environmentally friendly lean burn combustion, but can carry high risk failure modes particularly through their control valves. The purpose of the developed technology is the rapid detection of these failure modes, such as valve sticking or impending sticking, and therefore to reduce this risk. However, sensing valve state is challenging due to hot environmental temperatures, which results in a low reliability for conventional position sensing.

Starting with the business needs elicited from stakeholders, a quality functional deployment process is performed to derive sensing system requirements. The process acknowledges the difference between test-bed and in-service aerospace needs through weightings on requirements and maps these customer requirements to systems performance metrics. The design of the system must therefore optimise the sensor suite, on- and off-board signal processing and acquisition strategy.

Against this systems engineering process, two sensing strategies are outlined which illustrate the span of solutions, from conventional gas path sensing with advanced signal processing to novel non-invasive sensing concepts. While conventional sensing may be feasible within a test cell, the constraints of aerospace in-service operation may necessitate more novel alternatives. Acoustic emission (detecting very high frequency surface vibration waves) sensing technology is evaluated to provide a non-invasive, remote and high temperature tolerant solution. Through this comparison, the considerations for the end-to-end system design are highlighted to be critical to sensor deployment success in-service.

The paper provides insight into different means of addressing the important problem of monitoring faults in combustor systems in gas turbines. By casting of the complex design problem within a systems engineering framework, the outline of a toolset for solution evaluation is provided.

The paper provides three areas of significant contributions: a diversity of methods to diagnosing fuel system malfunctions by measuring changes fuel flow distributions, through novel means, and the combustor exit temperature profiles (cause and effect); the use of analytical methods to support the selection (types and quantities) and placement of sensors to ensure adequate state awareness while minimising their impact on the engine system cost and weight; and an end-to-end data processing approach to provide optimised information for the engine maintainers allowing informed decision-making.