Gas turbine engines for fixed-wing or rotary-wing aircraft are operated in a variety of harsh weather environments ranging from arctic, volcanic zones, to desert conditions. Operation under these degraded conditions leads to the undesired entrainment of complex particulates resulting in drastic performance losses. Hence, there is a critical need to understand the governing mechanisms to inform the development of durable thermal and environmental barrier coatings. The objective of the current work is to present a novel multiscale physics-based approach to study two-phase flows that take into account the underpinning particle transport and deposition dynamics. Sessile droplet models are presented and used to compute the contact angle at high temperatures and compared with experiments. The study also investigates the sensitivity of deposition patterns to the Stokes number and the results identify local vulnerability regions. The analysis suggests that particle size distributions and the initial trajectories of the particles are critically important in predicting the final deposition pattern.

Existing numerical schemes used to solve the governing equations for compressible flow suffer from dissipation errors which tend to smear out sharp discontinuities. Hybrid schemes show potential improvements in this challenging problem; however, the solution quality of a hybrid scheme heavily depends on the criterion to switch between the different candidate reconstruction functions. This work presents a new type of switching criterion (or selector) using machine learning techniques. The selector is trained with randomly generated samples of continuous and discontinuous data profiles, using the exact solution of the governing equation as a reference. Neural networks and random forests were used as the machine learning frameworks to train the selector, and it was later implemented as the indicator function in a hybrid scheme which includes THINC and WENO-Z as the candidate reconstruction functions. The trained selector has been verified to be effective as a reliable switching criterion in the hybrid scheme, which significantly improves the solution quality for both advection and Euler equations.

As the key part for energy amplification of high-power laser systems, disk amplifiers must work in an extremely clean environment. Different from the traditional cleanliness control scheme of active intake and passive exhaust (AIPE), a new method of active exhaust and passive intake (AEPI) is proposed in this paper. Combined with computational fluid dynamics (CFD) technology, through the optimization design of the sizes, shapes, and locations of different outlets and inlets, the turbulence that is unfavorable to cleanliness control is effectively avoided in the disk amplifier cavity during the process of AEPI. Finally, the cleanliness control of the cavity of the disk amplifier can be realized just by once exhaust. Meanwhile, the micro negative pressure environment in the amplifier cavity produced during the exhaust process reduces the requirement for sealing. This method is simple, time saving, gas saving, efficient, and safe. It is also suitable for the cleanliness control of similar amplifiers.

In chemical process engineering, surrogate models of complex systems are often necessary for tasks of domain exploration, sensitivity analysis of the design parameters, and optimization. A suite of computational fluid dynamics (CFD) simulations geared toward chemical process equipment modeling has been developed and validated with experimental results from the literature. Various regression-based active learning strategies are explored with these CFD simulators in-the-loop under the constraints of a limited function evaluation budget. Specifically, five different sampling strategies and five regression techniques are compared, considering a set of four test cases of industrial significance and varying complexity. Gaussian process regression was observed to have a consistently good performance for these applications. The present quantitative study outlines the pros and cons of the different available techniques and highlights the best practices for their adoption. The test cases and tools are available with an open-source license to ensure reproducibility and engage the wider research community in contributing to both the CFD models and developing and benchmarking new improved algorithms tailored to this field.

RANS models remain an attractive turbulence simulation method which could provide some open jet aerofoil interaction analysis at a fraction of the cost of a high-fidelity LES approach. The present work explores the potential and limitations of RANS in this context by simulating an open jet aerofoil noise experiment using the aerospace oriented Menter SST RANS model. This model’s tendency to transition at a critical Reynolds number lower than the experimental value was found to impact the boundary layer development. However, the introduction of a low-Re correction improved the prediction of surface pressure and skin friction, enabling the suction surface separation bubble to be captured. The free shear layer’s virtual origin characteristics exhibited sensitivity to the interaction with the aerofoil, which can be developed into a metric of the interaction. The main challenge for RANS was accounting for the rise in background disturbance level in the working section, which is caused by the high-turbulence intensity in the free shear layers.

In the current work, the effects of design (groove depth and groove width) and operational (temperature and velocity) parameters on aerodynamic performance parameters (coefficient of drag and coefficient of lift) of an isolated passenger car tire have been investigated. The study is conducted by using neural network-based Monte-Carlo analysis on computational fluid dynamics (CFD). The computer experiments are designed to obtain the causal relationship between tire design, operational, and aerodynamic performance parameters. The Reynolds-averaged Navier–Stokes equations-based Realizable K-ε model has been employed to analyze the variations in flow patterns around an isolated tire. The design parameters are varied over wide range and full factorial design, while considering temperature and velocity is completely explored to draw conclusive results. The multi-layer perceptron type neural network with the back-propagation algorithm is trained to map any non-linearity in causal relationships. The sensitivity analysis is performed to find the relationship between control variables and performance indicators. The importance of control variable is determined by both sensitivity and significance analyses and the paired interaction analysis is performed between selected control variables to find the interactive behavior of corresponding variables. The design parameter of groove width with 6.8% and 41% reduction in drag and lift coefficient, respectively, and conventionally overlooked operational parameter of velocity with 4% and 35% impact on drag and lift coefficient, respectively, are found to be the most significant variables. The air trapped between the longitudinal grooves and the road is found to follow the beam theory. The interaction of the groove depth and width is found to be significant with respect to coefficient of lift based on the air beam concept. The interaction of groove width and velocity is found to be significant with respect to both coefficients of lifts and drag.

The scramjet is a rather a new technology and there are many issues related to their operation, especially when it comes to the combustion processes. Combustion in high-speed flows causes various problems such as flame instability and poor fuel–air mixing efficiency. One of the methods used to overcome these problems is to recess a cavity in the combustor wall where a secondary flow is generated. In this study, a computational fluid dynamics (CFD) code is developed to analyse the reacting flow passing through the cavity-based scramjet combustor. The developed code is based on three-dimensional coupled Navier–Stokes and finite rate chemistry equations. An ethylene-air reduced chemical reaction model is used as a fuel–air combination. The Spalart–Allmaras model is utilised for turbulence closure. The non-dimensional form of the flow and chemical reaction equations are discretised using a finite volume method. The Jacobian-Free Newton–Krylov (JFNK) method is used to solve the coupled system of non-linear equations. The JFNK is a matrix-free solution method which improves the computational cost of Newton’s method. The parameters that affect the performance of the JFNK method are studied in the analysis of a scramjet combustor. The influence of the forcing term on the convergence of the JFNK method is studied in the analysis of scramjet combustor. Different upwind flux vector splitting methods are utilised. Various flux limiter techniques are employed for the calculations of higher order flux vectors. The effects of flux vector splitting and flux limiter methods on the convergence and accuracy of the JFNK method are evaluated. Moreover, the variations of the mixing efficiency with fuel injection angles are studied.

A numerical simulation has been carried out to investigate the effects of leading edge blowing upon heat alleviation on the surface of hypersonic vehicles. The initial phase of this work deals with the ability of the present CFD-based techniques to solve hypersonic flow field past blunt-nosed vehicles at hypersonic speeds. Towards this end, the authors selected three re-entry vehicles with published flow field data against which the present computed results could be measured. With increasing confidence on the numerical simulation techniques to accurately resolve the hypersonic flow, the boundary condition at the solid blunt surface was then equipped with the ability to blow the flow out of the solid boundary at a rate of at least 0.01–0.1 times the free stream (ρ∞u∞) mass flow rate. The numerical iterative procedure was then progressed until the flow at the surface matched this new ‘inviscid like’ boundary condition. The actual matching of the flow field at the ejection control surface was achieved by iterating the flow on the adjacent cells until the flow conformed to the conditions prescribed at the control surface. The conditions at the surface could be submitted as a ρ∞u∞ at the surface or could be equipped as a simple static pressure condition providing the desired flow rate. The comparison between the present engineering approach and the experimental data presented in this study demonstrate its ability to solve complex problems in hypersonic.

There is a wide variety of CFD grid types including Cartesian, structured, unstructured and hybrids, as well as, numerous methodologies of combining these to reduce the time required to generate high-quality grids around complex configurations. If the grid methodologies were implemented in different codes, they should be written in such a way as to obtain the maximum performance from the available computer resources. A common interface should also be required to allow for ease of use. However, it is very time consuming to develop, maintain and add extra functionally to different codes. This paper examines the possibility of taking an existing CFD solver, the Helicopter Multi-Block (HMB) CFD method, and implementing a new grid type while reusing as much as possible the original code base. The paper presents some of the challenges encountered in extending the code which was written for a single mesh type, to a more flexible solver that is still computationally efficient but can cope with a variety of grid types.

The next generation of civil large aero-engines will employ greater bypass ratios compared with contemporary architectures. This results in higher exchange rates between exhaust performance and specific fuel consumption (SFC). Concurrently, the aerodynamic design of the exhaust is expected to play a key role in the success of future turbofans. This paper presents the development of a computational framework for the aerodynamic design of separate-jet exhaust systems for civil aero-engines. A mathematical approach is synthesised based on class-shape transformation (CST) functions for the parametric geometry definition of gas-turbine exhaust components such as annular ducts and nozzles. This geometry formulation is coupled with an automated viscous and compressible flow solution method and a cost-effective design space exploration (DSE) approach. The framework is deployed to optimise the performance of a separate-jet exhaust for very-high-bypass ratio (VHBR) turbofan engine. The optimisations carried out suggest the potential to increase the engine’s net propulsive force compared with a baseline architecture, through optimum exhaust re-design. The proposed method is able to identify and alleviate adverse flow-features that may deteriorate the aerodynamic behaviour of the exhaust system.

Minimally invasive surgery is a developing direction of modern medicine. With the successful development of controllable capsule endoscopies, capsule robots are very popular in the field of gastrointestinal medicine. At present, the study of intestinal robots is aimed at the pipeline environment of a single-phase liquid flow. But there exist food residues (i.e. solid particles) or liquid foods in the actual intestine, so intestinal fluid should be liquid–solid or liquid–liquid two-phase mixed fluid. For inner spiral capsule robots with different internal diameters and outer spiral capsule robots, using computational fluid dynamics (CFD) method, the operational performance indicators (i.e. axial thrust force, circumferential resisting moment and maximum pressure to pipeline wall) of spiral capsule robots are numerically calculated in the liquid–solid or liquid–liquid two-phase mixed fluid. By the orthogonal experimental optimization method, the optimum design of spiral capsule robots is obtained in the liquid–solid mixed fluid. The experimental verification has been also carried out. The results show that in the liquid–solid two-phase fluid, the axial thrust force and circumferential resisting moment of the spiral capsule robots decrease with the increase of the size or concentration of solid particles. In the same liquid–solid or liquid–liquid mixed fluid, the operational performance indicators of outer spiral robots are much higher than those of inner spiral robots, and the operational performance indicators of inner spiral robots with bigger internal diameters are higher than those with smaller internal diameters. Adding solid particles of high concentration in the pipeline containing liquid will reduce the drive performance of spiral capsule robots, but adding another liquid of high viscosity will improve the drive performance of spiral capsule robots.

The laminar flow over a slender delta wing at incidence has been extensively studied both experimentally and theoretically using vortex sheet methods. These vortex sheet methods have generally been successful apart from the prediction of the secondary boundary-layer separation induced by the primary vortex. This paper revisits the problem using computational fluid dynamics (CFD) and focusses on the effects of the secondary flow separation. The modelling approach is briefly summarised, and the results are compared with flow measurements and results from vortex sheet methods. The computations show very good agreement with measurements for the surface pressures and total head contours. The results help to understand the complex structure of the leading edge vortex flow, and the associated secondary separation of the boundary layer. They indicate that inviscid mechanisms dominate the larger scale features, and highlight a possible mechanism for the development of an instability in the leading edge vortex sheet.

Full-scale simulations of a (Magnetorheological) MR damper are carried out for revealing its hysteretic behaviors associated with implementation of semi-active control using the routine of computational fluid dynamics. By virtue of the structural symmetry of the MR damper, a two-dimensional configuration for finite element simulation is built up. Herschel-Bulkley model is employed to represent the property of the MR fluid, of which the control parameters and their relevances to the input current are addressed. Typical cases involving sinusoidal and irregular displacements, steady and transient currents loaded upon the MR damper are investigated. Numerical investigations reveal that the damper force has a positive correlation with input current, excitation amplitude and excitation frequency. The full-scale simulation is proved to exhibit a sound accuracy through the validation of experimental data. It provides a logical manner revealing the true performance of MR dampers under desirable operating modes in practice, and can be readily integrated with the gain design of the associated semi-actively controlled structure. This progress bypasses the technical challenge inherent in the traditional tests with low-frequency cyclic loadings due to the limitation of experimental setup. Besides, comparative study between two-dimensional and three-dimensional configuration simulations of the MR damper shows that former has a better applicability, which can be carried out on a low-cost platform.

This study investigates numerically the performance of applying aerospike nozzle in a hydrogen peroxide mono-propellant propulsion system. A set of governing equations, including continuity, momentum, energy and species conservation equations with extended k-ε turbulence equations, are solved using the finite-volume method. The hydrogen peroxide mono-propellant is assumed to be fully decomposed into water vapor and oxygen after flowing through a catalyst bed before entering the nozzle. The aerospike nozzle is expected to have high performance even in deep throttling cases due to its self-compensating characteristics in a wide range of ambient pressure environments. The results show that the thrust coefficient efficiency (Cf,η) of this work exceeds 90% of the theoretical value with a nozzle pressure ratio (PR) in the range of 20 ~ 45. Many complex gas dynamics phenomena in the aerospike nozzle are found and explained in the paper. In addition, performance of the aerospike nozzle is compared with that of the bell-shape nozzle.

The rotorcraft industry needs Virtual Engineering first to ensure decisions made early in the life-cycle, at the requirements capture and preliminary design phases for example, are reliably informed. Later, in design, development and qualification, Virtual Prototypes can become the centre of attention for critical reviews and, ultimately, certification itself. A significant challenge is to ensure that model fidelity is good enough, not only for supporting design decisions but also in establishing requirements based on sufficiently mature technologies. This international conference, Rotorcraft virtual engineering; supporting life-cycle engineering through design and development, test and certification and operations and co-sponsored by the RAeS/AHS/A3F/DGLR/AIDAA addressed these themes and this paper reviews and assesses the value of the various contributions.

A new wing-tip concept with morphing upper surface and interchangeable conventional and morphing ailerons was designed, manufactured, bench and wind-tunnel tested. The development of this wing-tip model was performed in the frame of an international CRIAQ project, and the purpose was to demonstrate the wing upper surface and aileron morphing capabilities in improving the wing-tip aerodynamic performances. During numerical optimisation with ‘in-house’ genetic algorithm software, and during wind-tunnel experimental tests, it was demonstrated that the air-flow laminarity over the wing skin was promoted, and the laminar flow was extended with up to 9% of the chord. Drag coefficient reduction of up to 9% was obtained when the morphing aileron was introduced.

The geometry of a planar converging-diverging nozzle operating with dry air in dilute gas conditions is modified by the introduction of a small recessed step at the throat section. Pressure measurements along the nozzle axis, schlieren visualisations and numerical simulations are performed to investigate the influence of the recessed step on the supersonic flow-field. In the experiments, the height of the recessed step is 0.1 mm and the nozzle height at the throat is 10 mm. Numerical simulations examine also 0.05 mm and 0.2 mm step heights. From the numerical simulations, the flow Mach number at the step location is 1.04 and the Reynolds number computed using the sonic conditions and the throat half-height is Re = 3.73 × 105. A perturbation wave pattern originates from the step, which results in a perturbation of the measured pressure profile close to the throat section. In the diverging portion, sufficiently far from the throat section, the pressure profile of the recessed-step nozzle matches the one measured in the clean configuration.