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IN the first of these articles it was pointed out that normal supersonic flow can be described theoretically, to a first approximation, by the linearized equation of motion. This has the form of the wave equation and governs first order disturbances to fields of uniform flow; for example, flow past thin wings or slender bodies at small angles of incidence, and flow through ducts of varying cross‐section. In the same way small disturbances in a purely subsonic stream can be described by a linearized equation of motion, which can be reduced to Laplace's equation by contracting the co‐ordinate normal to the direction of flow. Transonic flow, in which regions of both supersonic and subsonic flow occur, is not so easily represented.

This paper aims to study passive control techniques for transonic flow over a backward-facing step (BFS) using square-lobed trailing edges. The study investigates the efficacy of upward and downward lobe patterns, different lobe widths and deflection angles on flow separation, aiming for a deeper understanding of the flow physics behind the passive flow control system.

Large Eddy Simulation and Reynolds-averaged Navier–Stokes were used to evaluate the results of the study. The research explores the impact of upward and downward patterns of lobes on flow separation through the effects of different lobe widths and deflection angles. Numerical methods are used to analyse the behaviour of transonic flow over BFS and compared it to existing experimental results.

The square-lobed trailing edges significantly enhance the reduction of mean reattachment length by up to 80%. At Ma = 0.8, the up-downward configuration demonstrates increased effectiveness in reducing the root mean square of pressure fluctuations at a proximity of 5-step height in the wake region, with a reduction of 50%, while the flat-downward configuration proves to be more efficient in reducing the root mean square of pressure fluctuations at a proximity of 1-step height in the near wake region, achieving a reduction of 71%. Furthermore, the study shows that the up-downward configuration triggers early spanwise velocity fluctuations, whereas the standalone flat-downward configuration displays less intense crosswise velocity fluctuations within the wake region.

The findings demonstrate the effectiveness of square-lobed trailing edges as passive control techniques, showing significant implications for improving efficiency, performance and safety of the design in aerospace and industrial systems.

This paper demonstrates that the square-lobed trailing edges are effective in reducing the mean reattachment length and pressure fluctuations in transonic conditions. The study evaluates the efficacy of different configurations, deflection angles and lobe widths on flow and provides insights into the flow physics of passive flow control systems.

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 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.

The purpose of this paper is to numerically study inflow turbulence effects on the transitional flow in a high pressure linear transonic turbine at the design incidence.

The three‐dimensional (3‐D) compressible turbulent flow in a turbine inlet guide vane is simulated using a finite volume based fluid solver coupled with dynamic large eddy simulation (LES) computations to investigate the effects of varying inflow turbulence length scale and the turbulence intensity on the aero‐thermal flow characteristics and the laminar‐turbulent transition phenomena. The computational analyses are extended to very high exit Reynolds number flow conditions to further study the effect of high exit Reynolds numbers on the transitional behavior of the present flow around the inlet guide vane cascades of the turbine. The calculations are performed with varying degree of inflow turbulence intensity values ranging from 0.8 to 6 percent and the inflow turbulence length scales ranging from one to five percent of pitch for different exit isentropic Mach and Reynolds numbers.

The numerical predictions in comparison with the experimental data demonstrate that the level of inflow turbulence closure provided by the present LES computations offers a reliable framework to predict complex turbulent flow and transition phenomena in high free‐stream turbulence environments of high pressure linear turbines.

This is the first instance in which both artificially modified random flow generation method in association with the dynamic procedure of LES application is employed to represent the realistic inflow turbulence conditions in the high pressure turbine and to resolve the transitional flow in a dynamic approach.

The approximate factorisation‐conjugate gradient squared (AF‐CGS) method has been successfully demonstrated for unsteady turbulent aerofoil flows and transonic inviscid flows in two and three dimensions. The method consists of a conjugate gradient solution of the linear system at each step with the ADI approximate factorisation as a preconditioner. In the present paper the method is adapted to obtain rapid convergence for steady aerofoil flows when compared to the basic explicit method. Modifications to the original method are described, convergence criteria are examined and the method is demonstrated for transonic flow including AGARD test case 9 for the RAE2822 aerofoil.

This paper aims to improve Reynolds-averaged Navier Stokes (RANS) turbulence models using a data-driven approach based on machine learning (ML). A special focus is put on determining the optimal input features used for the ML model.

The field inversion and machine learning (FIML) approach is applied to the negative Spalart-Allmaras turbulence model for transonic flows over an airfoil where shock-induced separation occurs.

Optimal input features and an ML model are developed, which improve the existing negative Spalart-Allmaras turbulence model with respect to shock-induced flow separation.

A comprehensive workflow is demonstrated that yields insights on which input features and which ML model should be used in the context of the FIML approach

Two numerical methods, based on high order finite volume formulations and upwind schemes, are used to compute the two‐ and three‐dimensional flow field in a transonic nozzle. The influence of numerical diffusivity, boundary treatment and mesh structure is explored for inviscid and turbulent configurations. First order computations provide significantly different inviscid results. However, high order methods lead to similar solutions. An explanation of the error generated through the shockwave is proposed in this case. The two‐dimensional interaction of the shock with the thin turbulent boundary layer developing on the bump wall is also presented. Good agreement between both approaches is obtained considering the rapid thickening of the boundary layer due to the shock. Furthermore, the downstream velocity recovery is almost identical. Only slight discrepancies occur in the main flow, near the outer edge of the boundary layer. These seem to be related to the way the turbulence model deals with the free stream turbulence. Finally, preliminary three‐dimensional unstructured turbulent results are presented and discussed.

Rotor performance analysis and design are complex due to the wide variation in flow characteristics. Design tools that can rapidly and accurately compute aerofoil data are needed for rotorcraft design and analysis purposes. The purpose of this paper is to describe a process which has been developed that effectively automates the generation of two‐dimensional (2D) aerofoil characteristics tables.

The process associates a number of commercial software packages and in‐house codes that employ diverse methodologies, including the Navier‐Stokes equation‐solving method, the high‐order panel method and Euler equations solved with the fully coupled viscous‐inviscid interaction (VII) method. The paper describes the development of a general automated generation method that extends from aerofoil shape generation to aerofoil characteristic analysis. The generated data are stored in C81 aerofoil characteristics tables for use in comprehensive rotorcraft analysis codes and rotor blade design. In addition, the methodology could be easily applied for fixed‐wing analysis and design, especially for transonic aircraft.

The method is demonstrated to achieve aerofoil characteristics quickly and accurately in automated process. Calculations for the SC1095 aerofoil section are presented and compared with existing experimental C81 data and previous studies.

The development of C81 tables is of interest to industry as they seek to update their airfoil tables as new designs. Automated processes to achieve this are helpful and applicable.

The paper presents an effective automated process to generate aerofoil characteristics tables quickly, and accurately.

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.