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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 study aims at enhancing the performance of a 16-stage axial compressor and improving the operating stability. The adopted approaches for upgrading the compressor are artificial neural network, optimization algorithms and computational fluid dynamics.

The process starts with developing several data sets for certain 2D sections by means of training several artificial neural networks (ANNs) as surrogate models. Afterward, the trained ANNs are applied to the 3D shape optimization along with parametrization of the blade stacking line. Specifying the significant design parameters, a wide range of geometrical variations are considered by implementation of appropriate number of design variables. The optimized shapes are analyzed by applying computational fluid dynamic to obtain the best geometry.

3D optimal results show improvements, especially in the case of decreasing or elimination of near walls corner separations. In addition, in comparison with the base geometry, numerical optimization shows an increase of 1.15 per cent in total isentropic efficiency in the first four stages, which results in 0.6 per cent improvement for the whole compressor, even while keeping the rest of the stages unchanged. To evaluate the numerical results, experimental data are compared with obtained data from simulation. Based on the results, the highest absolute relative deviation between experimental and numerical static pressure is approximately 7.5 per cent.

The blades geometry of an axial compressor used in a heavy-duty gas turbine is optimized by applying artificial neural network, and the results are compared with the base geometry numerically and experimentally.

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.

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.

A stall model to predict the performance of a blade row operating under rotating stall conditions, is proposed.

The experiments were carried out on an isolated rotor row of an axial flow compressor of a radius ratio of 0.66 hub/tip. Wall static pressure tappings were used for measurement of blade row pressure rise. The mass flow rate through the machine was determined from the pressure drop at the intake. Detailed flow measurements were made using a hot wire “V” probe and transducers. An online data acquisition system was used in which data sampling was phase‐locked with respect to stall cell trailing edge.

Measurements indicate that a pressure depression occurs in the stalled region. The assumption of uniform static pressure at the exit of a stalled blade row is not supported by the present work. The assumption of uniform static pressure at the exit of a stalled row together with the assumption that flow in unstalled regions operates at fixed point on the unstalled characteristic leads to the conclusion that total‐to‐static pressure rise during stalled operation is independent of blockage. This view is not supported by the experiments carried out on an isolated rotor.

Additional experimental studies for axial compressors having different rotor and blade geometries and rotor speeds, are required.

The results can be used in the design and operation of axial compressor rotors.

A new stall model is presented in which the behavior during stalled operation with large blockage is different from that during, low blockage.

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.

THE development of an airliner is an extremely complex operation calling for many different professional skills and a vast amount of testing and development facilities.

Active flow control on the NACA 0024 airfoil defined as suction-injection jet at the chord-based Reynolds number of 1.5 × 1e + 5 is studied.

The three-dimensional incompressible unsteady Reynolds-averaged Navier–Stokes equations with the SST k-ω turbulence model are used to study the effects of coflow-jet (CFJ) on the dynamic and static stall phenomena. CFJ implementation is conducted with several momentum coefficients to investigate their turnover. Furthermore, the current work intends to analyze the CFJ performance by varying the Reynolds number and jet momentum coefficient and comparing all states to the baseline airfoil, which has not been studied in prior research investigations.

It is observed that at the momentum coefficient (Cµ) of 0.06, the lift coefficients at low attack angles (up to a = 15) dramatically increase. Furthermore, the dynamic stall at the given Reynolds number and with the lowered frequency of 0.15 is explored. In the instance of Cµ = 0.07, the lift coefficient curve does not show a noticeable stall feature compared to Cµ = 0.05, suggesting that a more powerful stronger jet can entirely control the dynamic stall.

Furthermore, the current work intends to analyze the CFJ performance by varying the jet momentum coefficient and comparing all states to the baseline airfoil, which has not been studied in prior research investigations.

THE shortened runway has become an order of the day. Commercial operators want to offer jet service to communities with small airports. Military services seek to use small, unprepared fields—or no fields at all, just clearings.

Development of the Lockheed supersonic transport has followed the basic philosophy that an advance in air travel in terms of speed and economics should be accompanied by similar advances in aeroplane safety and flying qualities. To achieve these objectives, Lockheed's SST design work has been concentrated for many years on the development of a fixed‐wing design. The present configuration—called a double delta—provides a simple high lift system with low wing loading, excellent low speed stability and control, and large favourable ground effects in landing, with inherent advances in operational simplicity and safety.