Optimisation Method for Multistage Compressors

The paper presents an algorithm for seeking an optimal blade configuration for multistage axial-flow compressors. The primary tool behind the algorithm is 3D CFD simulation, augmented by commercial optimisation software. The core of the algorithm involves feeding an initial data vector to the parametric simulation module so as to form a "new" blade geometry, which is then transferred to 3D computational software. The results obtained are further processed in a program that implements the algorithm for seeking the optimum and forms a new input data vector to achieve the set goal. We present a method of parametrically simulation the blade shape, implemented in a software package, making it possible to describe the shape of the compressor blade profiles using a minimum number of variables and to automatically change the shape in the optimisation cycle. The algorithm developed allows the main parameters of compressor operation (efficiency, pressure ratio, air flow rate, etc.) to be improved by correcting the profile shape and relative position of the blades. The algorithm takes into account various possible constraints. We used the method developed to solve practical problems of optimising multistage axial compressors of gas turbine engines for various purposes, with the number of compressor stages ranging from 3 to 15. As a result, the efficiency, pressure ratio and stability margin of gas turbine engines were increased

Interstage Pressures of a Multistage Compressor with Intercooling

This paper considers the criterion of minimum compression work to derive an expression for the interstage pressure of a multistage compressor with intercooling that includes the gas properties, pressure drops in the intercoolers, different suction gas temperatures, and isentropic efficiencies in each compression stage. The analytical expression for the interstage pressures is applied to estimate the number of compression stages and to evaluate its applicability in order to estimate interstage pressures in the operation of multistage compressors, which can be especially useful when their measurements are not available.

An experimental study of rotor-stator wake unsteadiness in a multistage axial compressor

The flow in a multistage axial compressor is highly unsteady, three-dimensional and turbulent. The interaction between compressor blade rows results in rotor/stator wake unsteadiness, which is not typically considered in the computational fluid dynamics (CFD) models. To gain depth and insight into the inner flow mechanism in multistage compressors, specifically the wake variability driven by the rotor/stator and stator/stator interactions, a compound total-pressure pneumatic probe with both high and low response-frequency were designed and manufactured. Unsteady rotor and stator wake measurements between blade rows for the third stage were carried out with this probe installing on a 3-DOF displacement mechanism, to deepen the knowledge of unsteady interactions in the embedded stages of a four-stage low-speed axial compressor. By performing frequency spectrum analysis and ensemble-average methods, higher spectral magnitude of the blade passing frequency (fBPF) and higher root mean square values of total pressure (PtRMS) at both sides of the stator wake region caused by the shedding of upstream boundary layer are revealed. In addition, the high-order harmonics are strengthened by the stator/stator interactions, especially near the blade tip. The individual contributions of rotor geometry variations/interactions of the upstream rotor wakes and the effects of downstream stator potential modulation to the wake variations can be understood.

Methodology and Experience of Primary Design of a Transonic Axial Compressor

The work presents the main provisions underlying the program for axial compressors calculation anddesign. The calculation of head losses and grids deflection capacity is based on the formulas of A. Komarov. Themodel contains empirical coefficients, values of which are selected during verification of the program based onthe tests results of multistage compressors and compressor stages. The main equations and algorithm for pressuresand velocities calculation under the radial equilibrium condition are presented. The use of computer programsbased on these models in the design of a 4-stage gas turbine engine compressor of moderate power with a totalpressure ratio of 3.2 and a given velocity is shown. For the first compressor stage, two variants with differentflow coefficients were compared. Variant #1 is designed with the classic recommendation to approach the samemechanical gas energy at the exit of the stage along the radius. Variant #2 is designed for a smaller flowcoefficient, but in order to ensure radial equilibrium, it was necessary to introduce a significant unevenness ofmechanical energy supply along the radius. Due to the lower kinetic energy, variant 2 has a 1.9% higher stageefficiency. Despite the fact that the loss coefficients of the blade devices are lower for variant #1. The questionremains as to how much the unavoidable mixing losses of variant #2 will reduce its efficiency in the process ofgas mechanical energy equalizing.

Aerodynamic Performance Prediction of Transonic Axial Multistage Compressors Based on One-Dimensional Meanline Method

The one-dimensional meanline method is of great importance for the design and performance prediction of multistage axial compressors. The models adopted in it, such as incidence, deviation and loss, considering real-fluid effects, determine whether the compressors’ operating behavior can be simulated accurately or not. This paper describes an improved meanline stage-stacking approach for the modelling of modern transonic axial multistage compressors. The improvement embodied in this study is mainly focused on deviation and surge margin prediction, which is the result of a combination of the previous models and models’ correction. One of the coefficients in the deviation angle model is corrected. A new surge model, different from the well-known maximum static pressure rise method of Koch and Smith, is introduced into this program and its advantage lies in higher accuracy and direct calculation instead of proposing a judgment criterion. Three well-documented NASA axial transonic compressors are calculated by this meanline method, and the speedlines and aerodynamic parameters are compared with the experimental data to verify the method presented in this paper. A discussion of the result then follows.

Research of a Spatial Flow of a Low-Flow Stage of a Svd-22 Centrifugal Compressor by Computational Fluid Dynamics Methods Using Super-Computer Technologies

This paper provides the results of the study of a spatial flow in a low-flow stage of a SVD-22 centrifugal compressor of computational fluid dynamics methods using the Ansys CFX 14.0 software package. Low flow stages are used as the last stages of multistage centrifugal compressors. Such multistage compressors are widely used in boosting compressor stations for natural gas, in chemical industries. The flow features in low-flow stages require independent research. This is due to the fact that the developed techniques for designing centrifugal compressor stages are created for medium-flow and high-flow stages and do not apply to low-flow stages. Generally at manufacturing new centrifugal compressors, it is impossible to make a control measurement of the parameters of the working process inside the flow path elements. Computational fluid dynamics methods are widely used to overcome this difficulties. However verification and validation of CFD methods are necessary for accurate modeling of the workflow. All calculations were conducted on one of the SPbPU clusters. Parameters of one cluster node: AMD Opteron 280 2 cores, 8GB RAM. The calculations were conducted using 4 nodes (HP MPI Distributed Parallel startup type) with their full load by parallelizing processes on each node.

The optimal gas dynamic design system of industrial centrifugal compressors based on Universal modeling method

We present the modern stage of development of Universal Modeling Method, a complex of mathematical models and software for optimal design of centrifugal compressors - a new version of simplified mathematical model of efficiency and new software for variation calculations of multistage compressors. Based on this numerical calculation complex we have created a method for preliminary design of flow paths of stages - 2D and 3D impellers, vane and vaneless diffusers and return channels. The new, 9th version of its mathematical model features a quasi-3D calculation method of 2D and 3D impellers design, a new principle of pressure characteristic calculation, a new model of vaneless diffusers and much more. “Digital twin of a centrifugal compressor stage” and “3D compressor” software create digital descriptions of the flow part and its solid model (“digital twin”).

Numerical Study of the Penny Cavity Leakage Effects From Variable Guide Vanes in a High Pressure Compressor: Adjustment of Standard-RANS-CFD-Model to a Hybrid LES Calculation

Variable guide vanes are airfoils, normally in the front stages of multistage compressors, that can be restaggered to extend the compressor’s operational range. To allow the variable guide vanes to rotate the design will inevitably include a gap or cavity between the vane’s rotating mounting feature (Penny) and the stationary inner and outer sidewalls. These penny cavities cause additional leakages which impact losses, and airfoil turning, so reducing a compressor’s efficiency and stability. A compressor model which accurately simulates the penny cavity leakage is central to improving the design. This paper presents a study looking into how to accurately include penny cavity leakage effects during the design of multistage compressors. Multiple blade-row RANS-based flow simulations are the current state-of-the-art standard during the design of multistage compressors. However, it is unlikely that such models have the numerical accuracy to simulate the penny cavity leakage effects in detail: firstly the RANS-turbulence model cannot accurately recreate the turbulent mixing which takes place between the leakage and the main flow, and secondly because a typical multiple blade-row mesh is too coarse to resolve the details of the much smaller penny cavity. To compensate for this, the numerical modelling of penny cavities in the design of a compressor would need to be adjusted. On the other hand, a dedicated hybrid LES-model can more accurately simulate the secondary flows with its significant turbulent mixing, at the cost of computational capacity. In this paper, high resolution hybrid LES-simulations have been used as a benchmark to adjust RANS-calculations typical for the design of a multistage compressor. The paper presents the following steps: Using a standard Jet-in-Crossflow test case, a high resolution model was evaluated using both RANS and a hybrid LES model, and compared against measurements. The flow structures were analyzed and compared to measurements of this test case available from literature. These show that the hybrid LES-model performed significantly better than the RANS-model, in being able to predict the jet impact and flow structures. For a second model consisting of a generic compressor variable vane with penny cavity. RANS and hybrid LES simulations were performed with a highly refined mesh in the region of the penny cavity. The modelling is described in detail and the resulting penny cavity effects compared. Finally, the vane with standard mesh and penny cavity was run using RANS-turbulence CFD and compared to the above. From this conclusions were drawn on how to transfer experience from the higher-fidelity turbulence model to a more industry-standard RANS model, which could for instance be used during the design phase of a multistage compressor.