VIGV Angle Optimization for Subsonic Axial Flow Compressor

Science & Technology Development Journal - Engineering and Technology

Reducing the loss in the airflow clearance among the compressor blades of the rotor disk and stationary blades (guide vanes) is an urgent issue. Furthermore, additional losses of airflow through the clearances among the blades and airfoil losses are the main cause of reducing the efficiency of an axial flow compressor, especially the blade height is small. With a view towards the efficiency improvement of a multistage axial compressor with a high-pressure ratio, it is necessary to manufacture a highly economical compressor with a variety of compression stages. Airflow in the circulation clearances alternating among compressor blades has viscosity, unstable compression, and quite complex flow structure. This needs to be researched into the design with the assistance of modern software (ANSYS CFX, FlowER, etc.). Although this is an important step in the current design orientation, it requires additional practical elements to perform, especially the problem of optimizing the outer rim,...

In recent decade, industry had started to use intensively 3D simulation in compressor flow path and its components design. Designing an Axial Compressor is more an art than science, because one has to vary the parameters by his intuition. Slight variation in one parameter may lead to large variation in another parameter. There are several options to improve the design but it requires enormous time and effort, as each step of designing and analysis is pretty time consuming. Further this design can be optimized with several options available. This paper reports design and optimization of an axial flow compressor, using a uniform approach to solution on theoretical design problems. Preliminary design was carried out for a set of design parameters and tested for stall at different speed and mass flow rate for given pressure ratio. Modeling was performed using 1D and 2D simulations, as the loss parameters were in limits blade modeling was done using blade profiling and 2D analysis. 3D simulation both CFD and FEA were performed. The results of optimized design shown a close agreement with the theoretical design in terms of mass flow rate at given speed for rated pressure ratio with wide stall margin at low losses in less time for limited given input.

An axial flow compressor is one in which the flow enters the compressor in an axial direction (parallel with the axis of rotation), and exits from the gas turbine, also in an axial direction. The axial-flow compressor compresses its working fluid by first accelerating the fluid and then diffusing it to obtain a pressure increase. In an axial flow compressor, air passes from one stage to the next, each stage raising the pressure slightly. The energy level of air or gas flowing through it is increased by the action of the rotor blades which exert a torque on the fluid which is supplied by an electric motor or a steam or a gas turbine. In this thesis, an axial flow compressor is designed and modeled in 3D modeling software Pro/Engineer. The present design has 30 blades, in this thesis it is replaced with 20 blades and 12 blades. The present used material is Chromium Steel; it is replaced with Titanium alloy and Nickel alloy. Structural analysis is done on the compressor models to verify the strength of the compressor. CFD analysis is done to verify the flow of air

2015

Abstract: Axial flow compressor is one of the most important parts of Gas turbine. In design of Axial flow compressor the work presented comprises of basic flow parameters and dimensions of parts, this makes the further design process quite simple and the results will be helpful to take further changes or improvement at the time of detailed design. The objective of work presented is to design Axial flow compressor by using mean line method for a given mass flow rate and required pressure ratio. The parameters determined also include thermodynamic properties of the working fluid, stage efficiency, number of rotor and stator blades, tip and hub diameters, blade dimensions (chord, length and space) for both rotor and stator, Mach number, flow and blade angles (blade twist). The same parameters are also determined for all five stages. The twist of the blades can be calculated along the blade length at any required number of sections selected by the designers to obtain smooth blade twist...

Archives of Advanced Engineering Science

The present study elaborates on new concepts in designing an axial flow compressor. With the help of an innovative in-house design code, a rotor design is obtained. CFD was used to explore the level of achievement in the performance of the device with the new attributes. In that, the features including local angle of attack, chord length, camber, and thickness were radially adjusted to fulfil flow rate and pressure head requirements at a specific rotational speed, hub, and tip diameters. At the core of this methodology lies the effective use of a lookup table that interrelates airfoil shapes to their corresponding lift coefficients. The most important outcome was the machine attaining the endurance to stall. Furthermore, the rotor geometry is optimized so that the maximum efficiency coincides with the design point of the device, where a low-pitch air motion with a uniform pressure was observed downstream of the rotor, inducing improved aerodynamic conditions. This leads to the minim...

The purpose of this work is to develop and evaluate an inverse optimization algorithm, which designs two-dimensional compressor-blade shapes based on a prescribed pressure or velocity distribution. Thus, the design meets prescribed constraints imposed by the aerodynamic performance requirements that are suggested by the designer/stakeholder. The algorithm is coded using "Visual Basic". Its three main components are: (i) Surface-panel-method flow solver, (ii) Bezier curve-surface-definition routine, (iii) Optimization method. The procedure starts with a given compressor-blade shape C4 as a base-profile. Interesting findings and beneficial conclusions are presented.

Fluid mechanics research international journal, 2018

The purpose of this work is to provide a method for the design of an axial flow compressor stage. This latter represents the front stage of a multistage compressor of industrial gas turbine. The proposed 2D design approach is based on the mean line concept which assumes that mean radius flow conditions prevailing at all other radial stations. The different conservation equations of fluid mechanic; mass, momentum, and energy with ideal gas state equation are applied in conjunction with the NASA empirical relations to compute both incidence and deviation angles at design condition. A FORTRAN computer program was implemented with the inputs such as mass flow rate, tip speed, pressure ratio, ambient temperature and pressure under larger flow coefficient and high reaction ratio. Design process begins with calculation of the channel geometric form from inlet to outlet stage. Thermodynamic properties of the working fluid are determined at rotor inlet, stator outlet and the intermediate station. Detailed geometry of the cascade; chord, pitch, camber and stagger angle is identified for both rotor and stator. The generation of the blade coordinates is performed on the basis of NACA65 profile with circular mean camber line.

This dissertation presents a simulation-based analysis of the influence of hub cavities on the performance of shrouded stator axial compressors. The study investigates the impact of hub cavities on the efficiency of a three-stage axial compressor, utilizing Computational Fluid Dynamics (CFD) tools, including CAD for geometry creation, TurboGrid for meshing, and ANSYS CFX for the simulation. The methodology encompasses the modelling of two configurations: one with hub cavities and one without, allowing for a comparative analysis of compressor performance. Key performance indicators, such as efficiency, static pressure distribution, and leakage flow through streamlines, are analysed in detail. Static entropy plots are used to identify loss areas, while total pressure analysis from hub to shroud offers further insights into the effects of hub cavity presence on the overall flow behaviour. The results indicate that the inclusion of hub cavities influences compressor efficiency and blade loading, which is critical for optimizing axial compressor design. This study provides a comprehensive understanding of the role of hub cavities in axial compressors, offering valuable insights for future advancements in turbine and compressor technologies.

A compressor is a device capable of efficiently transferring energy to the fluid medium so that it can be delivered in large quantities at elevated pressure condition. Compressors have numerous applications ranging from aircraft and process industries to household appliances such as refrigerators and air conditioners. There are numerous types of compressors, each suitable for a particular application. The return channel provides the connection and carries the flow between two stages of a multistage centrifugal compressor. These return channel vanes are located between the two stages through which the fluid passes from one stage to another stage in a centrifugal compressor. Its main function is to rectify the outlet swirling flow from the former stage with possible little flow loss before it enters the impeller of next stage. There are many complexities involve designing the return channel passages, namely; the total pressure loss coefficient, static pressure recovery coefficient and uniform blade loading. The design parameter of a return channel on the performance of a multistage centrifugal compressor is studied through numerical investigations. To find an optimum configuration of compressor with minimization of losses the CFD analysis is performed by changing the design parameter. In the present work a new return channel vane profile is designed and its performance is evaluated using CFD code FLUENT. The aerodynamic studies include the geometry generation of the return channel vanes, meshing, CFD-calculations and performance evaluation. The parameters such as total pressure loss coefficient, static pressure recovery coefficient, vane load distribution on the return channel passages are studied.

2013

In this study an axial compressor consisting of four stages is studied. The endwall of the hub of the last rotor is modified in an automated optimization process in order to reduce the losses and gain surge margin. The endwall is allowed to take non-axisymmetric shapes and is parametrized in such a way, that a groove along the suction side and a fillet are the basic forms. Additionally the lower part of the blade is allowed to vary, while the upper part remains fixed. In order to save calculation time only a set up consisting of stator 3, rotor 4 and stator 4 is calculated during the optimization process. The boundary conditions of the calculations are taken from CFD calculations of the complete four stage rig. For CFD calculations the DLR in-house TRACE-code is used. The optimization tool is AutoOpti, a DLR developed tool, which is based on a genetic algorithm speeded up by surrogate models. In the optimization two operating points are taken into account. One is the aerodynamic design point (ADP), the other one is an operating point near the surge limit (OPSL). The basic forms which evolved during the optimization are a large fillet at the leading edge and the long stretched groove along the suction side of the blade. The optimization shows that a significant reduction of turbulent eddy viscosity and corner stall can be achieved. As the reduction of total pressure losses is limited to the hub region, thus is the overall gain in isentropic efficiency, which rises about 0.2 % in the ADP, for the OPSL an efficiency gain of about 0.5% can be reached. Especially the OPSL shows a strong reduction in secondary flow and correspondingly in corner stall. As only the hub region of the rotor is changed the overall total pressure characteristic of the stage remains basically the same. Using non-axisymmetric endwall shaping can contribute to enhancing axial stages with already highly efficient blading. NOMENCLATURE a Shape factor for Gauss curves ADP Aerodynamic design point CPUh hour of usage of a 2.