A detailed assessment of the effects of 3D radial stacking on the aerodynamic performance of NASA Stage 37 rotor blade

Abstract

Axial compressors are critical components of modern engineering systems, particularly used in applications such as gas turbines, jet engines, and industrial power plants. Compressors are designed to operate efficiently by moving and compressing large volumes of gas or fluid in the axial direction. The primary function of axial compressors is to increase the pressure of the working fluid while maintaining a continuous flow, which is essential for achieving high performance in energy conversion systems. The design and operation of axial compressors are governed by the principles of aerodynamics and thermodynamics. Rotor blades, which rotate around a central axis, impart kinetic energy to the fluid, while stator blades positioned between the rotors, convert this kinetic energy into pressure energy. This staged compression process allows axial compressors to achieve high pressure ratios with minimal energy losses, making them ideal for applications requiring high efficiency. Axial compressors offer advantages such as high mass flow rates, compact designs, and adaptability to various operating conditions. However, during the design phase, factors like blade geometry, flow stability, and mechanical stresses must be carefully considered. These factors are crucial for preventing issues such as flow separation, stall, and surge. Advanced computational tools and experimental methods are frequently employed to optimize the performance of axial compressors, ensuring their efficiency and reliability in demanding applications. This thesis aims to investigate in detail the effects of modifications to the three-dimensional geometry of the rotor blade from the Stage 37 study, designed and tested by the National Aeronautics and Space Administration (NASA). These modifications are achieved by altering the stacking of the two-dimensional airfoils along the radial axis without changing their original two-dimensional design. The primary objective is to comprehensively analyze how these changes influence the performance of the rotor blade comprehensively. The flowpath of the NASA Rotor 37 blade is created using the meridional section view and dimensions provided in the NASA documentation. The airfoils of the rotor blade are also generated based on the geometric parameters shared in the same documentation. These airfoils, designed in two-dimensional sections, are stacked along the radial axis to construct the three-dimensional geometry. After stacking the generated airfoils along the radial axis, a computational mesh must be genrated for conducting CFD analyses of the resulting geometry. Ensuring both the accuracy of the analysis and the computational efficiency requires attention to several critical factors during mesh generation. In regions with complex surfaces or narrow gaps, mesh density should be increased to capture surface curvature and details effectively. However, unnecessary mesh refinement can lead to increased computational cost and time, necessitating a balance between mesh resolution and computational resources. Additionally, an appropriate fine mesh structure is essential in boundary layer regions to accurately analyse flow characteristics and properly calculate y+ values to ensure the effective application of turbulence models. Sudden transitions in mesh size and excessively high aspect ratios should be avoided, as they can adversely affect the stability and accuracy of the solution. Performing a mesh independence study is also crucial to verify that the solution is independent of the mesh configuration. Lastly, the type of mesh (structured, unstructured, or hybrid) should be selected according to the CFD software capabilities and analysis requirements. Considering these factors, the rotor geometry is meshed in ANSYS Turbogrid with a structured mesh, targeting a y+ value of 1 on the blade surface to enhance boundary layer resolution accuracy. The critical step in CFD analysis is ensuring that the results are not influenced by the chosen mesh configuration. A mesh independence study is performed by solving the same physical problem by using meshes with varying densities and comparing the obtained results. The mesh quality and density are incrementally refined in small steps, and specific parameters are recorded for each mesh configuration. In this study, isentropic efficiency and pressure ratio are used as the evaluation metrics. If the results converge to a consistent value as the mesh density increases, the solution is considered independent of the mesh. However, it is essential to balance mesh density with computational cost and time. An optimized mesh density is selected where the results exhibit negligible changes, ensuring computational efficiency without sacrificing accuracy. The mesh independence test is vital for enhancing the reliability of CFD results, particularly in applications involving complex geometries or turbulent flows. This approach improves solution accuracy while avoiding unnecessary computational expenses. Consequently, a mesh with 11.9 million elements is finalized for further analyses in this study. After completing the mesh independence study, a validation study is conducted. For the validation process, the analysis setup is prepared in the ANSYS CFX solver, and the results are compared with the experimental data of Rotor 37. Achieving accurate results during the convergence process requires not only the proper application of boundary conditions but also the selection of an appropriate solver and turbulence model. Therefore, the k-ω shear stress transport solver and the gamma-theta turbulence model, known for their better boundary layer resolution, are chosen. The validation study compares the analysis results with the performance data provided in the literature for the Stage 37 rotor blade at different rotational speeds. Sufficient accuracy is achieved for results at 90% of the design rotational speed, and the study proceeded based on these validated conditions. In this study, the radial stacking of airfoils is performed in two main approaches, each with two distinct directions, resulting in a total of four configurations. The two main stacking approaches involved creating "bow" and "full leaned" geometries. The offset directions are defined along the chordwise and chord normal direction. For each direction, the stacking is performed in two orientations: from the leading edge to the trailing edge and from the trailing edge to the leading edge along the chord, and towards positive rotation and negative rotation directions perpendicular to the chord. In the bow geometries, the hub and tip sections of the rotor blade are fixed, while the airfoil at 50% blade height is offset by a specified amount. In the fully leaned geometries, the hub section is fixed, and the tip section is offset. The offset values for the sections between the fixed and offset sections are calculated proportionally based on their relative height percentages along the blade span. This approach ensured a smooth transition of aerodynamic profiles in the modified geometries. As a result of the study, the effects of radial stacking modifications, which caused changes in the three-dimensional geometry, on performance parameters are observed. The performance parameters considered included the mass flow range, which is indirectly related to the stall margin, near stall the mass flow rate, the choking mass flow rate, the pressure ratio, and the isentropic efficiency. It is found that the choking mass flow rate is associated with the position of shocks, which is influenced by the narrowing of the flowpath. Generally, in geometries offset to upstream, the shock position is also shifted to regions with a wider passage, leading to an increase in the choking mass flow rate. Conversely, in geometries offset to downstream, the shock position shifted towards the narrowing sections of the passage, causing a decrease in the choking mass flow rate. In the bow stacking type, there is no change in the blade height. However, in the full lean geometries, where the tip section is offset, variations in blade height occurred. This change in blade height, being directly related to the energy imparted to the flow, is found to influence the pressure ratio. An increase in peak isentropic efficiency value is observed only in the geometry where the profiles are shifted chordwise from the leading edge to the trailing edge. In other stacking types, losses varied depending on the type and magnitude of the shift. Regarding the mass flow range, an increase is observed only in the full lean geometry with stacking along the chordwise from the trailing edge to the leading edge. This came with a negligible pressure loss and minimal isentropic efficiency drop. This study aims to identify methods that can minimize the time and cost of designing a new aviation engine with minimal differences from an already existing axial compressor design in the industry. By leveraging these approaches, the study seeks to streamline the design process and achieve efficient outcomes in a resource-effective manner.