Compressible Direct Numerical Simulation of Low-Pressure Turbines: Part I — Methodology

Modern low pressure turbines (LPT) feature high pressure ratios and moderate Mach and Reynolds numbers, increasing the possibility of laminar boundary-layer separation on the blades. Upstream disturbances including background turbulence and incoming wakes have a profound effect on the behavior of separation bubbles and the type/location of laminar-turbulent transition and therefore need to be considered in LPT design. URANS are often found inadequate to resolve the complex wake dynamics and impact of these environmental parameters on the boundary layers and may not drive the design to the best aerodynamic efficiency. LES can partly improve the accuracy, but has difficulties in predicting boundary layer transition and capturing the delay of laminar separation with varying inlet turbulence levels. Direct Numerical Simulation (DNS) is able to overcome these limitations but has to date been considered too computationally expensive. Here a novel compressible DNS code is presented and validated, promising to make DNS practical for LPT studies. Also, the sensitivity of wake loss coefficient with respect to freestream turbulence levels below 1% is discussed.

Effects of Wet Compression on the Flow Behavior of a Centrifugal Compressor: A CFD Analysis

Wet compression has been emerging as a prominent method for augmenting net power output from land based gas turbine engine. It is proven more effective than the conventional inlet cooling methods. In this method, fine water droplets are injected just upstream of the compressor impeller. These water droplets absorb the latent heat of evaporation during the compression process of gas-water droplet two-phase flow, consequently reducing the temperature rise. Many gas turbine engineers have performed the feasibility and usefulness studies on this wet compression, but physical understanding on the wet compression process is highly lacking, and related compression flow mechanism remains ambiguous. In the present study, a computational fluid dynamics method has been applied to investigate the wet compression effects on a low speed centrifugal compressor. A Langrangian particle tracking method was employed to simulate the air-water droplet two-phase flow. The power saving achieved with different injection ratio of water droplets has been calculated and it is found that significant saving can be obtained with a water droplet injection ratio of above 3%. The vapor mass fraction varies linearly along the streamwise direction, making the assumption for a constant evaporation rate is valid. With the increase in the injection ratio the polytropic index for compression is coming down. The diffuser pressure recovery has been improved significantly with the wet compression; while the total pressure ratio across the impeller does not improve much. Contrary to the expectation, the evaporation rate is found to be coming down with the increase in the compressor mass flow rate. It is observed that the operating point, at which the peak pressure ratio occurs, shift towards higher mass flow rate during wet compression due to the local recirculation region within the vaneless space between the impeller and diffuser.

Extensive Experimental Study of Circumferential Single Groove in an Axial Flow Compressor

Circumferential single-groove casing treatment becomes an interesting topic in recent few years, because it is a good tool to explore the interaction between the groove and the flow in blade tip region. The stall margin improvement (SMI) as a function of the axial groove location has been found for some compressors, such a trend cannot be predicted by steady high-fidelity CFD simulations. Recent efforts show that to catch such a trend, multi-passage, unsteady flow simulations are needed as the stalling mechanism itself involves cross-passage flows and unsteady dynamics. This indicates a need to validate unsteady numerical simulation results. In this paper, an extensive experimental study of a total of fifteen single casing grooves in a low-speed axial compressor rotor is presented, the groove location varies from 0.4% to 98.3% of axial tip chord are tested. The unsteady pressure data both at casing and at the blade wake with different groove locations are measured and processed, including the movement of trajectory of tip leakage flow, the evolution of unsteadiness of tip leakage flow (UTLF), the unsteady spectrum signature during the stall process, and the outlet unsteady flow characteristic along the span. These data provide a case study for validation of the unsteady CFD results, and may be helpful for further interpretation on the stalling mechanism affected by circumferential casing grooves.

Investigation of Unsteady Flow Field in a Low-Speed One and a Half Stage Axial Compressor: Effects of Tip Gap Size on the Tip Clearance Flow Structure at Near Stall Operation

The primary focus of this paper is to investigate the effect of rotor tip gap size on how the rotor unsteady tip clearance flow structure changes in a low speed one and half stage axial compressor at near stall operation (for example, where maximum pressure rise is obtained). A Large Eddy Simulation (LES) is applied to calculate the unsteady flow field at this flow condition with both a small and a large tip gaps. The numerically obtained flow fields at the small clearance matches fairly well with the available initial measurements obtained at the Johns Hopkins University with 3-D unsteady PIV in an index-matched test facility which renders the compressor blades and casing optically transparent. With this setup, the unsteady velocity field in the entire flow domain, including the flow inside the tip gap, can be measured. The numerical results are also compared with previously published measurements in a low speed single stage compressor (Maerz et al. [2002]). The current study shows that, with the smaller rotor tip gap, the tip clearance vortex moves to the leading edge plane at near stall operating condition, creating a nearly circumferentially aligned vortex that persists around the entire rotor. On the other hand, with a large tip gap, the clearance vortex stays inside the blade passage at near stall operation. With the large tip gap, flow instability and related large pressure fluctuation at the leading edge are observed in this one and a half stage compressor. Detailed examination of the unsteady flow structure in this compressor stage reveals that the flow instability is due to shed vortices near the leading edge, and not due to a three-dimensional separation vortex originating from the suction side of the blade, which is commonly referred to during a spike-type stall inception. The entire tip clearance flow is highly unsteady. Many vortex structures in the tip clearance flow, including the sheet vortex system near the casing, interact with each other. The core tip clearance vortex, which is formed with the rotor tip gap flows near the leading edge, is also highly unsteady or intermittent due to pressure oscillations near the leading edge and varies from passage to passage. For the current compressor stage, the evidence does not seem to support that a classical vortex breakup occurs in any organized way, even with the large tip gap. Although wakes from the IGV influence the tip clearance flow in the rotor, the major characteristics of rotor tip clearance flows in isolated or single stage rotors are observed in this one and a half stage axial compressor.

3D Unsteady Multi-Stage CFD Analysis of Combustor-Turbine Hot Streak Migration

With gas temperatures far exceeding the melting point of nickel-base alloys, advanced cooling schemes are essential to meet the desired mission life of turbine airfoils. Naturally, combustion systems produce gas-temperature non-uniformity in the exiting flowfield. Downstream turbine components must be tolerant to the maximum anticipated gas temperatures. On the other hand, excessive use of cooling air reduces engine efficiency and compromises combustor durability. Throughout gas turbine design history it has been the desire of Turbine Aerodynamicists to be able to compute combustor hot streak migration and mixing through multiple turbine airfoil stages. Typically, hot streak migration studies have been performed using (a) mixing-plane models between rotating and stationery domains or (b) unsteady simulations in which the flowpath annulus is represented by a segment containing airfoil counts that are integer multiples in each blade row or (c) Non-Linear Harmonic methods. With the development of highly-parallelized Computational Fluid Dynamic (CFD) codes driving high performance computer clusters simulation of combustor hot streak migration through multiple High Pressure (HP) turbine stages using an unsteady, 360° (full-annulus) model can be achieved. To this end, Honeywell, in collaboration with Numeca Corporation, has performed a study to evaluate the state-of the art for computation of the effect on second-stage HP turbine nozzle metal temperatures of combustor hot streaks migrated through the first-stage of a two-stage HP turbine.

Enhancing the Centrifugal Compressor Performance Map Measurements Through a Developed One–Dimensional Calculation Tool to Analyse Local Flow Phenomena

Centrifugal compressors for automotive turbochargers have large influence on the operation characteristic of combustion engines. Especially the improvement of the surge margin is one of the most important development targets. Thereby, a reliable detection of local flow phenomena within the compressor stage is necessary and a procedure to gain this information from standard measurement data is discussed in this paper. A one–dimensional calculation methodology for a single-stage centrifugal compressor with a vaneless diffuser and casing treatment is presented. The tool calculates the flow properties at the impeller inlet and exit as well as at diffuser exit, based on the measured inlet and outlet data and the geometry information of the compressor. The calculated flow characteristics are plotted within the measured compressor performance map to show local flow parameters. The unsteady recirculation flow within the casing treatment, the inflow angle and the total pressure losses are considered. The tool is validated on different compressor sizes. Thereby the compressor is equipped with static pressure measuring points at the impeller inlet and exit as well as at the diffuser exit. The calculated static pressure correlated well with the measured data with an accuracy of 2 % to 5 % on 95 % of the operating range. In this paper an experimental parameter study is executed in order to improve the surge margin. Thereby the geometry of the diffuser and the casing treatment is varied and the compressor performance is measured on a turbocharger test rig. The calculation of the flow angles and other flow characteristics within the diffuser enables one to find out whether surge is triggered through the diffuser or the impeller.

Aerodynamic Analysis and Multi-Objective Optimization Design of a High Pressure Ratio Centrifugal Impeller

The design of high pressure ratio impellers is a challenging task. SRV2-O, a typical high pressure ratio centrifugal impeller is selected for the research. A good understanding of flow characteristics in the passage of SRV2-O is obtained by using 3D Reynolds-Averaged Navier-Stokes (RANS) solutions upon numerical validation. It confirms that tip leakage flow and shock wave boundary layer interactions produce the primary energy loss in this transonic impeller. A 3D multi-objective aerodynamic optimization and data mining method named BMOE is presented and programmed by integrating a self-adaptive multi-objective differential evolution algorithm SMODE, 3D blade parameterization method based on non-uniformed B-Spline, RANS solver technique and self-organization map (SOM) based data mining technique. Using BMOE, multi-objective aerodynamic design optimization and data mining is performed for SRV2-O. 14 Pareto solutions are obtained for maximizing isentropic efficiency and total pressure ratio of the impeller. Three typical Pareto solutions, Design A with the highest efficiency, Design B with the higher efficiency and larger pressure ratio and Design C with the maximum pressure ratio, are analyzed. Detailed analysis indicates that the aerodynamic performance of optimized designs is greatly improved. Furthermore, by SOM-based data mining on optimization results, trade-off relation between objective functions and parameter influence mechanism on impeller aerodynamic performance are visualized and explored.

Proposal and Experimental Verification of Design Guidelines for Centrifugal Compressor Impellers With Curvilinear Element Blades to Improve Compressor Performance

This study numerically and experimentally examines the effects of applying curvilinear element blades to fully-shrouded centrifugal impellers on the performance of the centrifugal compressor stages. The design suction coefficient of the target impellers was 0.073. Our previous study confirmed that the application of curvilinear element blades could improve the stage efficiency of similar types of centrifugal compressors. However, a detailed explanation of the relation between the stall margin and the application of the curvilinear element blades remains to be given. The purpose of this study is to investigate the effects of using these blades on the impeller flow field and the stall margin in further detail. The curvilinear element blades we developed for centrifugal turbomachinery were defined by the coordinate transformations between a revolutionary flow-coordinate system and a cylindrical coordinate system. All the blade sections in the transferred cylindrical coordinate system were moved and stacked spanwise in accordance with the given “lean profile,” which meant the spanwise distribution profile of movement of the blade sections, to form a new leaned blade surface. The effects of the curvilinear element blades on the impeller flowfield were investigated by conducting numerical simulations using this method. We next considered the optimum design guidelines for impellers with curvilinear element blades. Then we designed a new impeller using these design guidelines and evaluated the performance improvement of a new compressor stage by conducting numerical simulations. As mentioned in several papers, we numerically confirmed that curvilinear element blades with a negative tangential lean profile improved the velocity distribution and stage efficiency because they help to suppress the secondary flows in the impeller. The negative tangential lean mentioned in this paper represents the lean profile in which the blade hub end leans forward in the direction of the impeller rotation compared to the blade shroud end. At the same time, we also found that the stall margin of these impellers deteriorated due to the increase in relative velocity deceleration near the suction surface of the shroud in the forward part of the impeller. Therefore, we propose new design guidelines for impellers with the curvilinear element blades by applying a negative tangential lean to line element blades in which the blade loading of the shroud side in the forward part of the impeller is reduced. We confirmed from the numerical simulation results that the performance of the new compressor stage improved compared to that in the corresponding conventional one. The new design guidelines for the curvilinear element blades were experimentally verified by comparing the performance of the new compressor stage with the corresponding conventional one. The measured efficiency of the new compressor stage was 2.4 % higher than that of the conventional stage with the stall margin kept comparable. A comparison of the measured velocity distributions at the impeller exit showed that the velocity distribution of the new impeller was much more uniform than that of the conventional one.

Investigation of Inlet Bent-Pipe’s Effect on a Turbocharger Compressor: Unsteadiness of Mass Glow and Location of Main Flow Loss

As inlet bent-pipes are often used in vehicle turbocharger compressors, understanding of the inlet bent-pipe’s effect on centrifugal compressor and the mechanism of flow loss is important to improve the turbocharger compressor performance. Experiment and numerical simulation were carried out on a centrifugal compressor with straight pipe and bent-pipe. Numerical simulated compressor performance was compared with test data obtained from compressor flow bench test and to validate the numerical method. The analysis is mainly based on numerical results and it indicated that the inlet bent pipe induces a serious distortion at impeller inlet and increases the risk of blade high-cycle fatigue. The level of distortion changes to be more serious as the operation point of compressor varies from design point to high load. When an impeller channel enters distortion, the larger pressure difference between inlet and outlet makes this channel get clogged and the flow rate through this channel is reduced. At design point the bent pipe mainly influences the flow loss in the inlet pipe and impeller while at high load more serious flow distortion induced by bent-pipe is able to extend its stronger effect to the downstream diffuser and volute.

Numerical and Experimental Investigation of Pulsating Flow Effect on a Nozzled and Nozzleless Mixed Flow Turbine for an Automotive Turbocharger

To achieve better flow guidance into the turbine blades, nozzle vanes were added as an integral part of the stator design. However, the full investigation that directly addresses the comparison between the two turbine arrangements under pulsating flow conditions is still not available in literature. This work represents the first attempt to observe differences, particularly in the circumferential flow angle distribution between both volute arrangements under steady and pulsating flow operating conditions. Experimentally validated Computational Fluid Dynamics (CFD) simulations have been conducted in order to achieve this aim. As the experimental data within the Turbocharger Group at Imperial College are extensive, the simulation procedures are optimized to achieve the best compromise between the computational cost and prediction accuracy. A single operating pressure ratio is selected for the steady and pulsating environment in order to provide consistent comparison for both volute arrangements. The simulation results presented in this work are conducted at the turbine speed of 48000rpm and 60Hz flow frequency for the pulsating flow simulations. The results indicated that there are significant differences in the flow angle behavior for both volutes regardless of the flow conditions (steady or unsteady). It is also found that the differences in flow angle distribution between increasing and decreasing pressure instances during pulsating flow operation is more prominent in the nozzleless volute than its nozzled counterpart.