Open Source Axial Compressor Mean-Line Design Tool for Supercritical Carbon Dioxide

2021 ◽  
Author(s):  
Kaden Wells ◽  
Mark G. Turner
Keyword(s):  
Open Source ◽  
Axial Compressor ◽  
Design Tool ◽  
Working Fluid ◽  
Real Property ◽  
Flow Angle ◽  
Total Enthalpy ◽  
Blade Row ◽  
The Mean ◽  
Line Design

Abstract This paper presents an open-source axial compressor design code developed for applications using Supercritical CO2 (S-CO2). Real property tables are generated using REFPROP (Reference Fluid Thermodynamic and Transport Properties Database) linked to MATLAB. These tables are created and are provided for S-CO2 and could be created for any fluid in the database. At this time, only a single-phase fluid has been implemented. The tables are imported into the mean-line code and are interpolated with cubic splines to calculate real properties based on two given properties. The mean-line code is written in Python to allow portability and convenient plotting capability. The inputs are simple ascii files with the overall compressor details, stage data, and an optional IGV file. The code uses the axial flow equations of continuity, energy, and angular momentum in addition to velocity triangles to calculate state properties at every station. A free vortex assumption at each between-blade row station is used to calculate information at hub, pitch, and tip. The input for each stage includes the Mach number and absolute flow angle at the rotor leading edge in addition to the total enthalpy rise across each rotor. Loss coefficients, solidity, aspect ratio and axial spacing are also specified for each blade row along with blockage to account for wakes, boundary layers, and bleed. A hub radius is also specified. These parameters allow for a complete set of realistic inputs for the design of axial compressors using S-CO2 as the working fluid. The output can be used to assess the design and is used as the start of higher fidelity design.

Oxy-Fuel Turbine Through-Flow Design

2014 ◽  
Author(s):  
Majed Sammak ◽  
Srikanth Deshpande ◽  
Magnus Genrup
Keyword(s):  
Mach Number ◽  
Total Pressure ◽  
Tangential Velocity ◽  
Flow Angle ◽  
Power Turbine ◽  
Through Flow ◽  
The Mean ◽  
Line Design ◽  
Reaction Degree ◽  
Relative Flow

The objective of the paper is to present the through-flow design of a twin-shaft oxy-fuel turbine. The through-flow design is the subsequent step after the turbine mean-line design. The through-flow phase analyses the flow in both axial and radial directions, where the flow is computed from hub to tip and along streamlines. The parameterization of the through-flow is based on the mean-line results, so principal features such as blade angles at the mean-line into the through-flow phase should be retained. Parameters such as total inlet pressure and temperature, mass flow, rotation speed and turbine geometries are required for the through-flow modelling. The through-flow study was performed using commercial software — AxCent(™) from Concepts NREC. The rotation speed of the twin-shaft power turbine was set to 7200 rpm, while the power turbine was set to 4800 rpm. The mean-line design determined that the twin-shaft turbine should be designed with two compressor turbine stages and three power turbine stages. The through-flow objective was to study the variations in the thermodynamic parameters along the blade. The power turbine last-stage design was studied because of the importance of determining exit Mach number distribution of the rotor tip. The last stage was designed with damped forced condition. The term ‘damped’ is used because the opening from the tip to the hub is limited to a certain value rather than maintaining the full concept of forced vortex. The study showed the parameter distribution of relative Mach number, total pressure and temperature, relative flow angle and tangential velocity. Through-flow results at 50% span and mean-line results showed reasonable agreement between static pressure, total pressure, reaction degree and total efficiency. Other parameters such as total temperature and relative Mach number showed some difference which can be attributed to working fluid in AxCent being pure CO2. The relative tip Mach number at rotor exit was 1.03, which is lower than the maximum typically allowed value of 1.2. The total pressure distribution was smooth from hub to tip which minimizes the spanwise gradient of total pressure and thus reduces the strength of secondary vortices. The reaction degree distribution was presented in the paper and no problems were revealed in the reaction degree at the hub. Rotor blades were designed to produce a smooth exit relative flow angle distribution. The relative flow angle varied by approximately 5° from hub to tip. The tangential velocity distribution was proportional to blade radius, which coincided with forced vortex design. Through-flow design showed that the mean-line design of a twin-shaft oxy-fuel turbine was suitable.

The Off-Design Analysis of Axial-Flow Compressors

1967 ◽  
Vol 89 (4) ◽  
pp. 453-462 ◽  
Author(s):  
W. Jansen ◽  
W. C. Moffatt
Keyword(s):  
Equations Of Motion ◽  
Velocity Ratio ◽  
Axial Compressor ◽  
Axial Flow ◽  
Pressure Rise ◽  
Flow Angle ◽  
Pressure Losses ◽  
Axial Compressors ◽  
Displacement Thickness ◽  
Blade Row

This paper describes in some detail the methods used for generation of a computer program for analyzing the off-design performance of axial compressors. The purpose of the study was to predict such axial-compressor characteristics as pressure rise and efficiency as a junction of mass flow and rotor speed when only the mechanical geometry of the annulus and blades are given. The aerodynamic analysis is divided into two basic parts. The first part of the analysis involves the numerical solution of the equations of motion in the compressor to find the flow field, temperature and pressure rise, and efficiency of the compressor. The second part concerns the analysis of the blade elements, in particular their performance at various angles of attack. Several effects are included that have a bearing on this performance. The relative Mach number and the axial-velocity ratio across the blade element, together with the shape of the blade and its configuration in the compressor, affect the flow discharging from the blade. Therefore, an elaborate study has been included to find the flow angle at the blade discharge and the pressure losses of the flow passing through the blade row. The calculation procedures have been tested by applying them to a number of compressors for which sufficient experimental data were available. A one-stage, a two-stage transonic, and a 13-stage axial compressor have been investigated and good agreement with experiments has been found provided the correct boundary layer displacement thickness along the annulus wall is used.

Experimental Study of an Axial Compressor Rotor Transfer Function With Non-Uniform Inlet Flow

1978 ◽  
Author(s):  
J. Colpin ◽  
P. Kool
Keyword(s):  
Flow Field ◽  
Three Dimensional ◽  
Axial Compressor ◽  
Hot Wire ◽  
Flow Angle ◽  
Compressor Rotor ◽  
Rotating Blade ◽  
Absolute Flow ◽  
Blade Row ◽  
Upstream Flow

The aim of the investigation is to study the propagation of a non-uniform upstream flow field through a rotating blade row. The flow is investigated using classical pneumatic instrumentation and hot wire anemometry. The latter allows one to determine the average flow values as well as the instantaneous blade-to-blade flow field. These measurements were performed upstream and downstream of a low-speed axial compressor stage rotor. A triangular inlet total pressure distortion was generated with a grid system, movable in the circumferential direction. The hot-wire data were processed with a periodic sampling and averaging technique to obtain the three-dimensional blade-to-blade flow downstream of the rotor at mid-span. The modification of the blade wakes and the existence of large absolute flow angle fluctuations are evidenced for different relative positions of the distortion and stationary hot wire.

The Sears problem for a lifting airfoil revisited - new results

1984 ◽  
Vol 141 ◽  
pp. 109-122 ◽  
Author(s):  
H. M. Atassi
Keyword(s):  
Small Angle ◽  
Potential Flow ◽  
Angle Of Attack ◽  
Nonlinear Dependence ◽  
Mean Flow ◽  
Linear Superposition ◽  
Flow Angle ◽  
Independent Components ◽  
The Mean ◽  
Gust Response

It is shown that for a thin airfoil with small camber and small angle of attack moving in a periodic gust pattern, the unsteady lift caused by the gust can be constructed by linear superposition to the Sears lift of three independent components accounting separately for the effects of airfoil thickness, airfoil camber and non-zero angle of attack to the mean flow. This is true in spite of the nonlinear dependence of the unsteady flow on the mean potential flow of the airfoil. Specific lift formulas are derived and analysed to assess the importance of mean flow angle of attack and airfoil camber on the gust response.

scholarly journals Through-Flow Analysis for Axial-Stage Design Including Streamline-Slope Effects

1993 ◽  
Author(s):  
Theodosios Korakianitis ◽  
Dequan Zou
Keyword(s):  
Mass Flow ◽  
Three Dimensional ◽  
Momentum Equation ◽  
Total Enthalpy ◽  
Streamline Curvature ◽  
Flow Balance ◽  
Through Flow ◽  
Blade Row ◽  
Iteration Loop ◽  
Predictor Corrector

This paper presents a new method to design (or analyze) subsonic or supersonic axial compressor and turbine stages and their three-dimensional velocity diagrams from hub to tip by solving the three-dimensional radial-momentum equation. Some previous methods (matrix through-flow based on the streamfunction approach) can not handle locally supersonic flows, and they are computationally intensive when they require the inversion of large matrices. Other previous methods (streamline curvature) require two nested iteration loops to provide a converged solution: an outside iteration loop for the mass-flow balance; and an inside iteration loop to solve the radial momentum equation at each flow station. The present method is of the streamline-curvature category. It still requires the iteration loop for the mass-flow balance, but the radial momentum equation at each flow station is solved using a one-pass numerical predictor-corrector technique, thus reducing the computational effort substantially. The method takes into account the axial slope of the streamlines. Main design characteristics such as the mass-flow rate, total properties at component inlet, hub-to-tip ratio at component inlet, total enthalpy change for each stage, and the expected efficiency of each streamline at each stage are inputs to the method. Other inputs are the radial position and axial velocity component at one surface of revolution through the axial stages. These can be provided for either the hub, or the mean, or the tip location of the blading. In addition the user specifies the azimuthal deflection of the flow from the axial direction at each radius (or as a function of radius) at each blade row inlet and outlet. By construction the method eliminates radial variations of total enthalpy (work) and entropy at each blade row inlet and outlet. In an alternative formulation enthalpy variations across radial positions at each axial station are included in the analysis. The remaining three-dimensional velocity diagrams from hub to tip, and the radial location of the remaining streamlines, are obtained by solving the momentum equation using a predictor-corrector method. Examples for one turbine and one compressor design are included.

scholarly journals Navier-Stokes and Potential Calculations of Axial Spacing Effect on Vortical and Potential Disturbances and Gust Response in an Axial Compressor

1995 ◽  
Author(s):  
Meng-Hsuan Chung ◽  
Andrew M. Wo
Keyword(s):  
Lower Order ◽  
Spacing Effect ◽  
Axial Compressor ◽  
Interaction Effects ◽  
Moving Frame ◽  
Navier Stokes ◽  
Blade Row ◽  
Vortical Disturbance ◽  
Blade Row Interaction ◽  
Gust Response

The effect of blade row axial spacing on vortical and potential disturbances and gust response is studied for a compressor stator/rotor configuration near design and at high loadings using 2D incompressible Navier-Stokes and potential codes, both written for multistage calculations. First, vortical and potential disturbances downstream of the isolated stator in the moving frame are defined; these disturbances exclude blade row interaction effects. Then, vortical and potential disturbances for the stator/rotor configuration are calculated for axial gaps of 10%, 20%, and 30% chord. Results show that the potential disturbance is uncoupled; the potential disturbance calculated from the isolated stator configuration is a good approximation for that from the stator/rotor configuration for all three axial gaps. The vortical disturbance depends strongly on blade row interactions. Low order modes of vortical disturbance are of substantial magnitude and decay much more slowly downstream than do those of potential disturbance. Vortical disturbance decays linearly with increasing mode except very close to the stator trailing edge. For a small axial gap, lower order modes of both vortical and potential disturbances must be included to determine the rotor gust response.

scholarly journals Effects of Endwall Suction and Blowing on Compressor Stability Enhancement

1989 ◽  
Author(s):  
N. K. W. Lee ◽  
E. M. Greitzer
Keyword(s):  
Flow Injection ◽  
Axial Compressor ◽  
Pressure Rise ◽  
Primary Objective ◽  
Stall Margin ◽  
Casing Treatment ◽  
Blade Row ◽  
Stall Margin Improvement ◽  
Stability Enhancement ◽  
Suction And Blowing

An experimental investigation was carried out to examine the effects on stall margin of flow injection into, and flow removal out of, the endwall region of an axial compressor blade row. A primary objective of the investigation was clarification of the mechanism by which casing treatment (which involves both removal and injection) suppresses stall in turbomachines. To simulate the relative motion between blade and treatment, the injection and removal took place through a slotted hub rotating beneath a cantilevered stator row. Overall performance data and detailed (time-averaged) flowfield measurements were obtained. Flow injection and removal both increased the stalling pressure rise, but neither was as effective as the wall treatment. Removal of high blockage flow is thus not the sole reason for the observed stall margin improvement in casing or hub treatment, as injection can also contribute significantly to stall suppression. The results also indicate that the increase in stall pressure rise with injection is linked to the streamwise momentum of the injected flow, and it is suggested that this should be the focus of further studies.

Vesseg: An Open-Source Tool for Deep Learning-Based Atherosclerotic Plaque Quantification in Histopathology Image

2021 ◽  
Author(s):  
J.M. Murray ◽  
P. Pfeffer ◽  
R. Seifert ◽  
A. Hermann ◽  
J. Handke ◽  
...  
Keyword(s):  
Deep Learning ◽  
Open Source ◽  
Atherosclerotic Plaque ◽  
State Of The Art ◽  
The Mean ◽  
Time Savings ◽  
Plaque Quantification ◽  
Atherosclerosis Research ◽  
Hematoxylin Eosin ◽  
Microscopy Images

Objective: Manual plaque segmentation in microscopy images is a time-consuming process in atherosclerosis research and potentially subject to unacceptable user-to-user variability and observer bias. We address this by releasing Vesseg a tool that includes state-of-the-art deep learning models for atherosclerotic plaque segmentation. Approach and Results: Vesseg is a containerized, extensible, open-source, and user-oriented tool. It includes 2 models, trained and tested on 1089 hematoxylin-eosin stained mouse model atherosclerotic brachiocephalic artery sections. The models were compared to 3 human raters. Vesseg can be accessed at https://vesseg .online or downloaded. The models show mean Soerensen-Dice scores of 0.91±0.15 for plaque and 0.97±0.08 for lumen pixels. The mean accuracy is 0.98±0.05. Vesseg is already in active use, generating time savings of >10 minutes per slide. Conclusions: Vesseg brings state-of-the-art deep learning methods to atherosclerosis research, providing drastic time savings, while allowing for continuous improvement of models and the underlying pipeline.

Evaluation of a Flow Measurement Probe Influence on the Flow Field in High Speed Axial Compressors

2021 ◽  
Author(s):  
Ryosuke Seki ◽  
Satoshi Yamashita ◽  
Ryosuke Mito
Keyword(s):  
Flow Field ◽  
Approximation Method ◽  
High Speed ◽  
Static Pressure ◽  
Axial Compressor ◽  
Pressure Change ◽  
Operating Point ◽  
Blade Row ◽  
Probe Insertion ◽  
Probe Geometry

Abstract The aerodynamic effects of a probe for stage performance evaluation in a high-speed axial compressor are investigated. Regarding the probe measurement accuracy and its aerodynamic effects, the upstream/downstream effects on the probe and probe insertion effects are studied by using an unsteady computational fluid dynamics (CFD) analysis and by verifying in two types of multistage high-speed axial compressor measurements. The probe traverse measurements were conducted at the stator inlet and outlet in each case to evaluate blade row performance quantitatively and its flow field. In the past study, the simple approximation method was carried out which considered only the interference of the probe effect based on the reduction of the mass flow by the probe blockage for the compressor performance, but it did not agree well with the measured results. In order to correctly and quantitatively grasp the mechanism of the flow field when the probe is inserted, the unsteady calculation including the probe geometry was carried out in the present study. Unsteady calculation was performed with a probe inserted completely between the rotor and stator of a 4-stage axial compressor. Since the probe blockage and potential flow field, which mean the pressure change region induced by the probe, change the operating point of the upstream rotor and increase the work of the rotor. Compared the measurement result with probe to a kiel probe setting in the stator leading edge, the total pressure was increased about 2,000Pa at the probe tip. In addition, the developed wake by the probe interferes with the downstream stator row and locally changes the static pressure at the stator exit. To evaluate the probe insertion effect, unsteady calculations with probe at three different immersion heights at the stator downstream in an 8-stage axial compressor are performed. The static pressure value of the probe tip was increased about 3,000Pa in the hub region compared to tip region, this increase corresponds to the measurement trend. On the other hand, the measured wall static pressure showed that there is no drastic change in the radial direction. In addition, when the probe is inserted from the tip to hub region in the measurement, the blockage induced by the probe was increased. As a result, operating point of the stator was locally changed, and the rise of static pressure of the stator increased when the stator incidence changed. These typical results show that unsteady simulations including probe geometry can accurately evaluate the aerodynamic effects of probes in the high-speed axial compressor. Therefore, since the probe will pinpointed and strong affects the practically local flow field in all rotor upstream passage and stator downstream, as for the probe measurement, it is important to pay attention to design the probe diameter, the distance from the blade row, and its relative position to the downstream stator. From the above investigations, a newly simple approximation method which includes the effect of the pressure change evaluation by the probe is proposed, and it is verified in the 4-stage compressor case as an example. In this method, the effects of the distance between the rotor trailing edge (T.E.) and the probe are considered by the theory of the incompressible two-dimensional potential flow. The probe blockage decreases the mass flow rate and changes the operating point of the compressor. The verification results conducted in real compressor indicate that the correct blockage approximation enables designer to estimate aerodynamic effects of the probe correctly.

The Impact of Manufacturing Variations on Performance of a Transonic Axial Compressor Rotor

2018 ◽  
Author(s):  
Marcus Lejon ◽  
Niklas Andersson ◽  
Lars Ellbrant ◽  
Hans Mårtensson
Keyword(s):  
Flow Field ◽  
Geometric Mean ◽  
Pressure Ratio ◽  
Axial Compressor ◽  
Mean Deviation ◽  
Design Intent ◽  
Compressor Rotor ◽  
Measuring Machine ◽  
The Mean ◽  
The Impact

In this paper, the impact of manufacturing variations on performance of an axial compressor rotor are evaluated at design rotational speed. The geometric variations from the design intent were obtained from an optical coordinate measuring machine and used to evaluate the impact of manufacturing variations on performance and the flow field in the rotor. The complete blisk is simulated using 3D CFD calculations, allowing for a detailed analysis of the impact of geometric variations on the flow. It is shown that the mean shift of the geometry from the design intent is responsible for the majority of the change in performance in terms of mass flow and total pressure ratio for this specific blisk. In terms of polytropic efficiency, the measured geometric scatter is shown to have a higher influence than the geometric mean deviation. The geometric scatter around the mean is shown to impact the pressure distribution along the leading edge and the shock position. Furthermore, a blisk is analyzed with one blade deviating substantially from the design intent, denoted as blade 0. It is shown that the impact of blade 0 on the flow is largely limited to the blade passages that it is directly a part of. The results presented in this paper also show that the impact of this blade on the flow field can be represented by a simulation including 3 blade passages. In terms of loss, using 5 blade passages is shown to give a close estimate for the relative change in loss for blade 0 and neighboring blades.

Sign in / Sign up

Export Citation Format

Share Document