Experimental Investigation of a Stall Prediction System Using a Transonic Multistage Compressor

2007 ◽  
Author(s):  
Tomofumi Nakakita ◽  
Masahiro Kurosaki ◽  
Yukio Kamiyoshihara
Keyword(s):  
Axial Compressor ◽  
Prediction System ◽  
Test Results ◽  
Stall Margin ◽  
Pressure Transducers ◽  
Stator Vane ◽  
Multistage Compressor ◽  
Stall Warning ◽  
Stall Control ◽  
Multistage Axial Compressor

This paper describes the test results of a stall prediction system using a transonic multistage axial compressor. The test results show that there is a clear relationship between the stall margin and the stall-warning index measured at the first stage. The stall warning index is derived from auto correlation of casing wall pressure signals above the rotor tip. In the tests, the compressor installed with pressure transducers on the casing wall near the leading edges of the first three rotors was forced to stall while operating at a constant speed by closing the discharge valve. The test results where the stator vane settings of the first three stages were changed show that load distribution among stages does not have significant effects on the stall margin vs. stall-warning index relationship. Using some smoothing technique, undesired time variation of the stall-warning index can be reduced to the level necessary for practical active stall control with allowable response time delay.

scholarly journals МЕТОД РАСЧЁТА ТЕРМОГАЗОДИНАМИЧЕСКИХ ПАРАМЕТРОВ ТУРБОВАЛЬНОГО ГАЗОТУРБИННОГО ДВИГАТЕЛЯ НА ОСНОВЕ ПОВЕНЦОВОГО ОПИСАНИЯ ЛОПАТОЧНЫХ МАШИН. ЧАСТЬ II. ОПРЕДЕЛЕНИЕ ПАРАМЕТРОВ СТУПЕНЕЙ И МНОГОСТУПЕНЧАТЫХ КОМПРЕССОРОВ

2019 ◽  
pp. 18-28 ◽  
Author(s):  
Людмила Георгиевна Бойко ◽  
Александр Евгеньевич Демин ◽  
Наталия Владимировна Пижанкова
Keyword(s):  
Gas Turbine ◽  
Calculation Method ◽  
Axial Compressor ◽  
Flow Path ◽  
Dynamic Parameters ◽  
Operating Characteristics ◽  
Program Complex ◽  
Gas Dynamic ◽  
Multistage Compressor ◽  
Multistage Axial Compressor

Gas Turbine Engine (GTE) operating characteristics such as thrust (or power), specific fuel consumption and other cycle parameters on different regimes, can be determined by engine modeling and applying correspondent calculation method. Its accuracy is the function of the engine’s element maps definition precision. So these maps representations influence for engines investigation results significantly. Main points and equation system for engine performances calculation method were represented in Part I of this article. The method gives an opportunity for the flow path thermodynamical parameters and engine integral values analyzing by using multistage axial blade machines blade-to-blade descriptions. The compressor and gas turbine and parameters are getting by special program modules, adding to the engine operating characteristics investigation program complex. These modules use the flow path and cascade middle radius geometrical parameters as the data for calculation. The goal of this article is the representation of the method for axial stages and multistage compressors performances definition. The calculation technique is based on one-dimensional (1D) multistage axial compressor flow description. Proposed 1D flow analysis method allows to get the multistage axial compressor maps taking into account the blade-to-blade gaps flow bleeding and by-pass. The method including is founded on the thermal and gas dynamic equations and turbomachinery theoretical dependences and empirical functions for losses and deviation angles determination. Besides, the representing method allows to calculate gas dynamic parameters, velocity triangles, angles of attack, evaluate their deviations from optimal values, hydraulic losses. Also, it can show accordance of stages working on different regimes, find the stage, which is a reason for compressor instability, and stall margin. This method can be used in GTE mathematic simulation, founded on blade-to-blade description multistage blade machines or also in multistage compressor designing. The proposed method gives the opportunity to control the stator variable vanes stagger angles control and to analyze its influence for stage and multistage compressor gas dynamic parameters and maps.

scholarly journals Influence of Blade Aspect Ratio on Axial Compressor Efficiency

2019 ◽  
Vol 3 ◽  
pp. 639-652 ◽  
Author(s):  
Markus Peters ◽  
Tobias Schmidt ◽  
Peter Jeschke
Keyword(s):  
Aspect Ratio ◽  
Numerical Study ◽  
Axial Compressor ◽  
Compressor Cascade ◽  
Secondary Loss ◽  
First Case ◽  
Lower Reynolds Number ◽  
Stator Vane ◽  
Negative Effect ◽  
Multistage Compressor

A numerical study on the influence of compressor blade aspect ratio on profile and secondary loss has been conducted. In order to more accurately estimate the change in secondary loss, a new analytical model has been developed. The aspect ratio has been increased by reducing blade chord while maintaining blade height and solidity. A simplified compressor cascade geometry and an engine-like HPC stage geometry (rotor blade and stator vane) have been analysed with 3D CFD simulations. For these simulations, the solver TRACE has been used together with the k-ω turbulence model and a Low-Reynolds approach. A negative effect of increased aspect ratio on profile loss due to the lower Reynolds number has been observed as expected from literature. Moreover a decrease of secondary loss at increased aspect ratio due to smaller endwall regions has been noticed. While this effect is also well known, a significant influence of the assumptions regarding the incoming boundary layer thickness has been observed based on the cascade simulations. This leads to the conclusion that changing the aspect ratio of all blades and vanes of a multistage compressor causes a much stronger decrease in secondary loss per blade row than changing the aspect ratio of a single rotor or stator within the compressor. In literature so far only the first case is considered in common loss correlations. However considering the latter would increase the accuracy of secondary loss estimation for a non-uniform change in aspect ratio within a compressor.

scholarly journals Simulation of Casing Treatments of a Transonic Compressor Stage

2008 ◽  
Vol 2008 ◽  
pp. 1-10 ◽  
Author(s):  
M. Hembera ◽  
H.-P. Kau ◽  
E. Johann
Keyword(s):  
Three Dimensional ◽  
Axial Compressor ◽  
Navier Stokes ◽  
Compressor Stage ◽  
Flow Solver ◽  
Stall Margin ◽  
Transonic Compressor ◽  
Flow Mechanisms ◽  
Multistage Compressor ◽  
Casing Treatments

This article presents the study of casing treatments on an axial compressor stage for improving stability and enhancing stall margin. So far, many simulations of casing treatments on single rotor or rotor-stator configurations were performed. But as the application of casing treatments in engines will be in a multistage compressor, in this study, the axial slots are applied to a typical transonic first stage of a high-pressure 4.5-stage compressor including an upstream IGV, rotor, and stator. The unsteady simulations are performed with a three-dimensional time accurate Favre-averaged Navier-stokes flow solver. In order to resolve all important flow mechanisms appearing through the use of casing treatments, a computational multiblock grid consisting of approximately 2.4 million nodes was used for the simulations. The configurations include axial slots in 4 different variations with an axial extension ranging into the blade passage of the IGV. Their shape is semicircular with no inclination in circumferential direction. The simulations proved the effectiveness of casing treatments with an upstream stator. However, the results also showed that the slots have to be carefully positioned relative to the stator location.

A Novel Stall Warning Indicator: Part II — Underlying Flow Mechanism

2019 ◽  
Author(s):  
Mario Eck ◽  
Roland Rückert ◽  
Marc Lehmann ◽  
Dieter Peitsch
Keyword(s):  
Axial Compressor ◽  
Propagation Speed ◽  
Flow Coefficient ◽  
Spectral Signature ◽  
The Novel ◽  
Time Resolved ◽  
Pressure Transducers ◽  
Flow Disturbances ◽  
Stall Warning ◽  
Rotating Instability

Abstract The aim of the present paper is to improve the physical understanding of flow irregularities in the blade passing signal of turbomachinery rotors, since the novel stall warning method presented in part I is based upon those irregularities. For this purpose, a complementary instrumentation was used in a single stage axial compressor. A set of pressure transducers evenly distributed along the circumference surface mounted in the casing near the rotor tip leading edges is measuring the time-resolved wall pressures simultaneously to an array of transducers recording the chord-wise static pressures. The latter allows for plotting quasi-instantaneous 2D-pressure contours. Any occurring flow disturbances causing the before mentioned irregularity can later be classified using validated frequency analysis methods being applied to the data from the circumferential sensors. While leaving the flow coefficient constant, a continuously changing number of prestall flow disturbances appears to be causing the very spectral signature which is known from investigations on Rotating Instability. Any arising number of disturbances is matching a specific mode order to be found within the spectral signature. While the flow coefficient is reduced the propagation speed of prestall disturbances increases linearly as the speed seems to be independent from the clearance size. Data taken beyond the stalling limit demonstrate a complex superposition of stall cells and flow disturbances which the title “prestall disturbance” therefore doesn’t fit to precisely any more. Different convection speeds allow the phenomena to be clearly distinguished from each other.

Deposition Pattern Analysis on a Fouled Multistage Test Compressor

2020 ◽  
Author(s):  
Alessio Suman ◽  
Alessandro Vulpio ◽  
Nicola Casari ◽  
Michele Pinelli ◽  
Rainer Kurz ◽  
...  
Keyword(s):  
Flow Field ◽  
Particle Deposition ◽  
Axial Compressor ◽  
Rotor Blades ◽  
Deposition Patterns ◽  
Compressor Inlet ◽  
Multistage Compressor ◽  
Multistage Test ◽  
Favorable Conditions ◽  
Multistage Axial Compressor

Abstract Compressor fouling is one of the main causes of gas turbine performance degradation. Microsized particles adhere to the blade surfaces increasing the surface roughness and modifying the airfoil shape. In the present work, the contamination of the Allison 250 C18 multistage compressor engine with four sorts of micrometric dust has been provided. The tests were performed changing the relative humidity at the compressor inlet and the unit rotational speed. After each test, a photographic inspection of the internal fouled parts has been realized and the digital pictures have been analyzed employing an image processing package. The deposits build-up of stator vanes and rotor blades have been post-processed and the most affected regions of each compressor stage have been highlighted. Besides, a numerical simulation of the machine has been performed. The numerical flow field has been used to highlight the blade regions which show the most favorable conditions for particle deposition. A theoretical model has been applied to the flow field to simulate the particle deposition. The combination of the deposition model with the results of the CFD simulations gives the chance to better understand the experimentally-founded deposition patterns. Those results have been finally compared to the pictures of the patterns. The possibility to detect and measure the deposition patterns on a rotating test rig and the comparison with models and experiments gave the possibility to assess in detail the particle deposition phenomenon on a multistage axial compressor flow path.

scholarly journals Generalized Multistage Compressor Characteristics

1960 ◽  
Author(s):  
G. L. Mellor
Keyword(s):  
Mach Number ◽  
Axial Compressor ◽  
Flow Coefficient ◽  
Restrictive Assumption ◽  
Characteristic Shape ◽  
Design Performance ◽  
General Properties ◽  
Multistage Compressor ◽  
Stage Characteristic ◽  
Multistage Axial Compressor

Presented below is a method of calculating the off design performance of multistage axial compressor characteristics that differs in concept from previous methods. A set of resulting calculations has certain general properties, namely, that it is independent of absolute stage work and flow coefficients, number of stages, and annulus geometry. Each set depends only on a stage characteristic “shape” which is normalized so that the work coefficient is unity when the flow coefficient is unity. The principle restrictive assumption that has been made is that the effect of Mach number on the density is negligible.

Deposition Pattern Analysis On a Fouled Multistage Test Compressor

2021 ◽  
Author(s):  
Alessio Suman ◽  
Alessandro Vulpio ◽  
Nicola Casari ◽  
Michele Pinelli ◽  
Rainer Kurz ◽  
...  
Keyword(s):  
Flow Field ◽  
Particle Deposition ◽  
Axial Compressor ◽  
Rotor Blades ◽  
Deposition Patterns ◽  
Compressor Inlet ◽  
Multistage Compressor ◽  
Multistage Test ◽  
Favorable Conditions ◽  
Multistage Axial Compressor

Abstract Compressor fouling is one of the main causes of gas turbine performance degradation. Microsized particles adhere to the blade surfaces increasing the surface roughness and modifying the airfoil shape. In the present work, the contamination of the Allison 250 C18 multistage compressor engine with four sorts of micrometric dust has been provided. The tests were performed changing the relative humidity at the compressor inlet and the unit rotational speed. After each test, a photographic inspection of the internal fouled parts has been realized and the digital pictures have been analyzed employing an image processing package. The deposits build-up of stator vanes and rotor blades have been postprocessed and the most affected regions of each compressor stage have been highlighted. Besides, a numerical simulation of the machine has been performed. The numerical flow field has been used to highlight the blade regions which show the most favorable conditions for particle deposition. A theoretical model has been applied to the flow field to simulate the particle deposition. The combination of the deposition model with the results of the CFD simulations gives the chance to better understand the experimentally-founded deposition patterns. Those results have been finally compared to the pictures of the patterns. The possibility to detect and measure the deposition patterns on a rotating test rig and the comparison with models and experiments gave the possibility to assess in detail the particle deposition phenomenon on a multistage axial compressor flow path.

Mixed-Flow Compressor Stage Design and Test Results With a Pressure Ratio of 3:1

1987 ◽  
Vol 109 (4) ◽  
pp. 513-519 ◽  
Author(s):  
D. S. Musgrave ◽  
N. J. Plehn
Keyword(s):  
Additional Data ◽  
Pressure Ratio ◽  
Axial Compressor ◽  
Test Results ◽  
Mixed Flow ◽  
Stage Design ◽  
History Of ◽  
Flow Compressor ◽  
Tip Radius ◽  
Multistage Axial Compressor

This paper presents a brief history of mixed-flow compressors, possible applications, and the design and measured performance of a recently tested 3:1 pressure-ratio stage. The stage is intended to run behind a multistage axial compressor; it has an envelope radius only 9.4 percent greater than the rotor tip radius. A tandem cascade diffusing system is used to promote flow range and thus aid matching to the axial stages. Compressor maps from the rig test are presented along with additional data (from static taps and exit rakes) that characterize the behavior of various elements of the stage.

scholarly journals Impact Investigation of Stator Seal Leakage on Aerodynamic Performance of Multistage Compressor

2021 ◽  
Vol 9 ◽  
Author(s):  
Qi Wang ◽  
Zhou Zhang ◽  
Qingsong Hong ◽  
Lanxue Ren
Keyword(s):  
Mass Flow Rate ◽  
Mass Flow ◽  
Flow Rate ◽  
Axial Compressor ◽  
Flow Distribution ◽  
Aerodynamic Performance ◽  
Flow Losses ◽  
Multistage Compressor ◽  
The Impact ◽  
Multistage Axial Compressor

In this paper, a numerical model based on the mass flow rate of seal leakage is presented, and a 3D numerical method of a multistage axial compressor with good engineering practicability is established. Validation consists of modeling a nine-stage axial compressor in all operating rotation speeds and calculating results of the performance characteristic curves in good agreement with test data. Comparisons are made against different cases of seal leakage mass flow rate for analyzing the impact of increasing seal leakage on the aerodynamic performance of the multistage axial compressor. The results indicate that the performance of the nine-stage axial compressor is degenerated faster and faster with seal leakage increasing in all operating working points, and the degeneration of performance of this compressor can be evaluated by the relationships of main performance parameters with the mass flow rate of seal leakage. Comparisons of flow distribution in the compressor for different cases of seal leakage also show that stators located in front stages of the multistage axial compressor are affected more seriously by the increasing seal leakage, and it can be confirmed that relatively larger flow losses in front stages bring significant impact on the decay of aerodynamic performance of a multistage axial compressor.

Performance Analysis of the Test Results on a Two-Stage Transonic Fan

1983 ◽  
Vol 105 (1) ◽  
pp. 125-129
Author(s):  
Baoshi Chen ◽  
Tianyi Zhang
Keyword(s):  
Mach Number ◽  
Performance Analysis ◽  
Radial Distribution ◽  
Flow Separation ◽  
High Load ◽  
Test Results ◽  
Two Stage ◽  
Stall Margin ◽  
High Incidence ◽  
Transonic Fan

Test results obtained from a two-stage fan are analysed and the reasons that caused the design performance target not to be attained are presented in this paper. Addition of a partspan shroud on rotor 1 caused higher losses and changed radial distribution of parameters. Modification on the flowpath and chord length of stator 1 resulted in excessively high inlet Mach number and flow separation in the hub region. The high load and high incidence at the hub of rotor 2 caused higher losses and reduced stall margin of the fan.

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