Quick Start of an Industrial Gas Turbine Engine Through the Development of “Silent Start” VGV Schedule

2020 ◽  
Author(s):  
Senthil Krishnababu ◽  
Vili Panov ◽  
Simon Jackson ◽  
Andrew Dawson
Keyword(s):  
Guide Vane ◽  
Pressure Ratio ◽  
Axial Compressor ◽  
Rotating Stall ◽  
Rotor Blades ◽  
Start Time ◽  
Turbine Engine ◽  
High Pressure Ratio ◽  
Compressor Map ◽  
Industrial Gas Turbine

Abstract In this paper, research that was carried out to optimise an initial variable guide vane schedule of a high-pressure ratio, multistage axial compressor is reported. The research was carried out on an extensively instrumented scaled compressor rig. The compressor rig tests carried out employing the initial schedule identified regions in the low speed area of the compressor map that developed rotating stall. Rotating stall regions that caused undesirable non-synchronous vibration of rotor blades were identified. The variable guide vane schedule optimisation carried out balancing the aerodynamic, aero-mechanical and blade dynamic characteristics gave the ‘Silent Start’ variable guide vane schedule, that prevented the development of rotating stall in the start regime and removed the non-synchronous vibration. Aerodynamic performance and aero-mechanical characteristics of the compressor when operated with the initial schedule and the optimised ‘Silent Start’ schedule are compared. The compressor with the ‘Silent Start’ variable guide vane schedule when used on a twin shaft engine reduced the start time to minimum load by a factor of four and significantly improved the operability of the engine compared to when the initial schedule was used.

Methods of Improving the Axial Compressor Flow Passage to Reduce the Flow Circumferential Nonuniformity

2015 ◽  
Author(s):  
Daria Kolmakova ◽  
Grigorii Popov
Keyword(s):  
Gas Flow ◽  
Guide Vane ◽  
Axial Compressor ◽  
Rotor Blades ◽  
Flow Passage ◽  
Turbine Engine ◽  
Engine Design ◽  
Pressure Peak ◽  
Engine Systems ◽  
Circumferential Nonuniformity

Circumferential nonuniformity of gas flow is one of the main problems that can occur in the gas turbine engine. Usually, the flow circumferential nonuniformity appears near the support, located in the flow passage of the engine. The presence of circumferential nonuniformity leads to the increased dynamic stresses in the blade rows and the blade damage. The goal of this research was to find the ways of the flow non-uniformity reduction, which would not require a fundamental changing of engine design. A new method for reducing the circumferential nonuniformity of gas flow was proposed. It has been suggested to increase the gap of the leading edges of support racks from the trailing edge of the upstream guide vane blades which will result in achieving the desired results. An important advantage of this method is that the internal cavities of racks remain unchanged for the placement of engine systems. Moreover, the proposed method allows the prediction of the pressure peak values after the rotor blades without .

On the Prediction of Rotating Stall in an Industrial Gas Turbine Compressor

2020 ◽  
Author(s):  
Senthil K. Krishnababu
Keyword(s):  
Gas Turbine ◽  
Large Scale ◽  
Guide Vane ◽  
Rotating Stall ◽  
Computational Domain ◽  
Turbine Engine ◽  
Design Cycle ◽  
Lower Test ◽  
Gas Turbine Compressor ◽  
Industrial Gas Turbine

Abstract An investigation is presented into the computation of rotating stall in an industrial gas turbine compressor using a hybrid whole annulus and single passage computational domain. The objective of this investigation is to demonstrate the use of large-scale unsteady computations with quicker turn-around times in the design cycle to develop and evaluate several variable guide vane schedules and/or bleed settings. This means that subsequent engine test campaign could be carried out with significantly lower test matrix size in terms of the number of variable guide vane schedules and/or the handling bleed settings thus reducing the overall development time and cost. Rotating stall that was measured and characterised during a previous compressor rig test (Krishnababu, et al. [1]) were successfully predicted by large-scale unsteady computations using TurboStream. The predicted number of stall cells and their speed agreed closely with the test data. The methodology validated was applied to predict and mitigate the rotating stall in the development of a compressor for a new gas turbine engine. Using this approach, it was possible to define bleed control system that ensured stall free operation.

Stall Behavior in an Ultra-High-Pressure-Ratio Centrifugal Compressor: Backward-Traveling Rotating Stall

2021 ◽  
pp. 1-51
Author(s):  
Yingjie Zhang ◽  
Xingen Lu ◽  
Yanfeng Zhang ◽  
Ziqing Zhang ◽  
Xu Dong ◽  
...  
Keyword(s):  
Boundary Layer ◽  
Shock Wave ◽  
Shock Intensity ◽  
Incidence Angle ◽  
Pressure Ratio ◽  
Leading Edge ◽  
Boundary Layer Separation ◽  
Rotating Stall ◽  
High Pressure Ratio ◽  
Ultra High Pressure

Abstract This paper describes the stall mechanism in an ultra-high-pressure-ratio centrifugal compressor. A model comprising all impeller and diffuser blade passages is used to conduct unsteady simulations that trace the onset of instability in the compressor. Backward-traveling rotating stall waves appear at the inlet of the radial diffuser when the compressor is throttled. Six stall cells propagate circumferentially at approximately 0.7% of the impeller rotation speed. The detached shock of the radial diffuser leading edge and the number of stall cells determine the direction of stall propagation, which is opposite to the impeller rotation direction. Dynamic mode decomposition is applied to instantaneous flow fields to extract the flow structure related to the stall mode. This shows that intensive pressure fluctuations concentrate in the diffuser throat as a result of changes in the detached shock intensity. The diffuser passage stall and stall recovery are accompanied by changes in incidence angle and shock wave intensity. When the diffuser passage stalls, the shock-induced boundary-layer separation region near the diffuser vane suction surface gradually expands, increasing the incidence angle and decreasing the shock intensity. The shock is pushed from the diffuser throat toward the diffuser leading edge. When the diffuser passage recovers from stall, the shock wave gradually returns to the diffuser throat, with the incidence angle decreasing and the shock intensity increasing. Once the shock intensity reaches its maximum, the diffuser passage suffers severe shock-induced boundary-layer separation and stalls again.

Characteristics of Stable Rotating Stall Cells in an Axial Compressor

2017 ◽  
Author(s):  
Adam R. Hickman ◽  
Scott C. Morris
Keyword(s):  
Flow Field ◽  
High Speed ◽  
Pressure Ratio ◽  
Pressure Sensors ◽  
Axial Compressor ◽  
Field Measurements ◽  
Rotating Stall ◽  
Flow Coefficient ◽  
Stable Operation ◽  
Double Phase

Flow field measurements of a high-speed axial compressor are presented during pre-stall and post-stall conditions. The paper provides an analysis of measurements from a circumferential array of unsteady shroud static pressure sensors during stall cell development. At low-speed, the stall cell approached a stable size in approximately two rotor revolutions. At higher speeds, the stall cell developed within a short amount of time after stall inception, but then fluctuated in circumferential extent as the compressor transiently approached a stable post-stall operating point. The size of the stall cell was found to be related to the annulus average flow coefficient. A discussion of Phase-Locked Average (PLA) statistics on flow field measurements during stable operation is also included. In conditions where rotating stall is present, flow field measurements can be Double Phase-Locked Averaged (DPLA) using a once-per-revolution (1/Rev) pulse and the period of the stall cell. The DPLA method provides greater detail and understanding into the structure of the stall cell. DPLA data indicated that a stalled compressor annulus can be considered to contained three main regions: over-pressurized passages, stalled passages, and recovering passages. Within the over-pressured region, rotor passages exhibited increased blade loading and pressure ratio compared to pre-stall values.

scholarly journals The Design and Evaluation of a High Pressure Ratio Radial Turbine

1992 ◽  
Author(s):  
K. R. Pullen ◽  
N. C. Baines ◽  
S. H. Hill
Keyword(s):  
High Pressure ◽  
High Speed ◽  
Pressure Ratio ◽  
Performance Measurements ◽  
Turbine Engine ◽  
High Pressure Ratio ◽  
Turbine Efficiency ◽  
Wide Range ◽  
Dimensional Parameters ◽  
Very High

A single stage, high speed, high pressure ratio radial inflow turbine was designed for a single shaft gas turbine engine in the 200 kW power range. A model turbine has been tested in a cold rig facility with correct simulation of the important non-dimensional parameters. Performance measurements over a wide range of operation were made, together with extensive volute and exhaust traverses, so that gas velocities and incidence and deviation angles could be deduced. The turbine efficiency was lower than expected at all but the lowest speed. The rotor incidence and exit swirl angles, as obtained from the rig test data, were very similar to the design assumptions. However, evidence was found of a region of separation in the nozzle vane passages, presumably caused by a very high curvature in the endwall just upstream of the vane leading edges. The effects of such a separation are shown to be consistent with the observed performance.

Surge Margin Tracking for Active Control of Quick Windmill Relighting

2004 ◽  
Author(s):  
Sog-Kyun Kim ◽  
Ian A. Griffin ◽  
Haydn A. Thompson ◽  
Peter J. Fleming
Keyword(s):  
High Pressure ◽  
Active Control ◽  
Pressure Ratio ◽  
Flight Test ◽  
Operating Pressure ◽  
Turbine Engine ◽  
Engine Model ◽  
Surge Margin ◽  
Stator Vane ◽  
Compressor Map

Surge margin tracking logic is developed for use in the control of quick windmill relighting (QWR) at sub-idle. Using existing high pressure compressor (HPC) characteristics (but without any gas turbine engine model), the surge margin can be calculated and used to approximate the air flow which is currently not measured in flight. During the QWR flight test, only limited measurements excluding the airflow measurement are available. Based on the fact that a beta value is equivalent to the position of the throttle valve in a compressor test rig, the role of the beta value is here to interrelate between the PRC (pressure ratio of compressor) and NDMF (non-dimensional mass flow) values for the measured CNH (corrected high pressure spool speed) and PRC values. Using the proposed scaling factors (SFs), the HPC map in terms of PRC is adaptively scaled with the engine parameters to cover the operating pressure ratio of the HPC. These account qualitatively for the effects of heat soakage and stability aids such as bleed and VSV (variable stator vane) on the compressor map. The simulation results show that the variable SF approach is more realistic in estimation of the surge margin, compared to the fixed SF approach. As a result of this proposed surge margin tracking logic, an active control for QWR may be possible using an estimated surge margin to adjust the fuel flow. This improves the pull-away time to reach idle power without danger of stall or surge during QWR.

Tailored Structural Design and Aeromechanical Control of Axial Compressor Stall: Part II — Evaluation of Approaches

2003 ◽  
Author(s):  
L. G. Fre´chette ◽  
O. G. McGee ◽  
M. B. Graf
Keyword(s):  
Structural Parameters ◽  
Axial Compressor ◽  
Coupled System ◽  
Rotating Stall ◽  
Rotor Blades ◽  
Tip Clearance ◽  
Theoretical Evaluation ◽  
Axial Compressors ◽  
Spatial Modes ◽  
Control Authority

A theoretical evaluation was conducted delineating how aeromechanical feedback control can be utilized to stabilize the inception of rotating stall in axial compressors. Ten aeromechanical control methodologies were quantitatively examined based on the analytical formulations presented in the first part of this paper (McGee et al, 2003a). The maximum operating range for each scheme is determined for optimized structural parameters, and the various schemes are compared. The present study shows that the most promising aeromechanical designs and controls for a class of low-speed axial compressors were the use of dynamic fluid injection. Aeromechanically incorporating variable duct geometries and dynamically re-staggered IGV and rotor blades were predicted to yield less controllability. The aeromechanical interaction of a flexible casing wall was predicted to be destabilizing, and thus should be avoided by designing sufficiently rigid structures to prevent casing ovalization or other structurally-induced variations in tip clearance. Control authority, a metric developed in the first part of this paper, provided a useful interpretation of the aeromechanical damping of the coupled system. The model predictions also show that higher spatial modes can become limiting with aeromechanical feedback, both in control of rotating stall as well as in considering the effects of lighter, less rigid structural aeroengine designs on compressor stability.

Design and Test of a Novel Highly-Loaded Compressor

2019 ◽  
Author(s):  
Chengwu Yang ◽  
Ge Han ◽  
Shengfeng Zhao ◽  
Xingen Lu ◽  
Yanfeng Zhang ◽  
...  
Keyword(s):  
Pressure Ratio ◽  
Axial Compressor ◽  
Pressure Rise ◽  
Tip Clearance ◽  
Leakage Flow ◽  
Flow Angle ◽  
Stall Margin ◽  
High Pressure Ratio ◽  
Numerical Predictions ◽  
Highly Loaded

Abstract The blades of rear stages in small size core compressors are reduced to shorter than 20 mm or even less due to overall high pressure ratio. The growing of tip clearance-to-blade height ratio of the rear stages enhance the leakage flow and increase the possibility of a strong clearance sensitivity, thus limiting the compressor efficiency and stability. A new concept of compressor, namely diffuser passage compressor (DP), for small size core compressors was introduced. The design aims at making the compressors robust to tip clearance leakage flow by reducing pressure difference between pressure and suction surfaces. To validate the concept, the second stage of a two-stage highly loaded axial compressor was designed with DP rotor according to a diffuser map. The diffuser passage stage has the same inlet condition and loading as the conventional compressor (CNV) stage, of which the work coefficient is around 0.37. The predicted performance and flow field of the DP were compared with the conventional axial compressor in detail. The rig testing was supplemented with the numerical predictions. Results reveal that the throttle characteristic of DP indicates higher pressure rise and the loss reduction in tip clearance is mainly responsible for the performance improvement. For the compressor with DP, the pressure and flow angle are more uniform on exit plane. What’s more, the rotor with diffused passage reveals more robust than the conventional rotor at double clearance gap. Furthermore, the experimental data indicate that DP presents higher pressure rise at design and part speeds. At design speed, the stall margin was extended by 7.25%. Moreover, peak adiabatic efficiency of DP is also higher than that of CNV by about 0.7%.

Complex Gas Dynamic Optimization of a Three Spool Axial Compressor of an Industrial Gas Turbine Engine

2019 ◽  
Author(s):  
Grigorii M. Popov ◽  
Igor Egorov ◽  
Dmitrii Dmitriev ◽  
Evgenii S. Goriachkin ◽  
Andrei A. Volkov
Keyword(s):  
Dynamic Optimization ◽  
Axial Compressor ◽  
Operating Conditions ◽  
Specific Task ◽  
Turbine Engine ◽  
Gas Dynamic ◽  
Impressive Result ◽  
Complex Models ◽  
Overall Efficiency ◽  
Industrial Gas Turbine

Abstract The paper provides a description of the algorithm, an example of a specific task, and the results of the optimization of a 15-stage three-spool compressor for a ground-based GTU by the efficiency criteria of the engine. It can be performed by using the proposed algorithm to find such a compressor configuration that will be not just the optimum compressor, but the best option for working as part of the engine under the specified operating conditions and with various types of required restrictions. Using the proposed algorithm, varying only the stagger angles of the profiles, the authors managed to find a way to increase the overall efficiency of the NK-36ST engine by 0.43%. Obviously, it is possible to achieve a more impressive result and at the same time to increase reliability, reduce the weight and cost of the engine by applying more complex models by changing the shape of the blade profiles.

Design and Test of a Transonic Axial Splittered Rotor

2015 ◽  
Author(s):  
Garth V. Hobson ◽  
Anthony J. Gannon ◽  
Scott Drayton
Keyword(s):  
Mass Flow ◽  
Pressure Ratio ◽  
Axial Compressor ◽  
Design Procedure ◽  
Tip Clearance ◽  
Turbine Engine ◽  
Transonic Compressor ◽  
Compressor Rotor ◽  
Transonic Axial Compressor ◽  
Commercial Off The Shelf

A new design procedure was developed that uses commercial-off-the-shelf software (MATLAB, SolidWorks, and ANSYS-CFX) for the geometric rendering and analysis of a transonic axial compressor rotor with splitter blades. Predictive numerical simulations were conducted and experimental data were collected in a Transonic Compressor Rig. This study advanced the understanding of splitter blade geometry, placement, and performance benefits. In particular, it was determined that moving the splitter blade forward in the passage between the main blades, which was a departure from the trends demonstrated in the few available previous transonic axial compressor splitter blade studies, increased the mass flow range with no loss in overall performance. With a large 0.91 mm (0.036 in) tip clearance, to preserve the integrity of the rotor, the experimentally measured peak total-to-total pressure ratio was 1.69 and the peak total-to-total isentropic efficiency was 72 percent at 100 percent design speed. Additionally, a higher than predicted 7.5 percent mass flow rate range was experimentally measured, which would make for easier engine control if this concept were to be included in an actual gas turbine engine.

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