scholarly journals Integrated Turbine Tip Clearance and Gas Turbine Engine Simulation

2016 ◽  
Author(s):  
Jeffryes W. Chapman ◽  
Ten-Huei Guo ◽  
Jonathan L. Kratz ◽  
Jonathan S. Litt
Keyword(s):  
Gas Turbine ◽  
Gas Turbine Engine ◽  
Tip Clearance ◽  
Turbine Engine ◽  
Engine Simulation

scholarly journals A Simplified Gas Turbine Engine Model With Heat Storage/Tip Clearance Effects

1993 ◽  
Author(s):  
T. H. Wong
Keyword(s):  
Gas Turbine ◽  
Heat Storage ◽  
Gas Turbine Engine ◽  
Tip Clearance ◽  
Metal Temperature ◽  
Transient Method ◽  
Transient Responses ◽  
Turbine Engine ◽  
Engine Simulation ◽  
Engine Model

A simplified turboshaft gas turbine engine model called the direct transient method (DTM) model has been developed. The DTM model consists of table look-up data generated from the actual engine data or the transient engine simulation. The DTM model accounts for heat storage, tip clearance and volume dynamics effects. It can, therefore, better predict engine transient responses and turbine metal temperature than the traditional engine horsepower extraction (HPX) model. This paper presents in detail the DTM methodology for generating accurate simplified engine models of transient performance. Comparisons of engine transient responses between the DTM and HPX models are provided.

Extensible object model for gas turbine engine simulation

2001 ◽  
Vol 21 (1) ◽  
pp. 111-118 ◽  
Author(s):  
Zhiwu Xie ◽  
Ming Su ◽  
Shilie Weng
Keyword(s):  
Gas Turbine ◽  
Gas Turbine Engine ◽  
Object Model ◽  
Turbine Engine ◽  
Engine Simulation

The Effect of Heat Transfer Coefficient Increase on Tip Clearance Control in H.P. Compressors in Gas Turbine Engine

2013 ◽  
Author(s):  
Godwin Ita Ekong ◽  
Christopher A. Long ◽  
Peter R. N. Childs
Keyword(s):  
Heat Transfer ◽  
Heat Transfer Coefficient ◽  
Transfer Coefficient ◽  
Gas Turbine ◽  
Time Constant ◽  
Gas Turbine Engine ◽  
Operating Conditions ◽  
Tip Clearance ◽  
Test Facility ◽  
Turbine Engine

Compressor tip clearance for a gas turbine engine application is the radial gap between the stationary compressor casing and the rotating blades. The gap varies significantly during different operating conditions of the engine due to centrifugal forces on the rotor and differential thermal expansions in the discs and casing. The tip clearance in the axial flow compressor of modern commercial civil aero-engines is of significance in terms of both mechanical integrity and performance. In general, the clearance is of critical importance to civil airline operators and their customers alike because as the clearance between the compressor blade tips and the casing increases, the aerodynamic efficiency will decrease and therefore the specific fuel consumption and operating costs will increase. This paper reports on the development of a range of concepts and their evaluation for the reduction and control of tip clearance in H.P. compressors using an enhanced heat transfer coefficient approach. This would lead to improvement in cruise tip clearances. A test facility has been developed for the study at the University of Sussex, incorporating a rotor and an inner shaft scaled down from a Rolls-Royce Trent aero-engine to a ratio of 0.7:1 with a rotational speed of up to 10000 rpm. The idle and maximum take-off conditions in the square cycle correspond to in-cavity rotational Reynolds numbers of 3.1×106 ≤ Reφ ≤ 1.0×107. The project involved modelling of the experimental facilities, to demonstrate proof of concept. The analysis shows that increasing the thermal response of the high pressure compressor (HPC) drum of a gas turbine engine assembly will reduce the drum time constant, thereby reducing the re-slam characteristics of the drum causing a reduction in the cold build clearance (CBC), and hence the reduction in cruise clearance. A further reduction can be achieved by introducing radial inflow into the drum cavity to further increase the disc heat transfer coefficient in the cavity; hence a further reduction in disc drum time constant.

An Integrated Approach for Gas Turbine Engine Simulation

2000 ◽  
Author(s):  
Zhiwu Xie ◽  
Ming Su ◽  
Shilie Weng
Keyword(s):  
Gas Turbine ◽  
Local Area Network ◽  
Integrated Approach ◽  
Local Area ◽  
Gas Turbine Engine ◽  
Future Research ◽  
Area Network ◽  
Turbine Engine ◽  
Engine Simulation ◽  
Engine Components

Abstract The static and transient performance of a gas turbine engine is determined by both the characteristics of the engine components and their interactions. This paper presents a generalized simulation framework that enables the integration of different component and system simulation codes. The concept of engine simulation integration and its implementation model is described. The model is designed as an object-oriented system, in which various simulation tasks are assigned to individual software components that interact with each other. A new design rationale called “message-based modeling” and its resulting class structure is presented and analyzed. The object model is implemented within a heterogeneous network environment. To demonstrate its flexibility, the codes that deal with different engine components are separately programmed on different computers running various operating systems. These components communicate with each other via a CORBA compliant ORB, which simulates the overall performance of an engine system. The resulting system has been tested on a Local Area Network (LAN) to simulate the transient response of a three-shaft gas turbine engine, subject to small fuel step perturbations. The simulation results for various network configurations are presented. It is evident that in contrast to a standalone computer simulation, the distributed implementation requires much longer simulation time. This difference of simulation efficiency is analyzed and explained. The limitations of this endeavor, along with some future research topics, are also reported in this paper.

A model for combustor dynamics for inclusion in a dynamic gas turbine engine simulation code

1997 ◽  
Author(s):  
David Costura ◽  
Tomas Velez ◽  
Patrick Lawless ◽  
Steven Frankel ◽  
David Costura ◽  
...  
Keyword(s):  
Gas Turbine ◽  
Gas Turbine Engine ◽  
Simulation Code ◽  
Turbine Engine ◽  
Engine Simulation

scholarly journals Improving Dynamic Response of a Single-Spool Gas Turbine Engine Using a Nonlinear Contoller

1992 ◽  
Author(s):  
O. F. Qi ◽  
N. R. L. Maccallum ◽  
P. J. Gawthrop
Keyword(s):  
Dynamic Response ◽  
Gas Turbine ◽  
Linear Models ◽  
Digital Simulation ◽  
Gas Turbine Engine ◽  
Open Loop ◽  
Nonlinear Controller ◽  
Nonlinear Simulation ◽  
Turbine Engine ◽  
Engine Simulation

This paper describes the design of a closed-loop nonlinear controller to improve the dynamic response of a single-spool gas turbine engine. The nonlinear controller is obtained by scheduling the gains of multivariable compensators as a function of engine non-dimensional shaft speed. The compensators, whose outputs are fuel flow and nozzle area, are designed using optimal control theory based on a set of linear models generated from a nonlinear engine simulation. Investigations are also made into developing simple algorithms to obtain an analytical expression for the compressor given its characteristic. The detailed process of developing a nonlinear simulation model for the engine is also described. The open-loop fuel controller is studied using the digital simulation.

Engine Ground Start Characteristics of Fighter Aircraft at High Altitude Airbase

2019 ◽  
Author(s):  
Santhosh Kasram ◽  
Sajath Kumar Manoharan ◽  
Mahesh P. Padwale ◽  
G. P. Ravishankar
Keyword(s):  
High Altitude ◽  
Gas Turbine ◽  
Weather Conditions ◽  
Gas Turbine Engine ◽  
Rotating Stall ◽  
Tip Clearance ◽  
Alternative Methods ◽  
Cold Weather ◽  
Oil Viscosity ◽  
Turbine Engine

Abstract The challenges faced during starting of an aircraft gas turbine engine using a Jet Fuel Starter (JFS) at high altitude airbase are discussed in this paper. Autonomous ground starts at high altitude airbase in soaked sub-zero temperature condition without any external ground support assistance is a challenge. Generally, the start cycle (sub-idle speed) at sub-zero temperatures of a gas turbine engine at high altitudes is influenced by several factors. Drag loads are estimated due to change in lube oil viscosity of engine gearbox and accessory gear box that affects available torque margin of a starter. These estimated loads are superimposed on starter characteristics to identify the available margins for successful starts. The cold start is particularly severe, since it increases the tip clearance between rotor and casing of the engine due to difference in its thermal growth. Higher tip clearances significantly degrade compressor surge margin and results in rotating stall. Inconsistent engine starts were resolved by adopting alternative methods without any change in hardware. This paper presents set of methods used to overcome inconsistent engine starts at high altitude cold weather conditions.

Implanted Component Faults and Their Effects on Gas Turbine Engine Performance

1992 ◽  
Vol 114 (2) ◽  
pp. 174-179 ◽  
Author(s):  
J. D. MacLeod ◽  
V. Taylor ◽  
J. C. G. Laflamme
Keyword(s):  
Gas Turbine ◽  
National Research Council ◽  
Engine Performance ◽  
Turbine Rotor ◽  
Gas Turbine Engine ◽  
Tip Clearance ◽  
Turbine Engine ◽  
Turbine Rotor Blade ◽  
Department Of National Defence ◽  
Project Objectives

Under the sponsorship of the Canadian Department of National Defence, the Engine Laboratory of the National Research Council of Canada (NRCC) has established a program for the evaluation of component deterioration on gas turbine engine performance. The effect is aimed at investigating the effects of typical in-service faults on the performance characteristics of each individual engine component. The objective of the program is the development of a generalized fault library, which will be used with fault identification techniques in the field, to reduce unscheduled maintenance. To evaluate the effects of implanted faults on the performance of a single spool engine, such as an Allison T56 turboprop engine, a series of faulted parts were installed. For this paper the following faults were analyzed: (a) first-stage turbine nozzle erosion damage; (b) first-stage turbine rotor blade untwist; (c) compressor seal wear; (d) first and second-stage compressor blade tip clearance increase. This paper describes the project objectives, the experimental installation, and the results of the fault implantation on engine performance. Discussed are performance variations on both engine and component characteristics. As the performance changes were significant, a rigorous measurement uncertainty analysis is included.

Sign in / Sign up

Export Citation Format

Share Document