Any prime mover exhibits the effects of wear and tear over time. The problem of predicting the effects of wear and tear on the performance of any engine is still a matter of discussion. Because the function of a gas turbine is the result of the fine-tuned cooperation of many different components, the emphasis of this paper is on the gas turbine and its driven equipment (compressor or pump) as a system, rather than on isolated components. We will discuss the effect of degradation on the package as part of a complex system (e.g., a pipeline, a reinjection station, etc.). Treating the gas turbine package as a system reveals the effects of degradation on the match of the components as well as on the match with the driven equipment. This article will contribute insights into the problem of gas turbine systent degradation. Based on some detailed studies on the mechanisms that cause engine degradation, namely, changes in blade surfaces due to erosion or fouling, and the effect on the blade aerodynamics; changes in seal geometries and clearances, and the effect on parasitic flows; and changes in the combustion system (e.g., which result in different pattern factors), the effects of degradation will be discussed. The study includes a methodology to simulate the effects of engine and driven equipment degradation. With a relatively simple set of equations that describe the engine behavior, and a number of linear deviation factors which can easily be obtained from engine maps or test data, the equipment behavior for various degrees of degradation will be studied. A second model, using a stage by stage model for the engine compressor, is used to model the compressor deterioration. The authors have avoided to present figures about the speed of degradation, because it is subject to a variety of operational and design factors that typically cannot be controlled entirely.

Degradation in Gas Turbine Systems

In the early development of gas turbines, many empirical design rules were used; for example in obtaining fluid deflection using the deviation from blading angles, in the assumption of zero radial velocities (so-called radial equilibrium) and in expressions for clearance loss (the Lakshminarayana formulas). The validity of some of these rules, and the basic fluid mechanics behind them, is examined by use of modern ideas and computational fluid dynamics (CFD) codes. A current perspective of CFD in design is given, together with a view on future developments.

A Review of Some Early Design Practice Using Computational Fluid Dynamics and a Current Perspective

This paper describes the introduction of 3D blade designs into the core compressors of the Rolls-Royce Trent engine series with particular emphasis on the use of sweep and dihedral in the rotor designs. It follows the development of the basic ideas in a university research project, through multistage low-speed model testing, to their application to the high pressure engine compressor. An essential element of the project was the use of multistage CFD and some of the development of the method to allow the designs to take place is also discussed. Part I of the paper concentrated on the fundamental university-based research and the methods development. Part II describes additional low-speed multistage design and testing and the high-speed engine compressor designs and tests.Copyright © 2002 by ASME

The Use of Sweep and Dihedral in Multistage Axial Flow Compressor Blading: Part II — Low and High Speed Designs and Test Verification

The tandem airfoil has potential to do more work as a compressor blade than a single airfoil without incurring higher losses. The goal of this work is to evaluate the fluid mechanics of a tandem rotor in the rear stages of a core compressor. As such, the results are constrained to shock-free, fully turbulent flow with thick endwall boundary layers at the inlet. A high hub-to-tip ratio 3D blade geometry was developed based upon the best-case tandem airfoil configuration from a previous 2D study. The 3D tandem rotor was simulated in isolation in order to scrutinize the fluid mechanisms of the rotor, which had not previously been well documented. A geometrically similar single blade rotor was also simulated under the same conditions for a baseline comparison. The tandem rotor was found to outperform its single blade counterpart by attaining a higher work coefficient, polytropic efficiency and numerical stall margin. An examination of the tandem rotor fluid mechanics revealed that the forward blade acts in a similar manner to a conventional rotor. The aft blade is strongly dependent upon the flow it receives from the forward blade, and tends to be more three-dimensional and non-uniform than the forward blade.© 2008 ASME

3D Numerical Investigation of Tandem Airfoils for a Core Compressor Rotor

The purpose of this paper is to set out some of the basic principles and rules associated with the design of axial flow compressors, principally for aero-engines, as well as the practical c...

Axial flow compressor design

Effects of tip clearance, secondary flow, skew, and corner stall on the performance of a multistage compressor with controlled diffusion blading have been studied experimentally Measurements between 1 and 99 percent annulus height were carried out in both the first and the third stages of a four-stage low-speed compressor with repeating-stage blading Measurements were obtained at a datum rotor tip clearance and at a more aerodynamically desirable lower clearance The consequences of the modified rotor tip clearance on both rotor and stator performance are examined in terms of loss coefficient and gas exit angle Stator losses close to the casing are found to increase significantly when the clearance of an upstream rotor is increased These increased stator losses cause 30 percent of the stage efficiency reduction that arises with increased rotor tip clearance

Endwall Effects at Two Tip Clearances in a Multistage Axial Flow Compressor With Controlled Diffusion Blading

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Closure to “Discussion of ‘Endwall Effects at Two Tip Clearances in a Multistage Axial Flow Compressor With Controlled Diffusion Blading’” (1994, ASME J. Turbomach., 116, pp. 645–646)

Effects of tip clearance, secondary flow, skew and corner stall on the performance of a multi-stage compressor with controlled diffusion blading have been studied experimentally. Measurements between 1% and 99% annulus height were carried out in both the first and the third stages of a four-stage low-speed compressor with repeating-stage blading.Measurements were obtained at a datum rotor tip clearance and at a more aerodynamically desirable lower clearance. The consequence of the modified rotor tip clearance on both rotor and stator performance are examined in terms of loss coefficient and gas exit angle. Stator losses close to the casing are found to increase significantly when the clearance of an upstream rotor is increased. These increased stator losses cause 30% of the stage efficiency reduction which arises with increased rotor tip clearance.The deviation angles due to tip clearance from the multi-stage measurements are found to be similar to data from single stage machines with conventional blading which suggests that the unsteady flow phenomena associated with the multi-stage environment do not dominate the physics of the flow.Copyright © 1993 by ASME

J. P. Barton

Papers

Endwall Effects at Two Tip Clearances in a Multistage Axial Flow Compressor With Controlled Diffusion Blading

Closure to “Discussion of ‘Endwall Effects at Two Tip Clearances in a Multistage Axial Flow Compressor With Controlled Diffusion Blading’” (1994, ASME J. Turbomach., 116, pp. 645–646)

Endwall Effects at Two Tip Clearances in a Multistage Axial Flow Compressor With Controlled Diffusion Blading