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93-GT-1 98

THE AMERICAN SOCIETY OF MECHANICAL ENGINEERS 345 E. 47th St., New York, N.Y. 10017 The Society shall not be responsible for statements or opinions advanced in i 7 papers or discussion at meetings of the Society or of its Divisions or Sections, ® or printed in its publications. Discussion is printed only if the paper is published in an ASME Journal. Papers are available from ASME for 15 months after the meeting.

Printed in U.S.A.

Copyright © 1993 by ASME

BASE-METAL THERMOCOUPLE TECHNOLOGY FOR IMPROVED TEMPERATURE MEASUREMENT IN GAS TURBINE ENGINES C. Paul Furniss Incotherm Limited Hereford, United Kingdom

Figure 1 (Courtesy of Rolls Royce PLC Derby,England,Engine Division)

ABSTRACT

During the next decade there will be growing pressures placed upon the manufacturers of gas turbines to produce more operationally efficient engines. Omn,YE550A There are two main end-use groupings for gas turbines.The parameters for efficiency may prove to be quite different for these end-use groups,requiring a separate emphasis for engineering design.

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With respect to aircraft propulsion gas turbines,the efficiencies may tend towards greater fuel economy and unit power outputs. In contrast to this the ground based gas turbine units may require increased unit power output but be restricted by the tightening emission requirements being dictated by international pollution laws. One of the key areas of focus for engineering design,in order to satisfy such performance demands,is that of improved operational control of the turbine. The process variables requiring accurate,reliable and repeatable monitoring and control include rotational speed, linear speed,pressure mass flow rate and temperature.Whilst all of these phenomena require correct control, it may be argued that temperature is of extreme importance for both an operational efficiency and safety viewpoint.

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This paper will attempt to explore the problems associated with conventional methods of gas turbine temperature measurement and discuss possible solutions using novel new technologies that will allow the earlier realisation of these efficiency goals.

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INTRODUCTION

GROUND BASED GAS TURBINES Figure Emissions Legislation

Figure 2b (After Lefebre,1903)

2a (Courtesy of Siemens AG Erlangen, Germany)

Prior to 1980, emissions control of gas turbines worldwide had not been rigidly enforced.However, the legislation passed in several countries over the past few years to restrict NO, emissions requires the introduction of a new combustion control for gas turbines employing primary (preventive) measures in order to limit the thermal NO, production.

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Figure 1 illustrates the general arrangement of components in an aircraft gas turbine engine,detai ling the temperature,pressure and velocities of the combustion gases.Ground-based units are similar,but do not possess a by-pass fan ("turbofan").

Influence of temperature on CO and NOx emissions

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There are two different sources of NO,. 1. The oxidation of organic nitrogen compounds bonded in fuel,such as ammonia in gaseous fuels or nitrogenous hydrogen compounds in liquid fuels,is unavoidable.

The flame temperature ( Figure 1,T: ) may be lowered by injecting water or steam into the combustion zone/chamber ( Figure 2a ) and so restricting combustion temperature to circa 1710 Kelvin.It follows that both NO, and CO may be minimised at this temperature ( Figure 2b.Ref 15 ).

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2. The thermal formation of NO„ ((SO and NO2) from the high-temperature reaction between the nitrogen and the oxygen in the intake air is considerably more serious in quantitative terms because the compressor mass flow is more than 50 times the fuel consumption.The thermal NO, production can,however, be effectively reduced by lowering the flame temperature. It follows that the effective control of temperature, through accuracy in its measurement,is of great importance for success in emission control. Depending on the engine manufacturer,a temperature measurement of better than +/- 12°C is required,in order to comply with the stringent emmissions levels legislation worldwide.

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NO, Pollution Gases and their Control

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Ground-Based Gas Turbine Temperature Ranges -

The range of temperatures present in a ground-based gas turbine may be from -40°C at the inlet to 1530°C within the combustion chamber.This may be represented throughout the engine by a diagram as shown in Figure 1.Crucial temperatures with respect to emissions control will be Ti and Ti .The compressor section of the gas turbine may raise the compressed air temperature to around 600°C at To.

Presented at the International Gas Turbine and Aeroengine Congress and Exposition Cincinnati, Ohio — May 24-27, 1993 Downloaded From: https://proceedings.asmedigitalcollection.asme.org on 04/26/2019 Terms of Use: http://www.asme.org/about-asme/terms-of-use

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AIRCRAFT GAS TURBINE ENGINES Performance and Efficiency In contrast to the case for ground-based units subject to stringent pollution control constraints,efficiency increase demands are pushing operational temperatures higher for flying engines. The gas turbine engine is basically a heat engine using air as a working fluid to provide thrust.To achieve this,the air passing through the engine has to be accelerated;this means that the velocity or kinetic energy of the air is increased.To obtain this increase,the pressure energy is first of all increased,followed by the addition of heat energy,before final conversion back to kinetic energy in the form of a high velocity jet efflux.The higher the temperature of combustion the greater is the expansion of the gases and the greater is the thrust.The temperature is,however,limited by the materials and design of the turbine assembly. From a safety viewpoint alone,accurate and reliable temperature measurement is essential for engine integrity. The trend in aircraft gas turbine technology is towards higher turbine engine temperatures (TETs),higher reheat temperatures and higher overall pressure ratios;there is also a trend to use higher bypass ratios.In the mid 1970s a common TET was 1320°C.It is believed that this will increase to 1520°C in the mid 1990s (Ref 1.). It is perceived that smaller aircraft engines with a greater thrust/weight ratio will result from advanced technologies over the coming years.These improvements in unit power output will also be accompanied by reductions in specific fuel consumption.The TET increases will be obtained by use of improved materials;increased cooling air(quantity & quality) and improved blade cooling techniques.

3. To thermally conduct the process temperature to the measuring thermocouple wires. The most popular material types currently in use are AISI 310 stainless steel, INCONEL*alloys 600 & 601,INCO*alloy HX and NIMONIC'alloy 75. PROBLEMS WITH CONVENTIONAL BASE-METAL THERMOCOUPLE MATERIALS Thermocouple Wire (Thermoelement) Problems There are three principal characteristic types and causes of thermoelectric instability in the standard base-metal thermoelement materials used in gas turbines (Ref 3;Table 1). 1. A gradual and cumulative drift in thermal emf on long exposure at elevated temperatures.This is observed in all conventional base-metal materials and is mainly due to oxidation and/or contamination of the thermoelements during use.The contamination is caused by the diffusion of high-vapour-pressure elements (such as aluminium (Al) and manganese) through the MgO insulant from one thermocouple wire to the other.Both type J thermocouple wires contain manganese (Mn) and the type K negative wire contains both Al & Mn (Figures 3 & 4). Type K can give a 16°C error after 700 hours at 1200°C. It should be noted that in the nickel-base ANSI types KP & KN thermoalloys the development of compositional inhomogeneities can be quite severe as reactive solute elements,in particular chromium,manganese and aluminium,are depleted by internal oxidation. Figure 3 (:after NBS Monograph 161 1978)

It should be further noted that under supersonic conditions of flight, gas turbine blades may encounter temperatures in the range 880°C to 1080°C. Manufacturers of flying engines are seeking technolgies to allow extended warranties for continuous unit operation to be increased from 10,000 hours to 30,000 hours. Aircraft Gas Turbine Temperature Ranges Due to an emphasis on thrust maximisation and overall performance, flying engines may be subject to higher maximum temperatures than those found in ground-based units.The power and maximum thrust requirements at take-off,in both military and civil aircraft,and for combat roles,in military aircraft,necessitate this,albeit for periods of relatively short duration. After combustion,the temperature of the gases released is about 1800 to 2130°C,which is too hot for entry to the nozzle guide vanes of the turbine.The air not used for combustion,60 to 75% of the total airflow,is introduced progressively into the flame tube.Approximately 50% is used to lower the gas temperature before it enters the turbine and the other half is used for cooling the walls of the flame tube (Ref 2.). Despite this cooling,the continuous flow of gas to which the turbine is exposed may have an entry temperature of between 700 and 1200°C. The temperature of the exhaust gases may be between 550 and 850°C in normal flight.This will be raised to 1700°C or higher with the use of an afterburner during take-off,climbing or combat (military aircraft).

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