GB2314383A

GB2314383A - Regulating a turbojet engine

Info

Publication number: GB2314383A
Application number: GB9712842A
Authority: GB
Inventors: Joachim Dr Kurzke
Original assignee: MTU Motoren und Turbinen Union Muenchen GmbH
Current assignee: MTU Aero Engines GmbH
Priority date: 1996-06-18
Filing date: 1997-06-18
Publication date: 1997-12-24
Anticipated expiration: 2017-06-18
Also published as: DE19624171C1; GB9712842D0; FR2749884B1; FR2749884A1; GB2314383B

Description

Process for regulating a turboiet engine The invention relates to a

process for regulating a turbo jet engine having a high-pressure compressor, a combustion chamber and a high-pressure turbine, in which the rotor blades of the high-pressure turbine are cooled by means of cooling air supplied by the high pressure compressor, and the rotor blade temperature of the high-pressure turbine is monitored in order to prevent overheating.

In a turbo jet engine the output of the gas turbine is directly dependent on the combustion chamber outlet temperature, subject to the limitation that this temperature must not exceed a maximum value as otherwise there is the danger of overheating of turbine parts, especially of the inlet guide wheel. However, the combustion chamber outlet temperature cannot be directly measured; instead, therefore, in practice a temperature in the region of the turbine is measured.

For this purpose one can use thermal elements or pyrometric measuring devices arranged upstream or downstream of the low-pressure turbine and measuring the rotor blade temperature of the high-pressure turbine or the low-pressure turbine directly.

In pyrometric measurement, essentially the mean metal temperature of the rotor blades is measured at a representative radius of the turbine. In an engine in which the turbine blades of the high-pressure turbine are cooled by means of cooling air taken from the outlet of the high-pressure compressor, the temperature is dependent not only on the gas temperature of the main stream, but also on the cooling air temperature.

Since the cooling air of the high-pressure turbine is taken at the outlet of the high-pressure compressor, its temperature is directly dependent on the operating point of the compressor.

At a constant combustion chamber outlet temperature, as the air and power consumption by the high-pressure shaft varies, the operating point of the high-pressure compressor, and hence also the cooling air temperature, changes, whereas the relative temperature for the rotor of the high-pressure turbine remains constant. Because the cooling air temperature changes with the air and power consumption, the metal temperature also changes. Regulation to a predetermined metal temperature, as is normally used nowadays with military thrust engines, therefore results in the combustion chamber outlet temperature changing with the air and power consumption. There therefore exists a risk of overheating of static turbine parts, especially of the inlet guide wheel, at high air and power consumption.

The aim of the invention is to provide a process for regulating a turbo jet engine that makes it possible to regulate the combustion chamber outlet temperature to a constant value, irrespective of air and power consumptions, without measuring the amounts of air and power consumption.

According to the invention, a representative relative temperature at the rotor inlet is determined (calculated) from measurement values for the compressor outlet temperature and the rotor blade temperature and from a value n,, for the cooling efficiency of the rotor blades. The relative temperature is used to regulate the engine in such a way that it does not become overheated. A significant advantage of the process of the invention is that the static parts of the high pressure turbine are reliably protected against excessive temperatures without air and power consumptions having to be measured.

According to a further development of the invention, a combustion chamber outlet temperature or an absolute rotor inlet temperature is derived from the relative temperature at the rotor inlet by means of an empirical correlation, and the engine is regulated in such a way that the combustion chamber outlet temperature or the rotor inlet temperature does not exceed a predetermined value. The advantage of this is that the combustion chamber outlet temperature or the rotor inlet temperature is given an absolute temperature value which must not exceed a predetermined value or a maximum value, so that damage to the turbine due to overheating is reliably prevented.

In some embodiments of the invention, provision is made for the rotor blade temperature to be measured by means of a radiation pyrometer, and the pyrometrically measured value of the rotor blade temperature is used for determining the representative relative temperature.

According to an embodiment of the invention the value for the effectiveness of the cooling is assumed to be constant. Alternatively, however, the value of the cooling efficiency is determined as a function of the operational parameters of the engine. This has the advantage that the value of the representative relative temperature at the rotor inlet, and possibly the values of the combustion chamber outlet temperature or the rotor inlet temperature derived from it, can be determined with even greater accuracy, and hence the engine power can be increased even more without the danger of damage due to overheating.

In further developments of the invention the combustion chamber outlet temperature or the absolute rotor inlet temperature are derived from the relative temperature at the rotor inlet by means of a functional relationship which contains the rotational speed of the engine as a variable.

The absolute rotor inlet temperature can be -4 derived from the relative temperature at the rotor inlet by means of the functional relationship T41/T..., = f (NH/T,111) where NH/T,134 is the corrected rotational speed N. of the highpressure turbine.

The high-pressure compressor outlet temperature may be measured directly or be derived from the operational parameters of the engine.

According to some embodiments of the invention, provision is made for the high-pressure outlet temperature to be used directly to determine the relative temperature at the rotor inlet.

is Finally, according to another embodiment of the invention, a value for the cooling air temperature in the region of the rotor blades of the high-pressure turbine is derived from the high-pressure compressor outlet temperature and from other operational parameters of the engine and used for determining the relative temperature at the rotor inlet.

For a better understanding of the invention embodiments of it will now be described, by way of example, with reference to the accompanying drawing, in which the single Figure shows a highly schematic representation of a jet engine.

As the Figure shows, a jet engine which is to be regulated in accordance with the process of the invention contains a combustion chamber 2, a high pressure compressor 1 connected upstream of the combustion chamber 2 in the flow direction, and a high pressure turbine 4 connected downstream of the combustion chamber 2 in the flow direction, which sets into rotation the gas stream leaving the combustion chamber 2 and drives the high-pressure compressor 1 by means of a shaft 6. A part of the air stream leaving 5_ the high-pressure compressor 1 bypasses the combustion chamber 2 in order to cool the rotor blades of the high-pressure turbine 4. A low-pressure compressor can be connected upstream of the high-pressure compressor 1, this not being shown in the Figure, and a low pressure turbine can be connected downstream of the high-pressure turbine 4, this also not being shown in the Figure.

The air drawn in by the jet engine at a temperature T1 and a pressure P, is compressed by the high-pressure compressor 1 and emitted at a temperature T3 and a pressure P3 The main part of the air emitted from the high pressure compressor 1 enters the combustion chamber 2 and is combusted there together is with the fuel supplied. As it leaves the combustion chamber 2, the gas stream has a combustion chamber outlet temperature T, and a pressure P4 - In an inlet guide wheel 3 of the high-pressure turbine 4 part of the cooling air is mixed with the main stream. The main stream temperature T, drops to T4, and the pressure P4 to P41' The gas stream entering the high-pressure turbine 4 thus has the temperature T41 and the pressure P41 The part of the compressed air delivered by the high-pressure compressor 1 that is guided past the combustion chamber 2 to cool the rotor blades of the high-pressure turbine 4 also leaves the high-pressure compressor 1 at the temperature T3 and reaches the rotor blades as cooling air at a temperature T,. The temperature of the gas stream can be described by a representative relative temperature T,,l at the rotor inlet. Between this representative relative temperature Tr,1, the cooling air temperature TK and the temperature Tmetall measured on the metal of the rotor blades there is a relationship which is given by the cooling efficiency of the rotor blades:

Trel - Tmetall ------------- T,., - TK The temperature of the rotor blades Tmetall can be measured by means of a pyrometric measuring device 5 mounted on the housing so as to receive radiation from the blades.

Since the cooling air temperature TK is closely linked to the compressor outlet temperature T3, the above formula can be solved with respect to the representative relative temperature Tre, at the rotor inlet:

is T.etall - 77K T3 Trel - - - - - - - - - - - - - - 1 - 77, (2) The representative relative temperature Tre, at the rotor inlet can thus be determined from the measured values for the compressor outlet temperature T3 and the rotor blade temperature Tmetall and from the value 77,, for the cooling effectiveness of the rotor blades. This value for the relative temperature Trel is the parameter which is used to regulate the engine in such a way that overheating of the high-pressure turbine 4, and in particular of the static inlet guide wheel 3, is prevented.

By means of an empirical correlation, the combustion chamber outlet temperature T,, or the rotor inlet temperature T,1 closely linked to it, can be derived from the relative temperature Tre, at the rotor inlet. In order to prevent overheating of the high pressure turbine parts 3 and 4, the engine is regulated in such a way that the combustion chamber outlet temperature T, or the rotor inlet temperature T,1 does not exceed a predetermined or maximum value.

The cooling effectiveness % is approximately a constant. However, on closer consideration it can be seen that 77, is also, albeit weakly, dependent on the Reynolds number. Also the radial temperature profile of the gas stream emerging from the combustion chamber 2 can influence the cooling effectiveness % locally. Furthermore, there is also a dependence on the outlet pressure P3 and outlet temperature T3 of the high-pressure compressor 1.

There is a close connection between the relative rotor inlet temperature T,,l and the absolute rotor inlet temperature T,,. As can be shown from applying the continuity equation [constant air system] to the is turbine, the relationship of the absolute rotor inlet temperature T,1 and the representative relative temperature T,el is dependent only on the corrected 112 rotational speed NI4/T,l, where NH is the rotational speed of the high-pressure turbine:

T,1/Trel = f(NH/T,11/2) (3) This functional relationship can be recorded, for instance, in the form of a table and used to regulate the engine. The rotor inlet temperature T,1 can then be determined as a function of the turbine speed NH and the representative relative temperature T,,j. The latter can in turn be calculated from the temperature T,,Iall measured by the pyrometric measuring device, of the rotor blades, the compressor outlet temperature T3 and the cooling effectiveness % Of the rotor blades, as described above.

The combustion chamber outlet temperature T, can easily be determined from the rotor inlet temperature T, by multiplication by a constant factor.

According to this process, a limiting value for the combustion chamber outlet temperature T, or for the rotor inlet temperature T,1 is used to regulate the engine instead of a limiting value for the temperature T.et.11 Of the turbine blades, measured with the pyrometric measuring device. The parts of the high pressure turbine, in particular its static parts 3, can therefore be more reliably protected from the excessive temperatures that occur as a result of high air and power consumption with a conventional method of regulation for constant rotor blade temperature.

Moreover, the measurement of air and power consumption is also not necessary.

Claims

1. A process for regulating a turbo jet engine having a high-pressure compressor, a combustion chamber and a high-pressure turbine, in which the rotor blades of the high-pressure turbine are cooled by means of the cooling air supplied by the high-pressure compressor and the rotor blade temperature of the high-pressure turbine is measured so as to prevent overheating, characterised in that a representative relative temperature Tr,l at the rotor inlet is determined from measured values for the compressor outlet temperature and the rotor blade temperature and f rom a value nk for the cooling effectiveness of the rotor blades, and in is that the relative temperature Tr,l is used for regulating the engine for preventing overheating.

2. A process according to claim 1, in which a combustion chamber outlet temperature T, is derived from the relative temperature T,,l at the rotor inlet by means of an empirical correlation, and in that the engine is regulated in such a way that the combustion chamber outlet temperature T, does not exceed a predetermined value.

3. A process according to claim 1, in which an absolute rotor inlet temperature T,, is derived from the relative temperature T,,l at the rotor inlet by means of an empirical correlation, and in that the engine is regulated in such a way that the rotor inlet temperature T,l does not exceed a predetermined value.

4. A process according to any preceding claim, in which the rotor blade temperature is measured by means of a radiation pyrometer, and the pyrometrically measured value of the rotor blade temperature T,etall is used for determining the representative relative temperature Trel -

5. A process according to any preceding claim, -lo- in which the value for the cooling effectiveness nk 'S assumed to be constant.

6. A process according to any preceding claim, in which the value for the cooling effectiveness nk 'S determined as a function of the operational parameters of the engine.

7. A process according to claim 2 or 3, in which the combustion chamber outlet temperature T, or the absolute rotor inlet temperature T41 is derived from the relative temperature T,,l at the rotor inlet by means of a functional relationship which contains the rotational speed of the engine as a parameter.

8. A process according to claim 7, in which the absolute rotor inlet temperature T,, is derived from the relative temperature T,e, at the rotor inlet by means of the functional relationship T41/Trel f (NH/T41M) where N./T41" 'S the corrected rotational speed NH of the high-pressure turbine.

9. A process according to any preceding claim, in which the high-pressure compressor outlet temperature T3 is measured directly.

10. A process according to any of claims 1 to 8, in which the high-pressure compressor outlet temperature T3 is derived from operational parameters of the engine.

11. A process according to any preceding claim, in which the high-pressure compressor outlet temperature T3 is used directly to determine the operational temperature T,,l at the rotor inlet.

12. A process according to any one of claims 1 to 10, in which a value derived from the high-pressure compressor outlet temperature T3 and from other operational parameters of the engine is used for the cooling air temperature in the region of the rotor blades of the high-pressure turbine in order to determine the relative temperature T,,l at the rotor inlet.

13. A process according to any preceding claim, in which the high-pressure compressor, combustion chamber and high-pressure turbine are a part of a multi-shaft gas turbine.

14. A process substantially as described herein with reference to the accompanying drawing.