US3903691A - Method and devices for avoiding the formation of thermal imbalances in turbine engines - Google Patents

Abstract

A device for avoiding the formation of thermal imbalances affecting the rotating parts of turbine engines. The device include passages and jets for injecting compressed air into the engine in proximity to the rotating parts in order to render the temperatures therein uniform, such injection being made as soon as the engine stops.

Description

[451 Sept. 9, 1975 United States Patent Szydlowski 415/DIG. 1 60/39 66 60/3966 60/3966 415/115 nut u m mmhm u t mmw m d nwm imam m 6 am AJMSS 4680 56677 9w999 H/HHH 8 02 1 11 9 230 26778 56558 88632 ,235 23333 METHOD AND DEVICES FOR AVOIDING THE FORMATION OF THERMAL IMBALANCES IN TURBINE ENGINES [76] Inventor: Joseph Szydlowski, Usine Turbomeca, Bordes (Basses Pyrenees), France Mar. 19, 1973 Appl. No.: 342,881

Primary ExaminerWil1iam L. Freeh Assistant Examiner-Robert E. Garrett Attorney, Agent, or FirmMurray Schaffer [22] Filed:

ABSTRACT [30] Foreign Application Priority Data May 26, 1972 France A device for avoiding the formation of thermal imbalances affecting the rotating parts of turbine engines.

The device include passages and jets for injecting 68 11 /7 C w 1m 1. m (mm 9m 3 l- 0 m L C St on UH 55 [58] Field of Search........................ 60/3966, 39.19;

compressed air into the engine in proximity to the rotating parts in order to render the temperatures therein uniform the engine stops.

, such injection being made as soon as References Cited UNITED STATES PATENTS 2,468,461 415/116 4 Claims, 3 Drawing Figures METHOD AND DEVICES FOR AVOIDING THE FORMATION OF THERMAL IMBALANCES llN TURBINE ENGTNES BACKGROUND OF THE INVENTION The present invention relates to internal combustion engines of the turbine type.

As soon as a gas turbine stops, a thermal imbalance occurs in its rotating components. This occurs because the hot gases are no longer swirled within the combustion chamber and the upper part of the rotating components become hotter than their lower part. This results in a thermal imbalance which, when the engine is started again, can cause stresses in the rotating components leading to more or less extensive damage.

The degree of thermal imbalance depends on the time during which the turbine is stopped. While it occurs on all internal combustion engines utilizing a turbine, the magnitude of its consequences varies according to the engine structure. It is particularly marked in the case of engines which, in the course of being speeded up, cross a critical level of rotation speeds. During the passage of the engine through this critical speed a significant thermal imbalance can have serious consequences.

In order to avoid these drawbacks, turbine manufacturers have heretofore recommended taking various precautions, including, in particular, the following:

a. The engine must not be restarted until a clearly specified time has elapsed since it was last shut off.

b. Certain ventilation operations must be carried out in order to render the temperatures of the rotating components uniform.

c. The engine must be rotated by means of an engine spinner for a certain time before start-up until the temperatures have become uniform.

These palliatives lead to a compulsory period of nonutilization of the engine each time the same is shut off and create a severe constraint operationally and economically on the operators.

The present invention has as its object the reduction of such downtimes by provision of simple and effective means and by a method of utilizing same, whereby the temperatures of the different parts of the rotating engine components are rendered uniform, as soon as the engine has stopped.

SUMMARY OF THE INVENTEON To this end, a gas turbine is provided which according to the invention includes jets and passages connected to compressed air delivery pipes originating from a source external to the turbine engine. The jets inject air at judiciously selected points inside the engine, preferably where the temperature is highest, thereby producing a swirling of the atmosphere surrounding the rotating components which are most liable to be affected by any thermal imbalance.

The invention further provides a method for utilizing such devices, in which the air injection is effected as soon as the engine stops, thus preventing the onset of thermal imbalances in the rotating engine components.

BRIEF DESCRiPTION OF DRAWINGS The description which follows with reference to the accompanying non-limitative exemplary drawings will give a clear understanding of how the invention can be carried into practice.

In the drawings:

FIG. 1 is an axial half-sectional view of a gas turbine engine with an annular combustion chamber and central centrifugal fuel delivery, having air injection devices according to the invention;

FIG. 2 shows on an enlarged scale that portion of the turbine engine in FIG. 1 in which the devices according to the invention are situated; and

FIG. 3 is a partial cross-sectional view through III- III of FIG. 2.

DESCRIPTION OF THE INVENTION In FIG. I a gas turbine is shown having an annular combustion chamber formed between an outer wall 1 and an inner wall 2 and debouching into the turbine proper, generally depicted by the numeral 18, through a fixed inlet nozzle diaphragm 3 positioned up-stream of a first stage turbine wheel 4. The inlet nozzle diaphragm comprises fixed hollow stator blades 5 which are restrained in slots formed in an outer ring 6 and an inner ring 7 and which are suitably welded to the rings.

In the exemplary embodiment considered herein, the turbine 18 has three stages, the wheel 4 comprising the first stage, while the two other stages comprise turbine wheels 8 and 9. Between the consecutive wheels 4, 8 and 9 are interposed fixed nozzle diaphragms l0 and 11 of fixed stator blade construction. The whole assembly is located within a turbine casing 12 which forms an envelope for the air flow delivered by a centrifugal compressor 13. The entire volume of air from the compressor enters the combustion chamber through slits, openings and tubes with which both outer wall 1 and inner wall 2 are provided. The air which enters the combustion chamber through inner wall 2 first passes through the hollow blades 5 of the first-stage nozzle diaphragm 3 and cools them in so doing.

Between nozzle diaphragm 3 and the first stage 4 of turbine 18 there is provided at least one air inlet 14 which feeds through a nozzle or jet 15 formed in the outer ring 6 of nozzle diaphragm 3. By means of jet l5, compressed air from an external source (not shown) is conveyed through a pipe 17 associated to air inlet 14 and is injected into the engine and causes air to swirl over the turbine and more particularly over the firststage wheel 4 thereof, which is its hottest part. This renders the temperatures of the different parts of the rotating components uniform.

It is a further teaching of this invention that the air injection is effected immediately after fuel combustion in the combustion chamber of the gas turbine ceases. This can be accomplished by providing suitable valving, automatic controls and the like. Thus, instead of countering already existing thermal imbalances, as would be the case if the temperature equalization process were undertaken subsequently only, the very formation of such imbalances is prevented by this method. This results in significant time saving and correspondingly shortened engine down-time, and also in a more efficient process of temperature equalization, since the high temperatures of the rotating components assist in the heat exchange with the injected air. Further, judiciously localized injections limit energy expenditure to a minimum for maximum efficiency, in contrast 0t prior art processes involving either ventilations or the use of engine spinners, in which considerable quantities of air are involved in superfluous heat exchanges with engine components not directly affected by the imbal ance elimination process.

At least two jets must be provided in the envelope of nozzle diaphragm 3. Preferably, such jets are arranged evenly along the periphery of ring 6.

Such passages and jets are preferably directed obliquely rather than radially (FlG. 3), whereby the compressed air streams injected by them cause the air mass adjacent wheel 4 to be rotated about the turbine axis, thus assisting and accelerating the temperature equalization process. The passages and jets 15 are preferably inclined in the normal direction of rotation of the turbine, but may be inclined in the opposite direction if desired.

The jets 15 may be disposed at other points in the turbine engine, an example being proximate the second and third stages of turbine 18, or ahead of the upstream end of the combustion chamber.

It goes without saying that changes and substitutions may be made in the form of embodiment hereinbefore described without departing from the scope of the ineach having a rotable turbine wheel connected to said compressor and driving the same, and a fixed nozzle diaphragm positioned upstream thereof, at least one closed passage extending from said casing through said engine into proximity with at least one stage. means for communicating said passage with a source of compressed air external of said engine, and a terminal jet on said passage for delivering a stream of said compressed air into said one stage about said associated rotable turbine wheel independent of the operation of said engine, said jet being positioned in proximity to the periphery of said associated rotatable wheel and having a port at an angle to the radius line from the axis of said rotatable turbine wheel inclined in the direction opposite to the normal direction of rotation of said rotatable turbine wheel.

2. The device as claimed in claim 1, in which said jet is positioned between said rotatable turbine wheel and the associated nozzle diaphragm thereof.

3. The device as claimed in claim 3, in which said jet is positioned between the rotatable turbine wheel and the nozzle diaphragm of the first stage.

4. The device as claimed in claim 1, in which there are at least two of said passages and jets associated with

Claims (4)

1. A gas turbine engine of the type having an outer casing, a compressor, a combustion chamber, a rotary fuel injector, and a turbine having a plurality of stages each having a rotable turbine wheel connected to said compressor and driving the same, and a fixed nozzle diaphragm positioned upstream thereof, at least one closed passage extending from said casing through said engine into proximity with at least one stage, means for communicating said passage with a source of compressed air external of said engine, and a terminal jet on said passage for delivering a stream of said compressed air into said one stage about said associated rotable turbine wheel independent of the operation of said engine, said jet being positioned in proximity to the periphery of said associated rotatable wheel and having a port at an angle to the radius line from the axis of said rotatable turbine wheel inclined in the direction opposite to the normal direction of rotation of said rotatable turbine wheel.

2. The device as claimed in claim 1, in which said jet is positioned between said rotatable turbine wheel and the associated nozzle diaphragm thereof.

3. The device as claimed in claim 3, in which said jet is positioned between the rotatable turbine wheel and the nozzle diaphragm of the first stage.

4. The device as claimed in claim 1, in which there are at least two of said passages and jets associated with each stage.

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