GB1581855A

GB1581855A - Turbomachine performance

Info

Publication number: GB1581855A
Application number: GB1576977A
Authority: GB
Original assignee: General Electric Co
Current assignee: General Electric Co
Priority date: 1976-08-02
Filing date: 1977-04-15
Publication date: 1980-12-31
Also published as: BE854122A; FR2360750A1; FR2360750B1; DE2718610A1; IT1076442B; JPS5317813A

Description

(54) IMPROVEMENTS IN METHODS OF IMPROVING TURBOMACHINE PERFORMANCE (71) We, GENERAL ELECTRIC COMPANY, a corporation organized and existing under the laws of the State of New York, United States of America, residing at 1, River Road, Schenectady, 12305, State of New York, United states of America, do hereby declare the invention, for which we pray that a patent may be granted to us, and the method by which it is to be performed, to be particularly described in and by the following statement: This invention relates generally to turbomachines and, more particularly, to a method for reducing the operational clearance between a turbine or compressor rotor and its surrounding shroud during predetermined conditions of operation.

In an effort to maintain a high degree of efficiency, manufacturers of turbine engines have strived to maintain the closest possible clearance between the engine rotor and the surrounding stator structure, since any gas which may pass therebetween represents a loss of energy to the system. If the system were to operate only under steady-state conditions, it would be a simple matter to establish the desired close clearance relationship between the rotor and the stator to obtain the greatest possible efficiency without allowing frictional interference between the elements. However, in reality, all turbine engines must initially be brought from a stand-still condition up to steady-state speed, and then eventually decelerate to the stand-still condition.Further, in turbine engines of the type used to propel jet aircraft, they must be capable of operation over variable transient conditions such as, for example, a cold rotor burst, a throttle chop, and a hot rotor burst. The difficulty in obtaining constant clearance between the shroud and the rotor during such transient operation, is caused by first, the variable mechanical expansion and shrinkage of the rotor as brought about by changes of speed, and secondly, by the relative thermal growth between the two structures as caused by the necessary difference in thermal inertia. One method of minimizing the turbine tip clearance of turbomachines has been to properly select the various materials which exhibit thermal properties that will assist in matching their radial responses at different engine operating conditions.Another method has been to direct variable temperature air on a shroud support structure so as to change the growth and shrinkage rate thereof in an effort to accommodate different engine operational conditions. However, it has generally been the practice to establish turbine tip clearances on the basis of transient engine conditions and maximum power settings. That is, these are the conditions under which the clearance between the turbine rotor and the shroud is at a minimum, and during all other conditions of operation the clearance will be greater than that required for safe, interference-free operation. It follows then that during these other periods of operation there is a loss of efficiency to the extent that this clearance is greater than it need be.It will be recognized that when the engine is operating in the cruise condition (reduced power), the tip clearance will be greater than is required and performance loss will result. This is particularly true when we consider that in most aircraft missions, a good percentage of the flying time occurs in the cruise condition.

It is therefore an object of this invention to provide a turbomachine which operates at increased overall efficiency and performance.

Accordingly, the invention provides in a turbomachine having a rotatable turbine or compressor and a closely surrounding shroud structure, an improved method of operating over a cruise range of speeds, comprising the steps of a) determining when the turbomachine is operating within a predetermined cruise speed range; and b) responsively cooling the shroud structure to cause thermal shrinkage thereof so as to thereby reduce the operational clearance between the turbine or compressor and the shroud structure.

The invention will now be described by way of example with reference to the accompanying drawing, wherein; Figure 1 is a partial longitudinal section view of the turbine portion of a turbomachine incorporating the present invention in accordance with the preferred embodiment.

Figure 2 is a graphic representation of a steady state turbine blade tip clearance as shown over a range of operating speeds.

Referring now to Figure 1, the invention is shown generally at 10 as being installed in the high pressure turbine portion of an engine having a single stage row of rotor blades or buckets 11, rotatably disposed in a flow path of the hot gases which flow from an annular combustor 12 rearwardly to a row of circumferentially spaced high pressure nozzles 13, through the circumferentially spaced row of buckets 11, through a stationary row of low pressure nozzles 14, and finally to flow downstream as indicated by the arrow to the low pressure turbine (not shown).

The combustor casing 16 and the combustor liner 17 mutually define the annular plenums 18 and 19 which receive air from the compressor discharge and provide cooling air to the high pressure turbine nozzles 13. The cooling air from the plenum 19 passes through the compartment 21 and into the cavity 22 forward of the turbine disc 23 as indicated by the arrows.

The air in the cavity 22 acts to cool the turbine disc 23 and to pressurize the cavity 22 so as to prevent leakage of the hot gases radially inwardly from the main flow stream.

In a manner similar to the high pressure nozzle 13, the low pressure nozzle 14 is also cooled by the flow of cooling air thereover.

This is generally accomplished by routing cooling air from a compressor bleed manifold 24 into a cooling plenum 28 from where it flows over the shroud through an impingement ring 29, into the annular compartment 31 and then down through the low pressure nozzle 14 to the cavity 32 below, where it assists in cooling the turbine disc 23 and pressurizes the cavity 32 to prevent the flow of hot gases inwardly from the main flow stream.

Closely surrounding the circumferentially spaced row of high pressure turbine buckets 11 is a shroud 33 which preferably has as part of its inner surface an abradable material which will allow the high pressure turbine buckets 11 to rub and abrade during certain conditions of operation. The shroud 33 is radially positioned and supported by forward and aft flanges 34 and 36 extending radially inward from a shroud support element 37 which in turn is attached to the combustor casing 16 by a plurality of bolts 38. It will be recognized that, since the shroud 33 is connected directly to the shroud support element 37, any variation in the temperature of the shroud support 37 will cause a thermal growth or shrinkage thereof and a resultant radial growth or shrinkage of the turbine shroud 33.If the shroud support 37 is continuously exposed directly to the compressor bleed in the cooling plenum 28, then it will tend to grow as the engine speeds increase (hotter bleed air) and shrink as the engine speeds decrease (cooler bleed air). This relationship is, of course, helpful when we consider the overall problem of maintaining proper clearances over a range of speeds, but taken alone it is not sufficient to provide for operation under transient conditions.

To accommodate these transient conditions, various other methods may be employed, such as for example, the application of hot turbine discharge air to the shroud support 37 to allow a quick increase in speed accompanied by an increase in temperature to cause the support to grow rapidly, and a rapid decrease in temperature when the engine is suddenly decelerated so as to cause a rapid shrinkage of the support element 37. However, whatever the means or method employed, to regulate the shroud position in order to accommodate transient operations, the turbine tip clearance 39 between the high pressure buckets 11 and the shroud 33 will be excessive when the engine is finally operating in a steady-state condition.It is primarily this condition for which this invention is intended, but, of course, the invention may also be applicable during nonsteady-state operating conditions if certain override provisions are made.

Disposed radially outside the shroud support element 37 is an impingement plate 41 having a plurality of holes 42 formed therein, in a direction and location so as to have the capability of impinging air on the shroud support element 37 as shown by the arrows in Figure 1. The impingement plate 41, which is connected at its ends to the shroud support element 37, partially defines an inner cavity 43 on its one side and an outer cavity 44 on its other side. The outer cavity 44 has an annular cover plate 46 into which a single or a plurality of cooling tubes 47 are connected to bring cooling air into the outer cavity 44. The cooling tubes 47 are, in turn, connected at their other ends to an air supply 50 such as, for example, the fan discharge or compressor bleed air.Airflow to the cooling tubes 47 is controlled by a valve 48 which is in turn operated by a control 49. The control 49 may be of the manual type wherein an operator opens or closes the valve 48 in response to the existence of certain engine conditions such as certain speeds or temperatures.

However, it may also be of the automatic type which operates in response to a speed or temperature sensor or a stator angle indicator to modulate the valve 48.

In operation, air from the compressor bleed 24 will charge the cooling plenum 28 and flow down into the cavity 32 during all conditions of engine operation. Cooling air from the supply 50 will flow through the cooling tubes 47 only during the periods in which the valve 48 is turned on. It is contemplated by this invention that the valve 48 will be turned on when the engine is operating in a speed range that is at or close to the designed cruise speed of the engine. Within this speed range, and with the inventive system of the present invention turned off (valve 48 closed), the turbine tip clearance 39 will be in excess of that required for such engine operation and will thereby result in an appreciable loss of performance.It is the intent of the present invention to reduce that clearance to a ininimum to thereby "fine tune" the system to obtain maximum performance. This is accomplished by determining when the engine speed approaches the engine-designed cruise speed, and then opening the valve^48 to provide cooling air to the shroud support 37 to thereby shrink it down to obtain the close clearance relationship with the turbine bucket row 11.

After passing through the impingement holes 42, into the inner cavity 43, and onto the shroud support 37, the cooling air will flow through the channel 49a toward the cavity 32.

When the speed of the engine is further increased to a point where it begins to exceed by a predetermined amount the designed engine cruise speed, then the valve 48 is closed and the system is turned off so as to allow the shroud support element 37 to again expand. This expansion will then provide the allowance for mechanical and thermal growth which is necessary when accelerating to maximum power settings.

When the system is operated in the manner described hereinabove, the clearance relationship between the turbine blades 11 and the surrounding shroud 33 will be modulated in the manner shown in Figure 2. It should be mentioned that the core speed and radial positions are indicated for engine steady-state operation, and if operation is contemplated over transient conditions, the relationship becomes much more complicated. However, since'the subject invention is intended primarily for steady-state conditions, the graph in Figure 2 will suffice for explanation purposes. It will be recognized that the radial clearance D between the rotor and the shroud when the impingement system is off, is required in order to accommodate transient condition operation, the required clearance decreasing as the speed of the engine increases.It will also be recognized that the clearance at full power is designed to be as low as possible. However, for operation at or near the designed cruise speed, the clearance y for steady-state conditions of operation is excessive and may be reduced by turning on the impingement cooling system and reducing the size of the shroud to allow for a minimum clearance z with the rotor. For purposes of explanation and for purposes of interpreting the claims, designed cruise speed is defined as a predetermined speed or speed range at which an engine is intended to operate in a steady-state condition over a substantial part of its flight schedule.

The particular type of aircraft and aircraft mission, as well as operating efficiency and economy, are primary factors in this determination. In relation to the operation as described above then, the system would sense from the associated compressor bleed temperature, the core speed A and would turn on the valve 48.

The system would remain on until the engine reached a speed B, wherein the increased compressor bleed temperature would cause the system to turn off the valve 48 and allow the shroud to expand to thereby increase the tip clearance sufficiently to allow for a maximum power setting.

It will, of course, be understood that, in addition to the temperature sensing method described hereinabove, there are various other methods by which the system may be responsively turned on and off. For example, it may be desirable to sense the engine speed directly as by a tachometer or the like. Another alternative would be for the valve to be selectively activated manually by an operator when certain conditions are met, or it could occur automatically in response to the engine throttle position.

Further, in order to accommodate transient condition operation within the cruise speed range for which the present cooling apparatus is designed to be operable, various lock-out devices such as, for example, delay mechanisms, may be employed so as to essentially lock out the system during transient periods of operation Also, it should be pointed out that the valve 48 may be of the on-off type or it may be of the variable flow type, wherein the degree to which it is open, and thus the amount of cooling air that flows, is dependent on the speed, temperature or other parameter of the engine. Of course, instead,of a single valve, there may be a plurality of valves operating simultaneously.

It will be understood that while the present invention is being described in terms of one embodiment, it may take on any number of other forms for remaining within the scope and intent of the invention. For example, while the present invention has been described in terms of use for speed ranges proximate the design cruise speed of the engine, it will be recognized that it could just as well be employed for different speed ranges such as, for example, loitering speed, take-off, climb, etc. Further, although described as applicable to a single turbine stage, the present system may be used for applying cooling air to a number of turbine stages or similarly to any number of compressor stages.

Further examples include the possible use of various other engine parameters, such as airflow or pressure, to which the valve may be made responsive, and the possible use of other cooling air sources such as, for example, fan air, for cooling.

It should also be understood that the cooling effect of the shroud structure may be brought about by the removal of relatively hot air rather than the injecting of a flow of relatively cool air. For example, in order to provide good transient matching of the stator and rotor diameters, high temperature air may be continuously circulated around the shroud structure. The present invention would provide for the removal of such hot air application during predetermined periods of operation.

Still another example of alternative structure would be the use of a cooling method other than that of the impingement method as shown, such as, for example, convection flow cooling.

Reference should be made to the description and claims of our co-pending application No.

15771/77 (Serial No. 1581566).

WHAT WE CLAIM IS: 1. A method of operating over a range of speeds a turbomachine having a rotatable turbine or compressor and a closely surrounding shroud structure, comprising the steps of: determining when the turbomachine is within a first predetermined cruise speed operating range; responsively cooling the shroud structure to cause thermal shrinkage thereof so as to thereby reduce the operational clearance between the turbine or compressor and the shroud structure, and determining when the turbomachine is within a second predetermined higher operating range and responsively discontinuing the cooling of the shroud support structure to cause thermal growth thereof so as to thereby increase the operational clearance between the turbine or compressor and the shroud structure.

2. The method as set forth in claim 1 wherein the turbomachine is a gas turbine engine and said first predetermined operating range is indicative of the engine cruise speed and said second predetermined operating range is indicative of a speed outside the range of the engine cruise speed.

3. The method as set forth in claim 1 wherein said first predetermined operating range is determined by sensing a temperature condition within the turbomachine.

4. The method as set forth in claim 3 wherein the turbomachine includes a compressor and cooling is accomplished by the use of bleed air from said compressor, and further wherein the temperature condition sensed is that of the bleed air from the compressor.

5. The method as set forth in claim 1 wherein the step of cooling is accomplished by way of routing a supply of cooling air to the shroud structure.

6. The method as set forth in claim 5 wherein the turbomachine includes a compressor with bleed air and further wherein said cooling air supply is said compressor bleed air.

7. In a turbomachine having a rotatable turbine or compressor and a closely surrounding shroud structure, an improved method of operating over a cruise range of speeds, comprising the steps of a) determining when the turbomachine is operating within a predetermined cruise speed range; and b) responsively cooling the shroud structure to cause thermal shrinkage thereof so as to thereby reduce the operational clearance between the turbine or compressor and the shroud structure.

8. The method as set forth claim 7 and including an additional step of determining when the turbomachine is operating at a speed higher than said predetermined cruise speed range and responsively discontinuing the cooling of the shroud support to cause thermal growth thereof so as to thereby increase the operational clearance between the turbine or compressor and the shroud structure.

9. The method as set forth in claim 8 wherein said determination of operation within a predetermined cruise speed range is made by sensing a temperature condition within the turbomachine.

10. The method as set forth in claim 9 wherein the turbo machine includes a compressor and cooling is accomplished by the use of bleed air from said compressor and further wherein the temperature condition sensed is that of the bleed air from said compressor.

11. The method as set forth in claim 7 wherein the step of cooling is accomplished by way of routing a supply of cooling air to the shroud structure.

12. The method as set forth in claim 11 wherein the turbomachine includes a compressor with bleed air and further wherein said cooling air supply is said compressor bleed air.

13. A method of operating a turbomachine as hereinbefore described with reference to the accompanying drawings.

14. A turbine type power plant having an engine casing, rotary machinery and seal means, said seal means including means for squirting cool air on the engine casing for impingement cooling thereof, and control means for turning on and off the cool air squirting means.

15. A power plant as claimed in claim 14, wherein the quirting means is external of the casing.

16. A power plant as claimed in claim 14, including means for supporting the seal on the casing.

17. A power plant as claimed in claim 14, wherein the control means responds to an engine operating parameter.

18. A power plant as claimed in claim 17, wherein the parameter is compressor speed.

19. A power plant as claimed in claim 14, including a fan discharge duct and a connection between the fan discharge duct and the cool air quirting means.

20. A power plant as claimed in claim 14, wherein the rotary machine is a turbine.

**WARNING** end of DESC field may overlap start of CLMS **.

Claims

**WARNING** start of CLMS field may overlap end of DESC **.

such as, for example, convection flow cooling.

Reference should be made to the description and claims of our co-pending application No.

15771/77 (Serial No. 1581566).

WHAT WE CLAIM IS: 1. A method of operating over a range of speeds a turbomachine having a rotatable turbine or compressor and a closely surrounding shroud structure, comprising the steps of: determining when the turbomachine is within a first predetermined cruise speed operating range; responsively cooling the shroud structure to cause thermal shrinkage thereof so as to thereby reduce the operational clearance between the turbine or compressor and the shroud structure, and determining when the turbomachine is within a second predetermined higher operating range and responsively discontinuing the cooling of the shroud support structure to cause thermal growth thereof so as to thereby increase the operational clearance between the turbine or compressor and the shroud structure.
2. The method as set forth in claim 1 wherein the turbomachine is a gas turbine engine and said first predetermined operating range is indicative of the engine cruise speed and said second predetermined operating range is indicative of a speed outside the range of the engine cruise speed.
3. The method as set forth in claim 1 wherein said first predetermined operating range is determined by sensing a temperature condition within the turbomachine.
4. The method as set forth in claim 3 wherein the turbomachine includes a compressor and cooling is accomplished by the use of bleed air from said compressor, and further wherein the temperature condition sensed is that of the bleed air from the compressor.
5. The method as set forth in claim 1 wherein the step of cooling is accomplished by way of routing a supply of cooling air to the shroud structure.
6. The method as set forth in claim 5 wherein the turbomachine includes a compressor with bleed air and further wherein said cooling air supply is said compressor bleed air.
7. In a turbomachine having a rotatable turbine or compressor and a closely surrounding shroud structure, an improved method of operating over a cruise range of speeds, comprising the steps of a) determining when the turbomachine is operating within a predetermined cruise speed range; and b) responsively cooling the shroud structure to cause thermal shrinkage thereof so as to thereby reduce the operational clearance between the turbine or compressor and the shroud structure.
8. The method as set forth claim 7 and including an additional step of determining when the turbomachine is operating at a speed higher than said predetermined cruise speed range and responsively discontinuing the cooling of the shroud support to cause thermal growth thereof so as to thereby increase the operational clearance between the turbine or compressor and the shroud structure.
9. The method as set forth in claim 8 wherein said determination of operation within a predetermined cruise speed range is made by sensing a temperature condition within the turbomachine.
10. The method as set forth in claim 9 wherein the turbo machine includes a compressor and cooling is accomplished by the use of bleed air from said compressor and further wherein the temperature condition sensed is that of the bleed air from said compressor.
11. The method as set forth in claim 7 wherein the step of cooling is accomplished by way of routing a supply of cooling air to the shroud structure.
12. The method as set forth in claim 11 wherein the turbomachine includes a compressor with bleed air and further wherein said cooling air supply is said compressor bleed air.
13. A method of operating a turbomachine as hereinbefore described with reference to the accompanying drawings.
14. A turbine type power plant having an engine casing, rotary machinery and seal means, said seal means including means for squirting cool air on the engine casing for impingement cooling thereof, and control means for turning on and off the cool air squirting means.
15. A power plant as claimed in claim 14, wherein the quirting means is external of the casing.
16. A power plant as claimed in claim 14, including means for supporting the seal on the casing.
17. A power plant as claimed in claim 14, wherein the control means responds to an engine operating parameter.
18. A power plant as claimed in claim 17, wherein the parameter is compressor speed.
19. A power plant as claimed in claim 14, including a fan discharge duct and a connection between the fan discharge duct and the cool air quirting means.
20. A power plant as claimed in claim 14, wherein the rotary machine is a turbine.