US20100329887A1

US20100329887A1 - Coolable gas turbine engine component

Info

Publication number: US20100329887A1
Application number: US12/575,171
Authority: US
Inventors: Andy Eifert
Original assignee: Rolls Royce Corp
Current assignee: Rolls Royce Corp
Priority date: 2009-06-26
Filing date: 2009-10-07
Publication date: 2010-12-30

Abstract

A gas turbine engine component coupled with a temperature member configured to assist in maintaining a temperature of the gas turbine engine component below a predetermined temperature. The temperature member can take the form of a phase changeable material that can change phase by melting or vaporization, among potential others. A variety of gas turbine engine components can be coupled with the temperature member.

Description

CROSS REFERENCE TO RELATED APPLICATIONS

The present application claims the benefit of U.S. Provisional Patent Application 61/220,791, filed Jun. 26, 2009, and is incorporated herein by reference.

TECHNICAL FIELD

The present invention generally relates to gas turbine engines, and more particularly, but not exclusively, to coolable gas turbine engine components.

BACKGROUND

Gas turbine engines include many components that operate in relatively high heating environments. Providing these and other gas turbine engine components with structures capable of providing cooling remains an area of interest. Some existing systems have various shortcomings relative to certain applications. Accordingly, there remains a need for further contributions in this area of technology.

SUMMARY

One embodiment of the present invention is a unique construction of a gas turbine engine component coupled with a phase changeable material. Other embodiments include apparatuses, systems, devices, hardware, methods, and combinations for cooling gas turbine engine components with phase changeable materials. Further embodiments, forms, features, aspects, benefits, and advantages of the present application shall become apparent from the description and figures provided herewith.

BRIEF DESCRIPTION OF THE FIGURES

FIG. 1 depicts one form of a gas turbine engine.

FIG. 2 depicts an embodiment of a gas turbine engine component coupled with a phase changeable material.

FIG. 3 depicts an embodiment of a gas turbine engine casing coupled with a phase changeable material.

FIG. 4 depicts an embodiment of a gas turbine engine component coupled with a phase changeable material.

FIG. 5 depicts an embodiment of a gas turbine engine blade.

FIG. 6 depicts an embodiment of a gas turbine engine blade coupled with a phase changeable material.

DETAILED DESCRIPTION OF THE ILLUSTRATIVE EMBODIMENTS

For the purposes of promoting an understanding of the principles of the invention, reference will now be made to the embodiments illustrated in the drawings and specific language will be used to describe the same. It will nevertheless be understood that no limitation of the scope of the invention is thereby intended. Any alterations and further modifications in the described embodiments, and any further applications of the principles of the invention as described herein are contemplated as would normally occur to one skilled in the art to which the invention relates.
With reference to FIG. 1, there is illustrated a schematic representations of one form of a gas turbine engine 50 used as a powerplant for an aircraft. The present application contemplates that a gas turbine engine may be, but is not limited to, an expendable or non-expendable engine. As used herein, the term “aircraft” includes, but is not limited to, helicopters, airplanes, unmanned space vehicles, unmanned combat aerial vehicles, fixed wing vehicles, variable wing vehicles, rotary wing vehicles, tailless aircraft, hover crafts, and other airborne and/or extraterrestrial (spacecraft) vehicles. Further, the present inventions are contemplated for utilization in other applications that may not be coupled with an aircraft such as, for example, industrial applications, power generation, pumping sets, naval propulsion and other applications known to one of ordinary skill in the art.
The gas turbine engine 50 includes a casing 52 that encloses a compressor 54, a combustor 56 and a turbine 58. In the illustrative embodiment the gas turbine engine 50 is in the form of a turbojet engine, but in other embodiments the gas turbine engine 50 can take other forms such as, but not limited to, turbofans, tubojets, turboshafts, and turboprops.
With reference to FIG. 2, there is illustrated a non-limiting schematic view of a gas turbine engine component 60. The term gas turbine engine component is intended to broadly include a number of components of the gas turbine engine examples of which include, but are not limited to, engine cases, cooling lines, combustor liners, exhaust nozzles, exhaust liners, valves, solenoids, and/or other cooled components. In one form the gas turbine engine component 60 includes a body or base 62 having a surface 64 and coupled to a temperature member 66 operable to maintain the base 62 within an operably important temperature (discussed below) when the gas turbine engine component 60 is subjected to a heating event. In one aspect the heating event can be caused by a failure of a cooling system, or can be caused by a relatively high temperature working environment, to set forth just two non-limiting examples. In one form the temperature member 66 can be disposed between the base 62 and a heating event such as, but not limited to, a relatively high temperature gas. In other forms, however, the base 62 can be disposed between the temperature member 66 and a heating event. The heating event can transfer heat to either the temperature member 66 and/or the base 62 by mechanisms such as convection, conduction, and radiation.
Temperature member 66 is formed at least in part from a material operable to undergo a change in phase such as by melting and vaporization, to set forth just two types of phase change. In some embodiments the temperature member 66 can be made from metals such as lead or tin. In other embodiments, the temperature marker can be made from wax. Other materials are also contemplated herein. The material composition of the temperature member 66 can be selected such that its phase change temperature is relatively lower than a predetermined temperature of the base 62 such as a critical temperature, fatigue temperature, or safety temperature, among other types of operably important temperatures. By having a phase change temperature below an operably important temperature of the base 62, the temperature member 66 can evaporate, melt, or otherwise change phase to maintain the base 62 at a temperature below the operably important temperature at least until the heating event passes or the phase change of the temperature member 66 is complete.
Temperature member 66 can be coupled to the component 60 using a variety of techniques. To set forth just two non-limiting examples, the temperature member 66 can be coupled by coating or casting it to the base 62. Additionally, in some forms surface 64 can be porous, or otherwise include depressions/apertures/openings to facilitate binding or attachment with temperature member 66. In some forms the temperature member 66 can be coupled using fasteners such as bolts, screws, and rivets, among other possibilities.
In some forms the temperature member 66 covers all of surface 64, but in other forms may only cover a portion of surface 64. Additionally, the thickness of the temperature member 66 can be constant, or may vary, across the surface 64 of the base 62. In still other forms, the surface 64 can include multiple temperature members 66, some of which may vary in attributes such as composition and thickness, among potential others. In one non-limiting embodiment, the temperature member 66 can have a relatively higher thickness near a location of the component 60 experiencing a relatively high heating event, and the temperature member 66 can have a relatively thin thickness near a location experiencing a relatively low heating event. Other combinations are also contemplated.
In one form of the present application the temperature member 66 is a sacrificial material and after changing phase can be drawn away from the component 60. In some embodiments the temperature member 66 can be discharged or otherwise exit the gas turbine engine 50 after changing phase. In some forms it is contemplated that the temperature member 66 can be contained in proximity to the component 60 for possible reuse. For example, in one embodiment the gas turbine engine component 60 can include a cover that defines an interior space between the surface 64 and the cover and within which is disposed the temperature member 66. The cover allows for retention of at least a portion of temperature member 66 after the phase change.
FIGS. 3-6 depict various embodiments of a gas turbine engine component and temperature member 66. With reference to FIG. 3, there is illustrated gas turbine engine 50 having a temperature member 66 coupled with a casing 68. In the embodiment illustrated in FIG. 3, temperature member 66 undergoes a phase change from a solid state to a liquid state when its temperature reaches the phase transition temperature. The phase change of the temperature member 66 draws heat away from the casing 68 thus assisting in maintaining the temperature of the casing 68 within an operably important temperature. The temperature member 66 is coupled to an outside surface of the casing 68 in the illustrated embodiment, but in other forms may be coupled to an inside surface of the casing 68.
Coupling the casing 68 of the gas turbine engine 50 with the temperature member 66 can be useful in engine applications because it assists in regulating the temperature of the component without necessarily adding of any additional cooling features, such as an active cooling system. In some embodiments the temperature member 66 can be added to existing engine case designs without modifications to the case.
Referring now to FIG. 4, there is illustrated another embodiment of a gas turbine engine component coupled with the temperature member 66. The gas turbine engine component of FIG. 4 includes a tube 70 having a fluid flow passage 72. Temperature member 66 is disposed on and surrounding the outer surface of the tube 70. Though the temperature member is depicted as coupled to the outer surface of the tube 70, in some forms the temperature member 66 can be coupled to the inside surfaces of the passage 72. In this form, fluid flowing through passage 72 can be in direct contact with the temperature member 66.
In operation, a fluid 74 at a relatively high first temperature enters the tube 70 at first end and flows through the fluid flow passage 72 towards a second end. As fluid 74 flows through fluid flow passage 72, heat is transferred from the fluid 74 and increases the temperature of the tube 70. Heat from the tube 70 is conducted from its inner surface to its outer surface and to temperature member 66. As the temperature of the temperature member 66 increases to the phase transition temperature, the temperature member 66 undergoes a change in phase. This change in phase continues to draw heat from the tube 70. The fluid 74 flowing through passage 72 can thus be cooled through the phase change of the temperature member 66. In one aspect of the present invention, the cooled fluid 74 can then be used to provide cooling to other gas turbine engine components. When the temperature member 66 completes a transition in phase, the fluid 74 flowing through passage 72 will no longer be cooled through the phase change process of the temperature member 66. As a result, after some amount of time the tube 70 will experience an increase in temperature. The thickness, volume, and placement, among other aspects of the temperature member, can be selectable based on desired or predicted operating conditions of the component 60.
With reference to FIGS. 5-6, there is illustrated a view of another embodiment of a gas turbine engine component 60 having a temperature member 66. The gas turbine engine component illustrated in FIGS. 5 and 6 is in the form of an actively cooled blade 76 having a plurality of cooling media exit openings 78 that allow for the discharge of cooling media across an outer surface of the blade 76. While the actively cooled blade 76 includes a number of openings 78, some forms of the blade 76 may not include the openings 78.
With specific reference to FIG. 6, the blade 76 further includes a cavity 80 at least partially filled with the temperature member 66 formed on an inner surface 82 of the cavity 80. The cavity 80 can be coupled with the openings 78 through passages (not shown) formed through the temperature member 66. The temperature member 66 can be evenly distributed around the inner surface 82 of the cavity or can be concentrated in certain locations. For example, if higher heating were expected at a leading edge of the blade 76 then additional thickness of the temperature member 66 can be used. Furthermore, the relative amount of temperature member 66 can vary from the hub of the blade 76 to its tip. In other embodiments, the blade 76 can be coupled with the temperature member 66 in locations other than the cavity 80. For example, the temperature member 66 can be coupled to the inner surfaces of the openings 78, or at selected locations on the exterior of the blade 76.
During operation and as the blade 76 increases in temperature, the temperature member 66 undergoes a phase change at its phase transition temperature. The temperature member 66 can be converted to a gas or liquid which thereafter can be expelled through openings 78 or through other apertures or conduits not otherwise explicitly depicted. In gas turbine engines, the temperature member 66 can be used without any openings 78 and still provide cooling to blade 76 by undergoing a phase change. The temperature member 66 can also be used a backup system to prevent damage to blades 76 in case a cooling media normally available to the blade 76 is otherwise interrupted.
One aspect of the present application provides a gas turbine engine component in thermal communication with a phase changeable material. The phase changeable material can be selected to change phase at a temperature below an operably important temperature of the gas turbine engine component.
Yet another aspect of the present application provides an apparatus comprising a gas turbine engine device including a gas turbine engine component having an envelope of operability and thermodynamically coupled with a thermal portion having size sufficient to prolong a useful life of the gas turbine engine component by keeping the gas turbine engine component within the envelope of operability during a relatively high heating rate mission event, the thermal portion operable to change phases from one physical state to another physical state during at least a portion of the relatively high heating rate mission event.
Still another aspect of the present application provides an apparatus comprising a gas turbine engine including a component subject to a heat stress, a temperature activated portion covering the gas turbine engine portion to protect the component from high temperatures during the heat stress, the temperature activated portion operable to be sacrificed by changing phases from one physical state to another.
Still a further aspect of the present application provides an apparatus comprising a gas turbine engine component operable to be located in a high temperature region of a gas turbine engine and means for sacrificially prolonging the life expectancy of the gas turbine engine component, the means coupled with the gas turbine engine component.
Yet another aspect of the present application provides a method comprising orienting a gas turbine engine component in preparation for coupling the component to a sacrificial temperature sensitive structure, identifying a portion of the gas turbine engine component to be coupled with the sacrificial temperature sensitive structure, selecting a quantity of the phase transition material sufficient for the gas turbine engine component to endure at least a portion of a heat flux event, and thermodynamically coupling the sacrificial temperature sensitive structure to the gas turbine engine component.
While the invention has been illustrated and described in detail in the drawings and foregoing description, the same is to be considered as illustrative and not restrictive in character, it being understood that only the preferred embodiments have been shown and described and that all changes and modifications that come within the spirit of the inventions are desired to be protected. It should be understood that while the use of words such as preferable, preferably, preferred or more preferred utilized in the description above indicate that the feature so described may be more desirable, it nonetheless may not be necessary and embodiments lacking the same may be contemplated as within the scope of the invention, the scope being defined by the claims that follow. In reading the claims, it is intended that when words such as “a,” “an,” “at least one,” or “at least one portion” are used there is no intention to limit the claim to only one item unless specifically stated to the contrary in the claim. When the language “at least a portion” and/or “a portion” is used the item can include a portion and/or the entire item unless specifically stated to the contrary.

Claims

1. An apparatus comprising:

a gas turbine engine device including a gas turbine engine component having an envelope of operability and thermodynamically coupled with a thermal portion having size sufficient to prolong a useful life of the gas turbine engine component by keeping the gas turbine engine component within the envelope of operability during a relatively high heating rate mission event, the thermal portion operable to change phases from one physical state to another physical state during at least a portion of the relatively high heating rate mission event.

2. The apparatus of claim 1, wherein the thermal portion is activated by temperature and is operable to be sacrificed during the relatively high heating rate mission event.

3. The apparatus of claim 1, wherein the gas turbine engine device includes a conduit for the passage of a working fluid, the thermal portion coupled with the conduit and operable to change phase when a temperature of a working fluid traversing the conduit is at or above a phase transition temperature.

4. The apparatus of claim 3, wherein the thermal portion is coupled to an outside wall of the conduit.

5. The apparatus of claim 1, wherein the gas turbine engine device is an engine case, the thermal portion operative to maintain the temperature of the engine case during an engine thermal condition.

6. The apparatus of claim 5, wherein the engine thermal condition occurs during an exigent condition.

7. The apparatus of claim 1, wherein the gas turbine engine device is a turbine vane having an internal cavity, the thermal portion occupying at least part of the internal cavity.

8. The apparatus of claim 1, wherein the gas turbine engine device is a portion of an actuation device, the thermal portion operative to prevent the actuation device from overheating during at least a portion of an engine thermal condition.

9. An apparatus comprising:

a gas turbine engine including a component subject to a heat stress; and

a temperature activated portion covering the gas turbine engine component to protect the component from high temperatures during the heat stress, the temperature activated portion operable to be sacrificed by changing phases from one physical state to another.

10. The apparatus of claim 9, wherein the temperature activated portion is structured to change phase through melting, vaporization, or sublimation.

11. The apparatus of claim 9, wherein the gas turbine engine is designed and manufactured as a single-mission, expendable gas turbine engine, the component subject to the heat stress during at least a portion of the single-mission of the gas turbine engine.

12. The apparatus of claim 9, wherein the temperature activated portion has a size sufficient to extend the lifetime of the gas turbine engine component during an operational exigency.

13. The apparatus of claim 12, wherein the operational exigency is characterized by a failure of a thermal transfer mechanism intended to maintain the gas turbine engine component within a temperature limit.

14. The apparatus of claim 12, wherein the temperature activated portion is sacrificed by having a portion that is separated from the remaining portion of the temperature activated portion.

15. An apparatus comprising:

a gas turbine engine component operable to be located in a high temperature region of a gas turbine engine; and

means for sacrificially prolonging the life expectancy of the gas turbine engine component, the means coupled with the gas turbine engine component.

16. A method comprising:

orienting a gas turbine engine component in preparation for coupling the component to a sacrificial temperature sensitive structure;

identifying a portion of the gas turbine engine component to be coupled with the sacrificial temperature sensitive structure;

selecting a quantity of the phase transition material sufficient for the gas turbine engine component to endure at least a portion of a heat flux event; and

thermodynamically coupling the sacrificial temperature sensitive structure to the gas turbine engine component.

17. The method of claim 16, wherein the selecting includes applying a thickness of the sacrificial temperature sensitive structure appropriate to protect the gas turbine engine component from a heating source for a duration.

18. The method of claim 16, wherein the thermodynamically coupling further includes affixing an amount of the sacrificial temperature sensitive structure to the gas turbine engine component.

19. The method of claim 18, wherein the affixing further includes thermally attaching the sacrificial temperature sensitive structure to a wall defining an airflow conduit.

20. The method of claim 18, wherein the affixing further includes placing the sacrificial temperature sensitive structure in a thermal communication path with a portion of an actuation device.

21. The method of claim 16, wherein the thermodynamically coupling further includes at least partially filling a cavity of the gas turbine component with the sacrificial temperature sensitive structure, the gas turbine engine component in the form of a turbine vane.

22. The method of claim 16, wherein the determining a thickness includes estimating the duration of an expendable aircraft mission.

23. The method of claim 16, wherein the determining a thickness includes estimating the duration of a gas turbine engine exigent condition.