Power plants – Internal combustion engine with treatment or handling of... – Material from exhaust structure fed to engine intake
Reexamination Certificate
1999-09-09
2001-06-05
Freay, Charles G. (Department: 3746)
Power plants
Internal combustion engine with treatment or handling of...
Material from exhaust structure fed to engine intake
C239S265330, C239S265370
Reexamination Certificate
active
06240720
ABSTRACT:
The invention herein described was made in the course of or under a contract, or a subcontract thereunder, with the United States Department of the Air Force.
FIELD OF THE INVENTION
This invention relates to gas turbine engines, and, more particularly, to the structure of the exhaust nozzle compression link used on some engines.
BACKGROUND OF THE INVENTION
In an aircraft gas turbine (jet) engine, air is drawn into the front of the engine, compressed by a shaft-mounted compressor, and mixed with fuel. The mixture is burned, and the hot exhaust gases are passed through a turbine mounted on the same shaft. The flow of combustion gas turns the turbine by impingement against an airfoil section of the turbine blades and vanes, which turns the shaft and provides power to the compressor. The hot exhaust gases flow from the back of the engine, driving it and the aircraft forwardly.
In some gas turbine engines, there is an exhaust nozzle flap structure at the back of the engine. The exhaust nozzle flap structure is formed of a series of individual exhaust nozzle flap segments arranged in a generally cylindrical fashion around the periphery of the engine at its exhaust end. The exhaust nozzle flap segments may be pivoted inwardly or outwardly, so as to change the size of the outlet passage through which the hot exhaust gases flow. This change in geometry alters the performance of the engine in a controllable manner. Each exhaust nozzle flap segment includes inner flaps which contact the hot combustion gas flow to shape it, and an outer flap which ensures aerodynamic efficiency.
Each of the exhaust nozzle flap segments is supported on a pivot structure at its forwardmost end. The inner and outer flaps are connected to each other and to the remainder of the engine structure in the form of a geometric linkage. The linkage includes an exhaust nozzle compression link. When the exhaust nozzle flap segment is moved by an actuator, the exhaust nozzle compression link aids in maintaining the correct geometrical relation of the inner flaps and the outer flap. The exhaust nozzle compression link is subjected to large compressive loads by the aerodynamic forces exerted by the combination of the hot combustion gas on the inner flaps and the external air flow on the outer flap.
The exhaust nozzle compression link is a vital component in the actuation of the flap structure. In conventional construction, the exhaust nozzle compression link is a rectangular-section hollow strut with a fitting at each end. The exhaust nozzle compression link is made of a nickel-base superalloy in order to withstand heating to about 800° F. resulting from the hot exhaust gases, while providing the necessary compressive strength to support the flap structure and the aerodynamic loading.
There is always a desire to reduce the weight of aircraft gas turbine engines, while retaining acceptable performance. Weight reductions in the area of the exhaust nozzle are highly beneficial to aircraft performance and maneuverability, because of the distance of the exhaust nozzle from the center of gravity of the aircraft. The need to reduce weight to achieve these benefits extends to individual components, such as the exhaust nozzle compression link. The present invention fulfills the need for reduced weight and acceptable performance, and further provides related advantages.
BRIEF SUMMARY OF THE INVENTION
The present invention provides a gas turbine exhaust nozzle compression link with fully acceptable performance and with about 40 percent less weight than a conventional exhaust nozzle compression link. The exhaust nozzle compression link of the invention has substantially the same fit and function as the conventional exhaust nozzle compression link. The design of the exhaust nozzle compression link allows it to utilize composite-materials technology to take advantage of its high elastic modulus and strength at moderately elevated temperatures of about 800° F.
A hybrid-composite gas turbine exhaust nozzle compression link comprises an elongated hollow tubular shell having an axis of elongation. The elongated hollow tubular shell has a structure in cross section taken perpendicular to the axis of elongation comprising at least one reinforcing layer (and preferably from one to ten reinforcing layers) comprising reinforcing fibers extending parallel to the axis of elongation, and a body comprising at least one titanium-base alloy in which the reinforcing fibers are disposed.
In one embodiment, the hybrid-composite exhaust nozzle compression link comprises an inner rim of an inner-rim titanium-base alloy, a reinforcing layer (preferably one to ten reinforcing layers) comprising reinforcing fibers extending parallel to the axis of elongation, and an outer rim of a outer-rim titanium-base alloy. The reinforcing layer preferably further includes a matrix in which the reinforcing fibers are disposed. The reinforcing fibers are preferably silicon carbide, and the matrix is preferably a matrix titanium-base alloy. The inner-rim titanium-base alloy, the outer-rim titanium-base alloy, and the matrix titanium-base alloy, where present, are preferably, but not necessarily, of the same nominal composition.
The elongated hollow tubular shell is preferably in the form of a hollow right circular cylinder. There are attachment fittings at each end of the elongated hollow tubular shell. Desirably, one of the attachment fittings is formed integrally with the tubular shell, and the other attachment fitting is joined to the tubular shell by an adjustable joint such as a threaded attachment.
The structure having a rim and reinforcing fibers extending parallel to the axis of elongation allows the hollow tubular shell to carry large compressive loadings without buckling. In most applications involving reinforcing fibers, the imposed loadings are tensile. The present approach has a cross-sectional rigidity that is 25 percent or more higher than that of pure titanium, resulting in a higher buckling load capacity and reduced weight for the structure. The preferred embodiment places the reinforcing fibers between the rims of titanium-base alloys, which constrains the reinforcing fibers against buckling but still utilizes their high elastic modulus and high strength. The reinforcing fibers are protected against damage of the hot exhaust gases and impact damage by the titanium-base alloys that overlie them.
The composite material of high-modulus, high-strength reinforcing fibers such as silicon carbide fibers in a titanium-alloy matrix is carefully selected to provide the required elastic modulus and strength in compressive loading at the operating temperature of about 800° F., while substantially reducing the weight of the exhaust nozzle compression link. The elastic modulus of the composite material provides a higher resistance to buckling due to the compressive loads that the exhaust nozzle compression link must carry in the exhaust nozzle flap structure.
Other features and advantages of the present invention will be apparent from the following more detailed description of the preferred embodiment, taken in conjunction with the accompanying drawings, which illustrate, by way of example, the principles of the invention. The scope of the invention is not, however, limited to this preferred embodiment.
REFERENCES:
patent: 6142416 (2000-11-01), Markstein et al.
Tseng Wu-Yang
Wojciechowski Charles R.
Freay Charles G.
General Electric Company
Hess Andrew C.
Narciso David L.
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