A dampening system for a miniature high-speed turbojet engine provides single or multiple resilient mountings for forward and aft bearings which mount a rotor shaft. The self-induced first engine order of the rotating rotor shaft is transmitted through the forward and aft bearing and is then absorbed in the resilient mountings.

Current U.S. Class:	415/119; 384/536; 415/229
Current CPC Class:	F01D 25/164 (20130101); F16C 27/04 (20130101); F16C 27/066 (20130101); F16C 19/548 (20130101); F05D 2250/82 (20130101); F16C 19/163 (20130101); F16C 2360/23 (20130101)
Current International Class:	F01D 25/16 (20060101); F01D 025/04 ()
Field of Search:	;415/119,229 ;416/244R,244A,500 ;384/535,536,581,582

Carlson, Gaskey & Olds

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Claims

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What is claimed is:

1. A rotor system for a miniature gas turbine engine comprising: an engine housing portion having a groove; a resilient member mounted within said groove; a bearing mounted adjacent said resilient member; a shaft rotationally mounted within said bearing; a compressor supported upon said shaft, said compressor facing forward toward an inlet; and a turbine supported upon said shaft, said turbine facing rearward toward an exhaust.

2. The rotor system as recited in claim 1, wherein said resilient member comprises an elastomeric member.

3. The rotor system as recited in claim 1, wherein said resilient member comprises an O-ring.

4. The rotor system as recited in claim 1, wherein said resilient member comprises an elastomeric strip.

5. The rotor system as recited in claim 1, wherein said groove is formed within an annular sleeve.

6. The rotor system as recited in claim 1, wherein said groove is formed within an annular sleeve mounted adjacent said engine housing portion and a second engine housing portion.

7. The rotor system as recited in claim 6, further comprising a second groove defined within said annular sleeve, a second resilient member mounted within said second groove, said resilient member and said second resilient member adjacent said bearing.

8. The rotor system as recited in claim 6, wherein said annular sleeve is mounted across an interface defined between a first engine housing portion and a second engine housing portion.

9. The rotor system as recited in claim 1, further comprising a second groove defined within said engine housing portion, a second resilient member mounted within said second groove, said resilient member and said second resilient member adjacent said bearing.

10. The rotor system as recited in claim 1, wherein said compressor and said turbine define a monorotor.

11. A miniature gas turbine engine comprising: an engine housing portion defining a first and a second circumferential groove; a first and a second resilient member mounted within said respective first and second circumferential groove; a first bearing mounted adjacent said first and a second resilient member; an annular sleeve mounted across an interface defined between said engine housing portion and a second engine housing portion, said annular sleeve defining a third and a forth circumferential groove; a third and a fourth resilient member mounted within said respective third and fourth circumferential groove; a second bearing mounted adjacent a third and a fourth resilient member; a shaft rotationally mounted within said first and said second bearing; a compressor supported upon said shaft, said compressor facing forward toward an inlet; and a turbine supported upon said shaft, said turbine facing rearward toward an exhaust.

12. The miniature gas turbine engine as recited in claim 11, wherein said compressor and said turbine define a monorotor.

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Description

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BACKGROUND OF THE INVENTION

The present invention relates to a miniature gas turbine engine and, more particularly, to a rotor shaft vibration dampening system.

Miniature gas turbine or turbojet engines (100 lbf thrust and smaller) are often utilized in single usage applications such as reconnaissance drones, cruise missiles, decoy and other weapon applications, including air-launched and ground-launched weapon systems. The use of such an engine greatly extends the range of the weapon in comparison to the more conventional solid fuel rocket engine. Miniature gas turbine engines are difficult to fabricate economically for general expendable usage in large numbers.

To achieve economically feasible extended range expendable propulsion sources for such applications, it is necessary that the gas turbine engines be manufactured relatively inexpensively yet provide a high degree of reliability and efficiency. One component that greatly affects performance yet is rather complicated to manufacture is the rotating component, such as the rotor system which typically includes a shaft mounted turbine and compressor wheel.

Rotor systems of miniature gas turbine engines typically operate at greater than 100,000 rpm. Such high rotations in the relatively small miniature gas turbine engines may generate high first engine order (1st EO) vibration characteristics that can be damaging to the engine, its vehicle, and/or the payload thereof.

Vibration dampening devices are commonly utilized in relatively large turbomachinery and conventional gas turbine engines. These applications typically operate at relatively lower speeds of below 100,000 rpm.

As the relatively large applications provide minimal packaging restrictions and are not designed to be expendable, relatively complicated and expensive dampening systems are feasible. As such, conventional dampening devices are impractical and cost prohibitive for miniature gas turbine engines. Heretofore, miniature gas turbine engines avoided most vibration damping systems altogether.

Accordingly, it is desirable to provide a reliable, inexpensive, and uncomplicated dampening system for an expendable gas turbine engine.

SUMMARY OF THE INVENTION

The dampening system according to the present invention for a miniature gas turbine engine provides single or multiple resilient mountings for the forward and aft shaft bearings. Self-induced first engine order vibrations of the rotating rotor shaft are transmitted through the forward and aft bearing's inner race, common balls and outer race and then absorbed in the resilient mountings.

The present invention therefore provides a reliable, inexpensive, and uncomplicated dampening system for an expendable gas turbine engine.

BRIEF DESCRIPTION OF THE DRAWINGS

The various features and advantages of this invention will become apparent to those skilled in the art from the following detailed description of the currently preferred embodiment. The drawings that accompany the detailed description can be briefly described as follows:

FIG. 1 is a general perspective view an exemplary vehicle embodiment for use with the present invention;

FIG. 2 is a schematic view of a gas turbine engine having a dampening system according to the present invention;

FIG. 3 is an expanded view of the dampening system of FIG. 2; and

FIG. 4 is a graphical representation of the reduction in high first engine order (1st EO) vibration characteristics achieved by a dampening system according to the present invention.

DETAILED DESCRIPTION OF THE PREFERRED EMBODIMENT

FIG. 1 illustrates a general schematic view of a vehicle 100 including a miniature gas turbine engine 10 according to the present invention. The vehicle 100 includes a body 102 and one or more aerodynamic surfaces 104. The engine 10 is coupled to, or within, the body 102. An intake 106 provides air to the engine 10, and an exhaust pipe 108 exhausts the thrust therefrom. The engine 10 of the invention may also be used in other single usage and reusable applications such as reconnaissance drones, cruise missiles, decoys and other weapon and non-weapon applications.

Referring to FIG. 2, the miniature gas turbine engine 10 generally includes a housing 14, a rotor shaft 16 rotationally mounted to a forward bearing 18 and an aft bearing 20, a combustion system 21 and an exhaust pipe (nozzle) 22. The rotor shaft 16 rotates about a longitudinal axis X. It should be understood that rotor systems other than the illustrated monorotor configuration, will also benefit from the present invention. In the illustrated rotor configuration, a rotor system 24 includes compressor blades 26 facing forward toward an inlet 28 and turbine blades 30 facing rearward toward the exhaust 22 to define a turbine wheel. The forwardly extending shaft 16 is received in the bearings 18, 20 and is preferably coupled to a fuel pump (illustrated schematically at 32) to provide fuel to an annular combustor liner 34 through a fuel manifold 36.

Referring to FIG. 3, an expanded view of the rotor system 24 is illustrated. The forward bearing 18 and the aft bearing 20 are installed on the shaft 16. The forward bearing 18, and aft bearing 20, are separated by a bearing spacer 38 and are held in place by a rotor nut 40. The forward bearing 18 is pre-loaded with a preload spring 42 and bearing preload spacer 44. It should be understood that although a particular component arrangement is disclosed in the illustrated embodiment, other arrangements will benefit from the instant invention.

The forward and aft bearings 18, 20 are preferably in contact with a dampening system 46 located between each bearing 18, 20 and the housing 14 or other static structure. That is, the bearings 18, 20 are closely mounted within a bore 49 in the housing 14 but are supported upon the dampening system 46. It should be understood that other bearing mounting static structure will also benefit from the present invention.

Preferably, each dampening system 46 includes one or more grooves 48 within the housing 14 which receives a resilient member 50 such as O-rings and/or elastomeric strips. It should be understood that other elastomeric, non-metallic, plastic, teflon, composite or other spring type resilient members will benefit from the present invention. The resilient members 50 provide for vibration absorption or dampening of the shaft 16 1.sup.st EO characteristics (FIG. 4).

Alternatively or in addition, an annular sleeve 52 includes the grooves 48 and the resilient member 50. The annular sleeve 52 is particularly preferred to provide a single mounting surface for a forward bearing 18 which is often located at an interface I between multiple engine housing portions such as the housing 14 and a forward inlet cap 54. The annular sleeve 52 thereby bridges the interface I to provide a stable and consistent mounting location for the bearing 18.

The foregoing description is exemplary rather than defined by the limitations within. Many modifications and variations of the present invention are possible in light of the above teachings. The preferred embodiments of this invention have been disclosed, however, one of ordinary skill in the art would recognize that certain modifications would come within the scope of this invention. It is, therefore, to be understood that within the scope of the appended claims, the invention may be practiced otherwise than as specifically described. For that reason the following claims should be studied to determine the true scope and content of this invention.

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