WO2006010200A1

WO2006010200A1 - Turbine engine

Info

Publication number: WO2006010200A1
Application number: PCT/AU2005/001061
Authority: WO
Inventors: Roger Clyde Webb
Original assignee: Poly Systems Pty Ltd
Priority date: 2004-07-29
Filing date: 2005-07-19
Publication date: 2006-02-02

Abstract

A turbine engine (10) comprising an engine body (3) disposed within a nacelle (2) and extending from adjacent the inlet end of the nacelle (2) to adjacent the outlet end (7) of the nacelle (2). The engine body (3) housing a combustion chamber (6) disposed intermediate of a plurality of forward turbine rotors (4) and a plurality of rear turbine rotors (5). A fluid passageway is substantially coiled about or within the engine body (3) from at least a forward end of the combustion chamber to the outlet end (7) of the engine body (3) and is adapted to receive fluid at a forward end via a fluid inlet (12) and to inject the fluid from the outlet end (13) of the fluid passageway into at least one of the rear turbine rotors (5).

Description

TURBINE ENGINE

Technical Field

The invention relates to a turbine engine, and a means by which exhaust gases passing there through are temperature moderated, such that the turbine engine operates as a "bi- turbine engine."

The invention has been developed primarily with respect to a turbine engine used on aircraft and will be described hereinafter with reference to this application. However, it will be appreciated that the invention is not limited to this particular field of use and is applicable to any vehicle having a turbine engine.

Background

Jet turbo-fan engines are very commonly used in military and non-military aircraft, most notably on large commercial aircraft. These engines are mounted to the aircraft, typically one or more on the underside of each wing. The turbine engines are turbo fan turbine engines which include an additional fan mounted to the output shaft of the turbine engine at a front end.

Air is compressed by the turbo fan and passed intermediate the engine and an engine cowling that extends from just forward of the turbo fan to cover approximately 80% the length of the turbine engine. Air is then further compressed within the engine by a plurality of turbine compressor rotors disposed intermediate the turbo fan and the combustion chamber. Combusted gases from the combustion chamber are exhausted through a rear set of turbine compressor rotors out of the back of the engine. The exhaust plume expands providing energy to the rear rotors and mixes with the by-pass airflow beyond the rear of the cowling and engine at the rear of the engine.

The turbine engines emit exhaust gases that are substantially hotter than the ambient air temperature into which they are expelled. The higher the rate of fuel being consumed by the engine, the more gases are emitted and the hotter the engine gets. In the case of jet turbo-fan engines, for example, an electric motor is typically used at start up to drive the

turbo fan and the turbine compressor rotors until the air has reached a sufficient pressure and the engine is ignited. The controlled expansion of ignited fuel in the combustion chamber is used to provide thrust and power to both the aircraft generally and the turbine compressor rotors. The gas is expelled via an exhaust outlet at the rear of the engine and part of the exhaust is used to drive the rear turbine compressor rotors.

The exhaust gases are expelled into an ambient environment being the atmosphere. As exhaust gases from the engines can be of the order of many hundred degrees Celsius, being significantly hotter than atmospheric gas temperature, there is provided a relatively high contrast 'heat signature' which is comprised of heated exhaust gas and heated atmospheric gases.

The heat of the engine itself also provides a substantial heat signature having a contrast sufficient to be targeted by a heat-seeking missile. This is the case even though engine or cowlings are used to minimise this problem by channeling air over the turbine engine. A further potential problem, albeit relatively minor with respect to the heat signature produced by the exhaust of the turbine engine, is the heating of a wing to which the turbine is mounted and/or proximate.

Both commercial and military jet turbo-fan engine aircraft operate in hostile and potentially hostile territory on a regular basis. As such, these aircraft are extremely vulnerable against attack from remotely fired missiles and projectiles that track an infrared emission or heat signature of an aircraft. The significant difference in temperature between the exhaust gases and the atmosphere provides a high contrast source for such missiles to lock on to.

Although counter-measures such as the strategically timed release of flares and other such devices are used by aircraft to avoid heat-seeking missiles, they only can play a very small part in providing a defence against such missiles, and at the very last minute. It is an object of the present invention to provide an engine that will overcome or substantially ameliorate one or more of the deficiencies of the prior art, or to provide a useful alternative.

Summary of Invention

According to a first aspect of the invention there is provided a turbine engine comprising: an engine body disposed within a nacelle and extending from adjacent the inlet end of said nacelle to adjacent the outlet end of said nacelle, said engine body housing a combustion chamber disposed intermediate of a plurality of forward turbine rotors and a plurality of rear turbine rotors; and wherein a fluid passageway is substantially coiled about or within said engine body from at least a forward end of the combustion chamber to the outlet end of said engine body, said fluid passageway is adapted to receive fluid at a forward end via a fluid inlet and to inject said fluid from the outlet end of said fluid passageway into at least one of said rear turbine rotors.

Preferably said engine body extends a first predetermined distance beyond said rear turbine rotors, and said fluid passageway extends at least partially therealong.

Preferably said fluid substantially comprises of water.

Preferably said plurality of rear turbine rotors comprises at least three rear turbine rotors.

Preferably the outlet of said fluid passageway is disposed adjacent at least one of the three rear-most rear turbine rotors.

Preferably a first turbo fan is mounted to a turbine engine shaft to pressurise air entering said engine.

Preferably in one embodiment an outer body is disposed substantially about said engine body. Preferably said nacelle extends a second predetermined distance beyond said outlet end of said engine body.

Preferably a second turbo fan is mounted adjacent the rear of said nacelle and the rear of said engine body.

In one preferred embodiment at least a portion of said fluid passageway is disposed within a fluid coil located about said engine body.

In another preferred embodiment at least a portion of said fluid passageway is cast or embedded within said engine body.

In another preferred embodiment at least a portion of said fluid passageway is disposed within said outer body.

According to a second aspect of the invention there is provided a cooling system for a turbine engine having a turbine body disposed within a nacelle, said turbine body extending from adjacent the inlet end of said nacelle to adjacent the outlet end of said nacelle, said turbine body having a combustion chamber disposed intermediate of forward and rear turbine rotors, said cooling system comprising a fluid passageway disposed about or within said turbine body from at least a forward end of the combustion chamber to the outlet end of said engine body, said fluid passageway is adapted to receive fluid at a forward end via a fluid inlet and to inject said fluid from the outlet end of said passageway into at least one of said rear turbine rotors.

According to a third aspect of the invention there is provided a method of suppressing heat in a turbine engine having an engine body disposed within a nacelle, the method comprising the steps of: providing a fluid passageway about a predetermined portion of the rear of said engine body; passing a fluid through said fluid passageway; and as said fluid exits said fluid passageway injecting same at a predetermined rear turbine rotor. It can therefore be seen that there is provided an engine in which it's associated heat signature is substantially minimised or eliminated. It can also be seen that there is provided a method of suppressing the heat signature of a turbine engine which can be applied to new engines or retrofitted to older engines.

Brief Description of Drawings

A preferred embodiment of the invention will now be described, by way of example only, with reference to the accompanying drawings in which: FIG. 1 is a cross-sectional schematic side view of a prior art conventional turbo- fan turbine engine; and

FIG. 2 is a cross-sectional side view of a turbo-fan turbine engine according to the preferred embodiment.

Mode of Carrying Out Invention

Referring to FIG. 1, there is shown a prior art conventional aircraft turbo-fan turbo engine 1. The engine 1 is typically adapted to be mounted to the underside of an aircraft wing. Typically an aircraft uses at least two such engines, each engine 1 being mounted to a corresponding wing.

The engine includes an engine nacelle 2. A turbine engine body 3 is mounted within the nacelle 2 and extends a predetermined distance beyond the rear of the nacelle 2. The turbine engine body 3 includes a plurality of spaced-apart forward turbine rotors 4 disposed at one end of the turbine engine body 3, and a plurality of spaced-apart rear turbine rotors 5 disposed at the other end of the turbine engine body 3. A combustion chamber 6 is disposed intermediate forward turbine rotors 4 and rear turbine rotors 5.

In use, the forward turbine rotors 4 compress air into the combustion chamber 6 where fuel is combusted and exhausted via the rear turbine rotors 5. The combusted gases are exhausted at the rear of the turbine engine body 3 at an outlet end 7. In the embodiment of FIG. 1, a turbo fan 8 is disposed within the front end nacelle 2 and is adapted for increasing the air intake of engine 1. Air is compressed by the turbo fan 8 and part of this air is fed through the turbine engine body 3 and part of this air is forced around the turbine engine body 3 intermediate nacelle 2 and engine body 3

As is commonly known, engines such as those of FIG. 1 provide a high contrast heat signature making them vulnerable to attack. Although the combusted gases base energy through expansion and provide mechanical energy to the rear turbine rotors 5, the temperature effect of the rotors 5 can be of the order of 800⁰C when taking off or landing. Therefore the heat signature is at its greatest when the aircraft is taking off or landing.

Turning now to FIG. 2, there is shown a turbo-fan turbine engine 10 according to a preferred embodiment of the present invention. It is noted that like reference numerals refer to like parts. In FIG. 2, a plurality of forward and rear turbine rotors 4 and 5 are shown. A fluid coil 11 is wrapped around the turbine engine body 3 from just forward of the combustion chamber 6 all the way along the turbine engine body 3 to its outlet end, thereby providing a fluid passageway therealong. It is noted that the turbine engine body 3 is extended beyond that conventionally known, as is the nacelle 2 extended a predetermined distance beyond the outlet end of the turbine body 3.

Fluid inlets 12 are provided on coil 11 at the forward end such that fluid may be passed through coil 11 to its end at the outlet end of the turbine engine body 3. The fluid having traversed the coil 11 is then fed back into the turbine engine body 3 and injected via fluid outlets 13 into the rear turbine rotors 5 and allowed to pass through the three rear-most of the rear turbine rotors 5.

It can therefore be seen that the fluid coil 11 provided around the turbine engine body 3 to remove heat therefrom, and this heated fluid is expelled at outlet 13 to mix with combusted exhaust gases expelled from the outlet end of the turbine engine body 3. That is, the injected heated fluid, in the form of steam when water fluid is used through the coil 11, is fed back into the three rear-most rotors of the rear turbine rotors converting the steam energy into mechanical energy. As the steam expands it mixes with hot exhaust gases thereby reducing the temperature of gases expelled from the outlet end of the turbine engine body 3. This system also advantageously uses the resulting mechanical steam energy provided to the rear turbine rotors 5, to increase engine efficiency and reduce the fuel consumption which results in a cooler exhaust. Static directional blades (not illustrated) disposed intermediate the nacelle 2 and engine body 3 can be used to add rotation to the by-pass air. This improves the mixing of by-pass air and exhaust gases aft of the turbine engine body 3.

It is noted that the fluid outlets 13 can be disposed adjacent any one or more of the rear turbine rotors or even at the outlet end of the turbine engine body 3.

It can further be seen that the cooled exhaust gases are expelled at outlet 7 and mixed with by-pass air flowing from the forward mounted turbo fan between the turbine engine body 3 and the nacelle 2.

It is noted that the rate of flow of fluid through the fluid coil 11 can be electronically controlled by a fluid controller (not illustrated) working in response to signals provided by temperature sensors (not illustrated) disposed at a predetermined location(s) within the turbine engine body 3, at its outlet end, and/or at a rear end of nacelle 2. It is noted that any preferred arrangement can be used in this regard.

The fluid flow through coil 11 may be controlled such that it operates continuously during operation of the engine or to operate selectively at certain times. For instance, fluid flow through fluid coil 11 may be controlled such that it automatically occurs during take-off and landing of the aircraft.

The fluid may be sourced from a fluid reservoir, preferably a water reservoir on the aircraft. However, water to be passed through the fluid coil 11 may also be scavenged from the air during flight by a scavenging unit (not shown). Such a scavenging unit may be disposed on the aircraft.

It can therefore be seen that there is provided an engine in which its associated heat signature is substantially minimised or eliminated. It can also be seen that there is provided a method of suppressing the heat signature of a turbine engine which can be applied to new engines or retrofitted to older engines. It can therefore also be seen that the engine 10 of the preferred embodiments can advantageously use relatively inexpensive materials in the rear turbine rotors 5 and blast (tail) cone 15 since the exhaust plume is cooled, as the fluid coil 11 provides a cooling system. Further, it is noted that the fluid coil 11 may be integrated into the turbine engine body 3 (not illustrated) and tail cone 15. This can be achieved by embedding a fluid coil 11, or by casting the fluid passageway into the engine body 3. Alternatively, the fluid coil 11 may be embedded or the fluid passageway cast into an outer body that is adapted to be fitted around the engine body 3.

The engine 10 of the preferred embodiment can include a rear turbo fan (not illustrated) to induce more efficient mixing of the exhaust plume having been cooled by the coil and the by-pass airflow.

Further, it is noted that the fluid coil 11 can be used in conventional jet engines, namely those without turbo fan(s) 8. In such embodiments (not illustrated) the fluid coil is disposed about the rear of a jet engine body and a fluid outlet is provided adjacent one or more of the rear turbine rotors.

The engine of the above-described embodiment may be referred to as a "bi-turbine engine." This is because the fluid in the form of steam being injected via outlet 13 into rear turbine rotors 5, adds a secondary source of steam energy that is converted into mechanical energy.

The foregoing describes only a preferred embodiment of the present invention and modifications, obvious to those skilled in the art, can be made thereto without departing from the scope of the present invention.

The term "comprising" (and its grammatical variations) as used herein is used in the inclusive sense of "having" or "including" and not in the exclusive sense of "consisting only of.

Claims

1. A turbine engine comprising:

an engine body disposed within a nacelle and extending from adjacent the inlet end of said nacelle to adjacent the outlet end of said nacelle, said engine body housing a combustion chamber disposed intermediate of a plurality of forward turbine rotors and a plurality of rear turbine rotors; and wherein a fluid passageway is substantially coiled about or within said engine body from at least a forward end of the combustion chamber to the outlet end of said engine body, said fluid passageway is adapted to receive fluid at a forward end via a fluid inlet and to inject said fluid from the outlet end of said fluid passageway into at least one of said rear turbine rotors.

2. A turbine engine as claimed in claim 1, wherein said engine body extends a first predetermined distance beyond said rear turbine rotors, and said fluid passageway extends at least partially therealong.

3. A turbine engine as claimed in claim 1, wherein said fluid substantially comprises of water.

4. A turbine engine as claimed in claim 1, wherein said plurality of rear turbine rotors comprises at least three rear turbine rotors.

5. A turbine engine as claimed in claim 4, wherein the outlet of said fluid passageway is disposed adjacent at least one of the three rear most rear turbine rotors.

6. A turbine engine as claimed in claim 1 , wherein a first turbo fan is mounted to a turbine engine shaft to pressurise air entering said engine.

7. A turbine engine as claimed in claim 1, wherein an outer body is disposed substantially about said engine body.

8. A turbine engine as claimed in claim 1, wherein said nacelle extends a second predetermined distance beyond said outlet end of said engine body.

9. A turbine engine as claimed in claim 1 wherein a second turbo fan is mounted adjacent the rear of said nacelle and the rear of said engine body.

10. A turbine engine as claimed in claim 1 , wherein at least a portion of said fluid passageway is disposed within a fluid coil located about said engine body.

11. A turbine engine as claimed in claim 1 , wherein at least a portion of said fluid passageway is cast within said engine body.

12. A turbine engine as claimed in claim 7, wherein at least a portion of said fluid passageway is disposed within said outer body.

13. A cooling system for a turbine engine having a turbine body disposed within a nacelle, said turbine body extending from adjacent the inlet end of said nacelle to adjacent the outlet end of said nacelle, said turbine body having a combustion chamber disposed intermediate of forward and rear turbine rotors, said cooling system comprising a fluid passageway disposed about or within said turbine body from at least a forward end of the combustion chamber to the outlet end of said engine body, said fluid passageway is adapted to receive fluid at a forward end via a fluid inlet and to inject said fluid from the outlet end of said passageway into at least one of said rear turbine rotors.

14. A method of suppressing heat in a turbine engine having an engine body disposed within a nacelle, the method comprising the steps of: providing a fluid passageway about a predetermined portion of the rear of said engine body;

passing a fluid through said fluid passageway; and

as said fluid exits said fluid passageway injecting same at a predetermined rear turbine rotor.