GB2451704A

GB2451704A - Gas turbine engine with compressor formed from a plurality of stacked surfaces

Info

Publication number: GB2451704A
Application number: GB0715667A
Authority: GB
Inventors: Keven Chappell
Original assignee: Individual
Current assignee: Individual
Priority date: 2007-08-10
Filing date: 2007-08-10
Publication date: 2009-02-11
Also published as: WO2009022103A2; WO2009022103A3; GB0715667D0

Abstract

Gas turbine engine 10 comprises a gas inlet 12, a compressor 14, a combustion chamber 16, and a gas outlet rotor 22 at or near the periphery of the combustion chamber 16. The gas inlet 12 is located at or near an axis of rotation AA' of the gas turbine engine 10. The compressor 14 comprises a plurality of mutually spaced stacked surfaces 30,32,34 each having a central aperture 60 operably coupled to the gas inlet 12 and aligned with the axis of rotation AA'. The compressor 14 is of the shear force, fluid friction or molecular drag type. The compressor 14 may comprise a plurality of symmetrical pairs of discs. The combustion chamber 16 may be toroidal and has walls integrally formed from peripheral portions of the stacked surfaces. The gas outlet rotor 20 may comprise a plurality of blades spaced around the periphery of the combustion chamber and each blade may comprise an aerofoil to impart rotation to the turbine engine due to aerodynamic lift.

Description

GAS TURBINE ENGINE

FIELD OF THE INVENTION

The present invention relates to gas turbine engines.

BACKGROUND OF THE INVENTION

Gas turbine engines are rotary engines which extract energy from gas flow. The basic components of a gas turbine engine are a compressor, a combustion chamber and a turbine.

In use, a gas (usually air) is drawn into the compressor where it is pressurised. The compressed air then passes from the compressor into the combustion chamber where fuel is burned so as to heat the air by combustion. The heated pressurised air then expands out of the combustion chamber through the turbine. As the air leaves, it impinges on the turbine rotor blades so as to impart energy to the rotor.

Some of the work extracted by the turbine is used to drive the compressor. In conventional gas turbine engines, there is a casing around the rotor to direct and control the gas flow.

Inefficiencies are introduced in gas turbine engines due to a variety of factors. For example, the casing around the rotor means that rotating element clearances and shaft sealing systems are required, and these elements can introduce inefficiencies. There are also limits imposed by the maximum temperatures which the construction materials can withstand, particularly in the combustion chamber and turbine.

The present invention seeks to provide a gas turbine engine with increased efficiency over those of the prior art. 4' V

SUMMARY OF THE INVENTION

According to a first aspect of the present invention, there is provided a gas turbine engine comprising a gas inlet, a compressor, a combustion chamber and a gas outlet rotor. The gas inlet is located at or near an axis of rotation of the gas turbine engine. The compressor comprises a plurality of mutually spaced stacked surfaces each having a central aperture operably coupled to the gas inlet and aligned with the axis of rotation. The stacked surfaces are arranged to act as Tesla discs as they co-rotate about the axis of rotation (Tesla discs are discussed further in the specific description). The combustion chamber has walls integrally formed from peripheral portions of the stacked surfaces. The gas outlet rotor is located at or near the periphery of the combustion chamber.

Advantageously, the peripheral portions of the stacked surfaces are arranged such that gas is introduced from the compressor into the combustion chamber at a plurality of entrance locations. More advantageously, the plurality of entrance locations are arranged around the edges of the combustion chamber. Optionally, the plurality of entrance locations comprise one or more perforations in the peripheral portions of the stacked surfaces.

In a preferred embodiment, the stacked surfaces comprise a plurality of pairs of symmetrical surfaces including a pair of inner surfaces and a pair of outer surfaces. The inner surfaces are in the centre/middle of the stack and the outer surfaces are at the outside of the stack, thereby defining the top and bottom of the stack.

The outer surfaces are larger than the inner surfaces.

Inside surfaces of the inner surfaces define an inner gas conduit from the gas inlet to the combustion chamber. The peripheral portions of the inner surfaces diverge from one another to form an inner entry portion of the combustion chamber located at the end of the inner gas conduit. Inside surfaces of the outer surfaces partially define a respective pair of outer gas conduits from the gas inlet to the Combustion chamber. Outer entry portions of the combustion chamber are located at the respective ends of the outer gas conduits. The outer gas conduits are formed so as to provide film cooling of parts of the combustion chamber walls in use. In other words, the hot surfaces which form the walls of the outer gas conduits are cooled by the flow of cool gas through the outer gas conduits.

Advantageously, the stacked surfaces further comprise one or more pairs of intermediate surfaces between the inner surfaces and the outer surfaces. The stacked surfaces increase in diameter from the inner surfaces to the outer surfaces. Inside surfaces of each pair of intermediate surfaces partially define a respective pair of intermediate gas conduits from the gas inlet to the combustion chamber.

Intermediate entry portions of the combustion chamber are located at the respective ends of the intermediate gas conduits.

Advantageously, the gas turbine engine further comprises a fuel conduit arranged to inject fuel into the combustion chamber. The fuel may be a liquid or gas.

In one embodiment, the fuel conduit extends substantially along a portion of the axis of rotation and is arranged to direct fuel into the inner gas conduit. In this way, the fuel mixes with gas as it passes between the inner surfaces of the compressor before reaching the combustion chamber. 4 I

In another embodiment, the fuel conduit is arranged to rotate with the turbine engine. Optionally, the fuel conduit extends directly into the combustion chamber from outside one of the outer surfaces. Advantageously, a portion of the fuel conduit is directly adjacent to an outside wall of the combustion chamber.

Advantageously, the gas outlet rotor comprises a plurality of blades spaced around the periphery of the combustion chamber. More advantageously, the blades each comprise an aerofoil shape arranged to impart rotation to the turbine engine due to aerodynamic lift in use.

Alternatively or additionally, the blades are arranged to direct expelled gas at least partially backwards with respect to the direction of rotation of the turbine engine in use. Advantageously, an orientation of the blades is

adjustable.

Preferably, an additional pair of stacked surfaces are provided around the outer surfaces so as to define a pair of cooling gas conduits which provide gas to cool the blades in use.

In one embodiment, the stacked surfaces comprise stacked discs. There is a substantially radial flow of gas in this embodiment (i.e. substantially no axial flow). In alternative embodiments, the stacked surfaces comprise stacked cones or stacked hemispheres. There is a component of axial gas flow in these embodiments as well as a radial component. The axial flow can provide thrust to the engine if required.

Optionally, the blades are arranged to direct expelled gas at least partially axially in use. Again, the axial flow can provide thrust to the engine if required.

Advantageously, the compressor is operably coupled to the gas inlet via at least one initial compression stage for raising the compression ratio of the turbine engine.

In one embodiment, the gas inlet, the compressor, the combustion chamber and the gas outlet rotor are arranged substantially linearly in a radial direction.

Other preferred features of the present invention are set out in the appended claims.

BRIEF DESCRIPTION OF THE DRAWINGS

Embodiments of the present invention will now be described by way of example with reference to the accompanying drawings in which: Figure 1 is a partial side view of a gas turbine engine according to one embodiment of the present invention; Figure 2 is a partial plan view of a rotor of the gas turbine engine shown in Figure 1; Figure 3 is a schematic plan view showing the direction of rotation of the gas turbine engine as compared to the direction of expelled gases; Figure 4 is a partial side view of an alternative fuel delivery system for the engine for Figures 1 to 3; Figure 5 is a cross-sectional side view showing how the fuel is delivered into the engine combustion chamber using the fuel delivery system of Figure 4; Figure 6 is a cross-sectional side view showing an alternative way of delivering fuel into the engine combustion chamber using the fuel delivery system of Figure 4; Figure 7 is a cross-sectional side view showing an initial compression stage coupled to the top of the engine shown in Figure 5; Figure 8 is a cross-sectional side view showing initial compression stages coupled to both the top and bottom of the engine shown in Figure 5; Figure 9 is a partial plan view of a gas outlet rotor of the engine; arid Figure 10 is a plan view of a blade of the gas outlet rotor showing the cooling airflow.

DETAILED DESCRIPTION OF A PREFERRED EMBODIMENT

Figure 1 shows a partial side view of a gas turbine engine 10 according to one embodiment of the present invention. The entire engine 10 rotates about an axis of rotation AA' in use. The rotating engine 10 comprises a gas inlet 12, a compressor 14, an annular combustion chamber 16, a gas outlet rotor 20, and a shaft 22. For simplicity, in Figure 1, the left hand side of the engine 10 is shown in full, but part of the right hand side (i.e. the combustion chamber and rotor) is omitted.

The shaft 22 is aligned with the axis of rotation AA' of the engine 10. Bearings 24 are used to mount the rotating engine 10 to non-rotating parts. A non-rotating fuel injection tube 18 is also shown in Figure 1.

The compressor 14 comprises three pairs of stacked co-axial discs. There is an inner pair of discs 30, an outer pair of discs 34, and an intermediate pair of discs 32 disposed between the inner discs 30 and the outer discs 34.

Each pair of discs is symmetrical with respect to a horizontal plane passing between the inner discs 30, as shown by line BE' in Figure 1. Each pair of discs is also symmetrical around the axis of rotation AA'. The inner discs 30 have a smaller diameter than the intermediate discs 32, which in turn have a smaller diameter than the outer discs 34.

Each adjacent pair of compressor discs defines a radial gas conduit from the gas inlet 12 to the combustion chamber 16. An inner gas conduit 40 is defined between inner surfaces of the inner discs 30. Two intermediate gas conduits 42 are defined between outer surfaces of the inner discs 30 and inner surfaces of the intermediate discs 32.

Two outer gas conduits 44 are defined between outer surfaces of the intermediate discs 32 and inner surfaces of the outer discs 43. It will be appreciated that a different number of pairs of discs, and therefore a different number of gas conduits, could also be used within the scope of the invention. In Figure 1, the spacing between each pair of adjacent discs 30, 32 and 34 (i.e. the width of each gas conduit 40, 42 and 44) is approximately the same. However, different spacings could also be used so long as the spacing allows the discs 30, 32 and 34 to act as Tesla discs as discussed further below.

Each disc of the compressor 14 comprises three regions: a central aperture 60, an annular compressor region 62 and an annular peripheral region 64. The disc central apertures are aligned with the axis of rotation AA'. Furthermore, the disc central apertures 60 define the gas inlet 12 of the engine 10. Each disc has a number of holes located at evenly spaced intervals around the central aperture 60 in the annular compressor region 62. The holes in each disc are aligned axially and are adapted to receive bolts 26 for rigidly attaching the discs to one another and to the shaft 22. The bolts 26 are received in an axial direction. In the annular compressor region 62, the discs 30, 32 and 34 are planar and parallel to one another. Thus, the gas conduits 40, 42 and 44 are linear in this region as shown in Figure 1.

The peripheral regions 64 of the discs are shaped so as to define the annular combustion chamber 16. Thus, walls of the annular combustion chamber 16 are integrally formed from peripheral portions 64 of the compressor discs so that the compressor 14 and the combustion chamber 16 rotate together.

The inner surfaces of the walls of the annular combustion chamber 16 are highly polished.

The inner discs 30 of the compressor 14 are dimensioned to set the initial compressor outlet pressure required by the engine 10. The peripheral portions 64 of the inner discs 30 diverge from one another (i.e. curve away from the line PB') so that they together form a divergent region which acts as an inner annular entry portion 50 of the combustion chamber 16 at the end of the inner gas conduit 40. The peripheral portions 64 of the intermediate discs 32 diverge from one another so as to maintain a separation between the inner discs 30 and the intermediate discs 32.

However, since the intermediate discs have a larger diameter than the inner discs 32, two intermediate annular entry portions 52 of the combustion chamber 16 are formed at the ends of the intermediate gas conduits 42. The peripheral portions 64 of the outer discs 34 initially diverge from one another so as to maintain a separation between the intermediate discs 32 and the outer discs 34. However, once there is no longer any overlap between the intermediate and outer discs 32 and 34 due to the larger diameter of the outer discs 34, the peripheral portions 64 of the outer discs 34 start to curve towards one another (i.e. curve towards the line PB') as shown in Figure 1. The outer discs 34 thereby enclose the toroidal combustion chamber 16. Thus a peripheral portion of the annular combustion chamber 16 is formed. In addition, two outer annular entry portions 54 of the combustion chamber 16 are formed at the ends of the outer gas conduits 44. Thus, in this embodiment, there are five different and distinct annular entry regions into the annular combustion chamber 16 from the compressor 14.

Also shown in Figure 1 are an outer pair of annular members 36. The annular members 36 are stacked coaxially around the outer compressor dscs 34. In a radial direction, the annular members 36 extend from the inner annular entry portion 50 of the combustion chamber 16 to the rotor 20. Therefore, the annular members 36 have a larger diameter than the outer discs 34. In the radial vicinity of the inner annular entry portion 50 of the combustion chamber 16, the annular members 36 are relatively far apart from one another and extend outwards from the axis of rotation AA' in a substantially radial direction so as to define gas entry regions 28. The annular members 36 then converge towards the rotor 20 to define cooling gas conduits 46 between inner surfaces of the annular members 36 and outer surfaces of the outer discs 34.

The turbine part of the gas turbine engine 10 is formed by the rotor 20, which is shown in more detail in Figure 2.

The rotor 20 is rigidly attached to the combustion chamber 16 so that the rotor 20 rotates together with the rotating compressor 14 and combustion chamber 16. The rotor 20 is annular and comprises a number of blades 70 evenly spaced around its circumference. Each adjacent pair of blades 70 defines a gas outlet therebetween in the form of a nozzle 72. Each blade has a first blade portion 74 in the form of an aerofoi]. which extends substantially radially outwards from the annular combustion chamber 16. Each blade also has -10 -a second blade portion 76 integrally formed with the first blade portion 74. The second blade portion 76 extends substantially circumferentially with respect to the rotor 20 and is used to direct the expelled gas in a substantially rearward direction with respect to the direction of rotation X, as shown in Figure 3. In an alternative embodiment, the first and second blade portions 74 and 76 may be separate so as to form two sets of rotor blades.

In use, the compressor discs 30, 32 and 34 co-rotate about the axis of rotation AA'. As mentioned above, the discs are sufficiently closed spaced together that they act as Tesla discs (otherwise known as friction discs) which use the boundary layer effect. In particular, as the compressor 14 rotates, gas from the gas inlet 12 is observed to be drawn radially into the gas conduits 40, 42 and 44 between the discs 30, 32 and 34. As the gas enters the compressor 14, friction between the rotating discs 30, 32 and 34 and the gas causes a boundary layer of gas near the surface of each disc 30, 32 and 34 to rotate with that disc. The spacing between each pair of adjacent discs 30, 32 and 34 (i.e. the width of each gas conduit 40, 42, 44) is small enough that adjacent boundary layers within each gas conduit 40, 42 and 44 interact. In this way, all of the gas within each gas conduit 40, 42 and 44 is made to rotate with the compressor 14. As the gas rotates, its angular momentum increases such that it moves away from the central aperture through the annular compressor region 62 and towards the annular peripheral region 64 of the discs 30, 32 and 34, gaining kinetic energy as it goes. Eventually, the gas is expelled outwards from the annular compressor region 62 into the combustion chamber 16 in the annular peripheral region -11 - 64. On entering the combustion chamber 16 at the annular entry portions 50, 52 and 54, the gas is decelerated so as to convert the kinetic energy it gained to provide a useful rise in pressure energy (i.e. there is a resultant increase in enthalpy). This configuration means that each gas conduit 40, 42 and 44 provides an associated high-pressure gas source in the combustion chamber 16. By small changes in geometry, it is possible to control the velocity, temperature and pressure of the gas in any one of the gas conduits 40, 42 and 44 independently of its neighbours.

While all the discs 30, 32 and 34 together form the compressor 14, the inner discs 30 additionally perform a fuel mixing and delivery function. In particular, fuel is delivered radially to the inner gas conduit 40 by means of the central fuel injection tube 18. Together with the gas, the injected fuel is transferred along the inner gas conduit to provide a pre-mixed gas/fuel mixture at the inner annular entry portion 50 of the combustion chamber 16. The compressed gas/fuel mixture is ignited in the combustion chamber by ignition means (not shown) as is known in the art. Combustion then occurs in the combustion chamber 16 such that the pressure of the gas in the combustion chamber 16 increases still further. The gas used must be combustible. Preferably the gas is air.

To complete the combustion process, secondary gas is supplied to the intermediate annular entry portions 52 of the combustion chamber from the intermediate gas conduits 42. The secondary gas is more compressed than the gas entering the combustion chamber 16 via the inner annular entry portion 50 since the intermediate gas conduits 42 are longer than the inner gas conduit 40. Tertiary gas is supplied to the outer annular entry portions 54 of the -12 -combustion chamber from the outer gas conduits 44. The tertiary gas is more compressed than the secondary gas since the outer gas conduits 44 are longer than the intermediate gas conduits 42.

The toroida]. form of the combustion chamber 16 tends to centralise the combustion gases. The ability of the combustion chamber 16 to control the combustion gases is further enhanced by the highly polished internal surfaces which reflect and focus the radiant thermal energy during combustion. The combustion chamber 16 has no circumferential restrictions to gas flow. Therefore, gas tends to move in a diagonal direction (i.e. partially radially outwards, and partially circumferentially) through the combustion chamber 16. This gives a longer dwell time of the gas in the combustion chamber 16 than in a non-integral, non-rotating combustion chamber of a similar size.

This allows for more efficient burning of the fuel and more efficient combustion without requiring a larger combustion chamber.

In the arrangement of Figure 1, the staged gas delivery into the combustion chamber (i.e. the gas delivery into the combustion chamber via the five distinct annular entry portions 50, 52 and 54) ensures that high gas pressures and velocities can be introduced at any point in the combustion chamber 16. In addition, the staged gas delivery promotes gas mixing by the creation of localised turbulence. This turbulence enhances the flame stability and mixing of combusting products to promote cleaner burning.

The staged gas delivery also promotes film cooling of the walls of the combustion chamber 16. In particular, the walls of the combustion chamber 16 adjacent to the inner annular entry portion 50 are cooled by gas from the inner -13 -gas conduit 40, the walls of the combustion chamber 16 adjacent to the intermediate annular entry portions 52 are cooled by gas from the intermediate gas conduits 42, and the walls of the combustion chamber 16 adjacent to the outer annular entry portions 54 are cooled by gas from the outer gas conduits 44. This film cooling of the combustion chamber walls is integral to the design of this gas turbine engine 10.

Combustion temperatures and pressures can be somewhat higher in this engine 10 due to the unbroken integral film cooling of the combustion chamber walls and the automatic centralisation of the hottest combustion gases, as described above. These elevated temperatures are not possible in small conventional gas turbine engines due to the limits in place to protect the non-cooled metal structures and facets in the gas path. Higher operating temperatures and pressures, and the reduction in parasitic losses in efficiency enables the gas turbine engine described herein to operate at a higher overall efficiency than a similarly sized, conventionally arranged gas turbine engine.

Rotation of the engine 10 and thus the operation of the compressor 14 is achieved by expelling the high pressure at high speed through the rotor 20. The outer discs 34 converge to direct and accelerate the gaseous products of combustion towards the peripherally mounted rotor 20. The annular members 36 direct gas along the cooling gas conduits 46 from outside the engine 10 into the hot gas stream from the combustion chamber 16 to the rotor 20. The cool gas from the cooling gas conduits 46 cools the hot gas stream to a level that the blades 70 can tolerate.

As the expelled gas leaves the combustion chamber 16 through the rotor 20, the aerofoil-shaped first blade -14 -portion 74 of the rotor 20 acts to impart a rotation to the engine 10 due to aerodynamic lift. The expelled gas then imparts a reaction force on the second blade portions on the engine 10 which helps to drive the rotation of the engine 10. Thus, both the first and second blade portions 74 and 76 of the rotor act to rotate the engine in the direction of rotation X around the axis of rotation AA'.

A useful power output maybe taken from the shaft 22 to drive a generator or other through a gearbox.

Alternatively, a fan or multi-bladed propulsor unit maybe mounted directly to the engine 10 to provide thrust, a variation of which utilises the exhaust gases to drive an unconnected second turbine attached to a propeller.

As an alternative to the fuel injection tube 18, another embodiment of a fuel delivery system is shown in the partial cross-sectional view of Figure 4. In this embodiment, the fuel is first used to lubricate and cool the bearings 24 (see the curved arrows in Figure 24) before passing inwards into a radial fuel conduit 80 in the shaft 22. The fuel then passes into an axial fuel conduit 82 which leads from the radial fuel conduit 80 through the shaft 22 towards one of the outer discs 34b of the compressor 14. The fuel then passes outwards into a radial fuel conduit 84 adjacent the outer disc 34b and into fuel injection tubes 19.

The fuel injection tubes 19 are thin bore pipes that deliver the atomised fuel from outside one of the outer surfaces 34b directly into the inner annular entry portion of the combustion chamber 16, as shown in Figure 5. As the fuel moves along a portion 19a of the respective fuel injection tube 19 within the combustion chamber 16, the fuel is momentarily heated to a high temperature which allows it -15 -to be burnt more readily, thus lowering the emissions of the engine 10. In the embodiment of Figure 5, the outer pair of annular members 36 are formed as an additional pair of outer discs which act to draw cool gas from the gas inlet 12 into the hot gas stream from the combustion chamber 16 to the rotor 20. As in the previous embodiment, the cool gas from the cooling gas conduits 46 cools the hot gas stream to a level that the blades 70 can tolerate. In addition, the cool gas within the cooling gas conduits 46 acts to cool the outside walls of the combustion chamber 16 which again enables higher combustion temperatures to be used within the combustion chamber 16 without impacting negatively on the integrity of the combustion chamber walls.

A further alternative fuel injection embodiment is shown in Figure 6. In this embodiment, the fuel injection tubes 19 each comprise a first portion 19b, a second portion 19c, a third portion 19d, a fourth portion 19e and a fifth portion 19f. The first portion 19b extends radially outwards from the axis of rotation AA' and is directly adjacent to an outside wall 36b of the combustion chamber.

The second portion l9c extends axially away from a peripheral portion of the combustion chamber 16 and bends round to form the radially inwardly extending third portion 19d which is spaced from the first portion 19b and the combustion chamber 16. The third portion 19d then curves and merges into the axially extending fourth portion l9e which is directed towards the combustion chamber 16. As the fourth portion passes through the discs 30, 32, 34 and' 36, it merges into the fifth portion 19f which is located within the combustion chamber 16. As for the embodiment of Figure 5, as the fuel moves along the fifth portion 19f it is momentarily heated to a high temperature which allows it -16 -to be burnt more readily, thereby lowering the emissions of the engine 10.

Another embodiment of a gas turbine engine 10 is shown in the partial cross-sectional side view of Figure 7. In this embodiment, there is an initial compression stage 90 coupled to the top of the engine 10. The initial compression stage 90 comprises an initial compressor 92, a gas return conduit 96, and a gas transference chamber 94 therebetween. The initial compressor comprises a plurality of mutually spaced stacked discs rigidly connected to the rotating engine 10. Each stacked disc of the initial compressor 92 has a central aperture coupled to the gas inlet 12. The stacked discs of the initial compressor 92 act as Tesla discs as they co-rotate, thereby drawing gas into the initial compressor 92 as shown by arrow A in Figure 7. As for the compressor 14, the gas within the initial compressor 92 starts to rotate with the discs due to the boundary layer effect and is forced radially outwards towards the annular gas transference chamber 94 located at the periphery of the initial compression stage 90. As the gas enters the gas transference chamber 94, kinetic energy is transferred to pressure energy so that there is a resultant increase in enthalpy. Due to this increased gas pressure, the gas is then forced inwardly from the gas transference chamber 94 into the annular gas return conduit 96 which leads back towards the axis of rotation AA'. As shown by arrow B in Figure 7, the already partially compressed gas is then drawn from the gas return conduit 96 into the compressor 14. In this way, the gas is more compressed on reaching the compression chamber 16 in the embodiment of Figure 7 than it is in the embodiment of -17 -Figures 5 or 6 which do not include an initial compression stage 90.

A further embodiment of a gas turbine engine 10 is shown in the partial cross-sectional side view of Figure 8.

In this case, there is an initial compression stage 90 coupled to the top of the engine 10, similarly to described above with reference to Figure 7. There is also an initial compression stage 98 coupled to the bottom of the engine 10.

The two initial compression stages 90 and 98 are substantially symmetrically disposed about the combustion chamber 16.

It is envisaged that further initial compression stages could be used at the top or bottom of the engine 10.

In the embodiments of Figures 5 to 8, the peripheral portions 64 of the discs 30 and 32 comprise a number of perforations 100 so as to provide an increased number of entry locations from the compressor 14 into the combustion chamber 16. This increases the staged nature of the gas delivery into the combustion chamber 16 to further enhance the mixing and cooling properties described above.

Figure 9 shows a partial plan view of the gas outlet rotor 20 of the engine 10. Three blades 70a, 70b and 70c are shown in Figure 9, however blades are actually located around the entire periphery of the rotor 20. Each blade 70 has an associated pivot point 71 by which it is attached to the rotor 20. The orientation of each blade 70 with respect to its respective pivot point 71 can be altered. For the purposes of illustration, each of the blades 70 in Figure 9 has a different orientation with respect to the engine 10.

However, in normal use, it is intended that each of the blades 70 be oriented at a similar angle with respect to the engine 10. The orientation shown for the blade 70a expels -18 -output gas more tangentially than the orientation shown for the blade 70b. Similarly, the orientation shown for the blade 70b expels output gas more tangentially than the orientation shown for the blade 7Cc. In this way, by altering the orientation of the blades 70 with respect to the engine 10, it is possible to provide a variable RPM engine 10. The variable orientation blades 70 enable optimisation of the gas exit velocity, and therefore the derivable power, at various rotational speeds.

Figure 10 is a plan view of a blade 70 of the gas outlet rotor 20. The blade 70 comprises blade cooling passages 73. In use, gas from the cooling gas conduits 46 passes through the blade cooling passages 73 so as to cool the blade 70.

In alternative embodiments, the engine 10 comprises stacked conical surfaces or hemispheres rather than the stacked discs 30, 32 and 34 of the previous embodiments.

The use of nested cones or hemispheres of sheet metal enables a more axial flow of the gases, which promotes a rearward facing exhaust as opposed to the tangential one of the previous embodiments.

The described embodiments of the present invention provide a small, reliable, lightweight, gas turbine engine of the reaction type. The engine described herein benefits from low production costs due to a pressed sheet construction which removes the requirement for expensive castings. Rotating as one complete unit, the engine minimises efficiency losses by removing the requirement for rotating element clearances and shaft sealing systems. The engine is designed to run efficiently on all currently available liquid or gaseous fuels. The engine is -19 -mechanically simple so is not complex and requires few ancillaries and can be considered almost disposable in its basic form. In its most advanced state, however, with the addition of modular components, it can achieve high fuel efficiency returns comparable to, or in excess of, existing gas turbine engines.

Although preferred embodiments of the invention have been described, it is to be understood that these are by way of example only and that various modifications may be contemplated.

Claims

-20 -CLAIMS: 1. A gas turbine engine comprising: a gas inlet located at or near an axis of rotation of the gas turbine engine; a compressor comprising a plurality of mutually spaced stacked surfaces each having a central aperture operably coupled to the gas inlet and aligned with the axis of rotation, the stacked surfaces being arranged to act as Tesla discs as they co-rotate about the axis of rotation; a combustion chamber having walls integrally formed from peripheral portions of the stacked surfaces; and a gas outlet rotor at or near the periphery of the combustion chamber.
2. The gas turbine engine of claim 1 wherein the peripheral portions of the stacked surfaces are arranged such that gas is introduced from the compressor into the combustion chamber at a plurality of entrance locations.
3. The gas turbine engine of claim 2 wherein the plurality of entrance locations are arranged around the edges of the combustion chamber.
4. The gas turbine engine of any preceding claim wherein the combustion chamber is substantially toroidal.
5. The gas turbine engine of any preceding claim wherein the stacked surfaces comprise a plurality of pairs of symmetrical surfaces including a pair of inner surfaces and a pair of outer surfaces, the outer surfaces being larger than the inner surfaces.

-21 -
6. The gas turbine engine of claim 5 wherein inside surfaces of the inner surfaces define an inner gas conduit from the gas inlet to the combustion chamber.
7. The gas turbine engine of claim 6 wherein the peripheral portions of the inner surfaces diverge from one another to form an inner entry portion of the combustion chamber located at the end of the inner gas conduit.
8. The gas turbine engine of any of claims 5 to 7 wherein inside surfaces of the outer surfaces partially define a respective pair of outer gas conduits from the gas inlet to the combustion chamber.
9. The gas turbine engine of claim 8 wherein outer entry portions of the combustion chamber are located at the respective ends of the outer gas conduits.
10. The gas turbine engine of claim 8 or claim 9 wherein the outer gas conduits are formed so as to provide film cooling of parts of the combustion chamber walls in use.
11. The gas turbine engine of any of claims 5 to 10 wherein the stacked surfaces further comprise one or more pairs of intermediate surfaces between the inner surfaces and the outer surfaces, the stacked surfaces increasing in diameter from the inner surfaces to the outer surfaces.
12. The gas turbine engine of claim 11 wherein inside surfaces of each pair of intermediate surfaces partially -22 -define a respective pair of intermediate gas conduits from the gas inlet to the combustion chamber.
13. The gas turbine engine of claim 12 wherein intermediate entry portions of the combustion chamber are located at the respective ends of the intermediate gas conduits.
14. The gas turbine engine of any preceding claim further comprising a fuel conduit arranged to inject fuel into the combustion chamber.
15. The gas turbine engine of claim 14 wherein the fuel conduit extends substantially along a portion of the axis of rotation and is arranged to direct fuel into the inner gas conduit.
16. The gas turbine engine of claim 14 wherein the fuel conduit is arranged to rotate with the turbine engine.
17. The gas turbine engine of claim 16 wherein the fuel conduit extends directly into the combustion chamber from outside one of the outer surfaces.
18. The gas turbine engine of claim 17 wherein a portion of the fuel conduit is directly adjacent to an outside wall of the combustion chamber.
19. The gas turbine engine of any preceding claim wherein the gas outlet rotor comprises a plurality of blades spaced around the periphery of the combustion chamber.

-23 -
20. The gas turbine engine of claim 19 wherein the blades each comprise an aerofoil shape arranged to impart rotation to the turbine engine due to aerodynamic lift in use.
21. The gas turbine engine of claim 19 or claim 20 wherein the blades are arranged to direct expelled gas at least partially backwards with respect to the direction of rotation of the turbine engine in use.
22. The gas turbine engine of any of claims 19 to 21 wherein an orientation of the blades is adjustable.
23. The gas turbine engine of any of claims 19 to 22 when dependent on claim 5 wherein an additional pair of stacked surfaces are provided around the outer surfaces so as to define a pair of cooling gas conduits which provide gas to cool the blades in use.
24. The gas turbine engine of any of claims 1 to 23 wherein the stacked surfaces comprise stacked discs.
25. The gas turbine engine of any of claims 1 to 23 wherein the stacked surfaces comprise stacked cones, each cone having a central aperture at its nose.
26. The gas turbine engine of any of claims 1 to 23 wherein the stacked surfaces comprise stacked hemispheres, each hemisphere having a central aperture.
27. The gas turbine engine of claim 25 or claim 26 when dependent on claim 19 wherein the blades are arranged to direct expelled gas at least partially axially in use.

-24 -
28. The gas turbine engine of any preceding claim wherein the compressor is operably coupled to the gas inlet via at least one initial compression stage for raising the compression ratio of the turbine engine.
29. The gas turbine engine of any preceding claim wherein the gas inlet, the compressor, the combustion chamber and the gas outlet rotor are arranged substantially linearly in a radial direction.
30. The gas turbine engine of any preceding claim when dependent on claim 2 wherein the plurality of entrance locations comprise one or more perforations in the peripheral portions of the stacked surfaces.
31. A gas turbine engine substantially as herein described with reference to Figures 1 to 10 of the accompanying drawings.

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