Improvements in or relating to radial turbines

Abstract

A rotor (20) for a radial turbine (1), comprising: a hub (22); and a plurality of turbine blades (24), extending from or affixed to at least part of the hub (22), providing a plurality of combustion gas passageways (28) between respective turbine blades (24) along which combustion gases may pass so as to rotate the rotor (20). The one or more of the plurality of turbine blades (24) comprise at least part of a cooling passageway (31), the combustion gas passageways (28) and cooling passageway (31) being separate passageways configured to provide movement of combustion and cooling gases. The rotor further comprises a shroud (23), the plurality of turbine blades (24) extending between at least part of the hub (22) and at least part of the shroud (23), providing a plurality of enclosed combustion gas passageways through which combustion gases may pass so as to rotate the rotor (20).

Description

[0001] Improvements in or Relating to Radial Turbines

[0002] The present invention relates to a rotor for a radial turbine and associated method for cooling a rotor of a radial turbine. In particular, the invention relates to a shrouded rotor for a radial turbine which provides improved rotor cooling, and an associated method for cooling. The present invention also relates to a gas turbine apparatus which includes the rotor, and an associated method for cooling a gas turbine apparatus.

[0003] Micro gas turbines (MGT) are known in the art. It is well known by those skilled in the art that there is a relationship between engine temperature and efficiency of a turbine apparatus, such that a higher temperature ratio provides higher efficiency. In particular, with respect to a jet engine, its efficiency is equivalent to the turbine entry temperature (TET).

[0004] MGTs generally use uncooled radial turbines and are, therefore, inefficient owing to low TETs, being, say, around 800°C to 1000°C. MGT manufacturers seek to compensate for this known problem by using exhaust heat recouperation to reduce engine workload and improve efficiency, this being the only viable option since small-sized cooled turbines do not presently justify the cost implications of improved cooling cycles.

[0005] There are various known methods for providing turbine cooling but all use a parallel flow of bleed air from the compressor. The bleed air from the compressor is, at most, no more than 10% of the total engine flow, which limits how much cooling these known methods may provide. In addition, bleed air from the compressor is already warmer than ambient air, as it has passed through the compressor. For example, internal convection cooling passes compressor bleed air through internal passages of a blade of a compressor of the turbine. Internal impingement cooling provides a variation of convection cooling, in which high-velocity air is directed at inner surfaces of the turbine. External film cooling provides pumping of cooled air through small holes or slots in a film coating an outside of the blade, to provide a thin film of cool air which seeks to isolate the combustion gases from the surface of the blade. Further, external transpiration cooling, which is similar to film cooling but uses a porous medium rather than machined holes or slots, provides the blade with more uniform cooling. All seek to allow the blade to work in temperatures in excess of the melting point of the material from which the blade has been manufactured.

[0006] MGT turbines have challenging geometry to effectively cool, which tends to make the cost of doing so prohibitive. Whilst it is not intended to be bound by theory, combustion temperatures between a large turbofan and an MGT are not linear compared to the amount of air flow. For example, a large turbofan might combust at around 2000°C but ingest more than 750kg / s of air and, so, a small proportion of that is much larger relative to the temperature increase than with an MGT, which combusts at around 1500°C but which only ingests 1 kg / s of air. Accordingly, a compressor of an MGT would actually have to work much harder to be able to achieve the same cooling effect (if it was geometrically possible).

[0007] In view of the above, and although for large scale turbines transpiration cooling is often the best option, present cooling methods do not provide adequate cooling and, therefore, a different mechanism has been sought to improve rotor blade cooling.

[0008] The following two paragraphs provide disclosures of prior art micro gas turbines, MGTs, with reference to Figures 1 to 3. As the following refers to known technology, it will only be briefly described. A first variant is shown in Figure 1a, being identified generally by reference 101 , and a second variant is shown in Figures 1 b, 2 and 3, being identified generally by reference 10T.

[0009] In Figure 1a, the turbine apparatus 101 includes a compressor 102, a combustor 103, a turbine 104 and a generator 105. As will be known by those skilled in the art, an air inlet 106 provides ambient air to the compressor 102, which compresses the ambient air and supplies the compressed air through a combustor inlet 107 to the combustor 103. The combustor 103 includes a fuel inlet 108 and, in the combustor 103, the fuel is burnt in the presence of compressed air supplied by the combustor inlet 107, creating combustion gases. The combustion gases are supplied to the turbine 104 through a turbine inlet 109 and work upon the turbine 104 creating a mechanical, rotational output 110. The rotational output 110 has two forms, being: 1 ) a rotational input 110a to the generator 105, for generating electricity; and 2) a rotational input 110b to the compressor 102, for rotating the compressor (and compressing further ambient air). The turbine 104 also includes an exhaust output 111 for exhausting combustion gases, which may or may not be fed to a heat-exchanger for additional cooling of the exhaust combustion gases I heating of air between the compressor 102 and combustor 103. This first variant is an MGT with uncooled turbine.

[0010] In Figures 1 b, 2 and 3, a second variant is shown, which is an MGT with cooled turbine. The turbine apparatus 10T has many common features with the turbine apparatus 101 , which common features will not be discussed in detail - only the main differences will be described further. The apparatus 10T includes a cooled turbine 104’, which has a rotor 112 - having a plurality of rotor blades 112a - that are each cooled through supply of bleed air 113, from the compressor 102. The turbine 104’ has a hollow shaft 114, extending from the compressor 102 to the turbine 104’, and the turbine 104’ is additionally hollowed providing one or more blade cooling channels 115 and central cooling channels 116, through which the bleed air 113 may pass, so as to cool the rotor 112. The channels 115 and 116 exit the rotor 112 through various outlets 117, which feed directly into the exhaust output 111 - there is no separate exhaust for the bleed air 113. It is important to note that cooling air provided to the rotor 112 is bleed air 113, supplied from the compressor only and is, therefore, not ambient air. In addition, the bleed air travels in the same direction as the combustion gases through the turbine, being from right to left, as shown in the figures.

[0011] Although cooling of rotors for radial turbines is already generally known, those present apparatus suffer from various disadvantages, which are as follows. One must use high-grade materials, such as a nickel-chromium-based supper alloy, in addition to a form of internal cooling to achieve the temperature performance required. One must use bleed air from the compressor to cool the turbine, which results in lost compressor work and pumping losses. Owing to the source of the cooling air being bleed air from the compressor, and that air being, thereby, compressed, this provides a reduction in the AT between the TET and the cooling (but compressed) air (1200°C 1200°C). Further, bleed air (cooling air) and exhaust combustion gases merge at the outlet I exhaust, cooling down the exhaust gases, which reduces heat recouperation efficiency, again, through a reduced AT.

[0012] In prior art MGTs, coolant flow to the rotor is approximately 2% to 7% of the core engine flow, significantly restricting the amount of coolant flow.

[0013] The present invention is, therefore, aimed at providing an improved turbine rotor and turbine cooling system which is applicable to various forms of radial turbine. According to a first aspect, the present invention provides a rotor for a radial turbine, comprising: a hub; and a plurality of turbine blades, extending from or affixed to at least part of the hub, providing a plurality of combustion gas passageways between respective turbine blades along which combustion gases may pass so as to rotate the rotor, wherein one or more of the plurality of turbine blades comprise at least part of a cooling passageway, the combustion gas passageways and cooling passageway being separate passageways configured to provide separate movement of combustion and ambient gases, wherein the rotor further comprises a shroud, the plurality of turbine blades extending between at least part of the hub and at least part of the shroud, providing a plurality of enclosed combustion gas passageways through which combustion gases may pass so as to rotate the rotor.

[0014] Preferably, the cooling passageway comprises an internal shape of the turbine blade being an approximate inversion to an external shape of the turbine blade at or through the same cross-section.

[0015] Preferably, the separate passageways are configured to provide movement of combustion and ambient gases in opposite directions acting upon at least part of the turbine blade. Preferably, the separate passageways are configured to provide movement of combustion and ambient gases in opposite directions along at least part of the turbine blade. Most preferably, the separate passageways are configured to provide movement of combustion and ambient gases in opposite directions along a substantial portion of the turbine blade I the whole blade.

[0016] Preferably, each turbine blade comprises a cooling passageway.

[0017] Preferably, the separate passageways are configured to prevent mixing of combustion and ambient gases.

[0018] Preferably, an external surface of the shroud comprises a cooling air inlet, or outlet, operatively connected to an internal channel of the one or more blades, which is operatively connected to a cooling air outlet, or inlet, respectively, provided on an external surface of the hub and / or a further surface or part of the shroud.

[0019] Preferably, the shroud is a frusto-conical shell which, together with the hub and plurality of turbine blades, provide enclosed combustion gas passageways having radial inlets, and axial outlets. Preferably, ‘radial’ and ‘axial’ is determined with respect to a shaft upon which the rotor rotates.

[0020] Preferably, the cooling passageway comprises: an inlet or outlet on an external surface of the shroud; and / or an outlet or inlet on an external surface of the hub, or further surface or part of the shroud.

[0021] Preferably: the inlet or outlet to the cooling passageway is formed on a / the frusto-conical shell surface of the shroud; and / or the outlet or inlet from the cooling passageway is formed on a radial surface of the hub.

[0022] Preferably, the cooling passageway comprises a hollowed portion of the rotor blade. Preferably, the cooling passageway comprises an aperture or slot (preferably elongate), extending axially along the shroud so as to provide a hollow region of the blade.

[0023] Preferably, the cooling passageway comprises an aperture or slot (preferably elongate), located radially on a disc portion of the hub and substantially in-line with at least part the blade.

[0024] Preferably, the inlet to and outlet from the cooling passageway are in fluid communication with each other, including the cooling passageway itself.

[0025] Preferably, the axial aperture or slot, and the radial aperture or slot are in fluid communication with each other, including the cooling passageway.

[0026] Preferably, the turbine blades comprise radially arranged blades of substantially arcuate, sinuous, or straight profile, or combinations thereof.

[0027] Preferably, the enclosed combustion gas passageways comprise arcuate, sinuous, straight and / or partially conical surfaces, or combinations thereof.

[0028] Preferably, an internal cooling surface area of one or more of the plurality of turbine blades is about 50% to about 85% of an external surface area of the same turbine blade exposed to heat, preferably, about 60% to about 75%.

[0029] Preferably, a cross-sectional surface area of the cooling passageway of the turbine blade is about 40% to about 60% of the total cross-sectional surface area of the turbine blade at the same cross-section.

[0030] Preferably, the invention provides heat-exchange but no mixing of ambient and combustion gases.

[0031] Preferably, the turbine rotor is mountable to a shaft for connecting with a compressor and generator, or other mechanical output means.

[0032] Preferably, the rotor is configured to receive (ambient) cooling air through a / the rotor blade travelling in a direction at least partially opposed to the direction of travel of combustion gas along or through combustion gas passageways formed between respective rotor blades.

[0033] Preferably, the rotor comprises: a first passageway, for combustion gas to move from a first end of the rotor to a second end of the rotor, causing rotation of the rotor; and a second passageway, for (ambient) air to move at least partially in an counter-flow condition between the second and first ends of the rotor, causing cooling of at least part of the rotor. Further preferably, the first and second passageways are separate passageways.

[0034] Preferably, the shroud and the hub are separate parts and manufactured separately, and then assembled. Alternatively, and also preferably, the shroud and hub may be unitary and formed in a single manufacturing step.

[0035] According to a second aspect, the invention provides a method for cooling a rotor of a radial turbine, comprising: passing combustion gases along passageways formed between respective turbine blades of the rotor, so as to rotate the rotor; passing cooling air through a cooling passageway formed in one or more turbine blades of the rotor, so as to cool at least part of the turbine blade, wherein combustion gas passageways and the cooling passageway are separate passageways providing separate movement of combustion and ambient gases wherein the rotor comprises a shroud, the turbine blades extending between at least part of a hub of the rotor and at least part of the shroud, providing a plurality of enclosed combustion gas passageways along which combustion gases pass so as to rotate the rotor. Preferably, the separate passageways provide movement of combustion and ambient gases in opposite directions along at least part of the turbine blade.

[0036] Preferably, combustion gas passageways and the cooling passageway are separate passageways preventing mixing of combustion and ambient gases.

[0037] Preferably, the method comprising supplying cooling I ambient air to an axial end of the rotor / cooling passageway, and passing the cooling I ambient air through the cooling passageway formed in the turbine blade towards the shroud.

[0038] Preferably, the method comprising a rotor according to the first aspect.

[0039] Preferably, subsequently passing the air in the cooling passageway to a compressor. Alternatively, subsequently passing the air in the cooling passageway to a combustor.

[0040] Preferably, supplying only cooling passageway air to the compressor or combustor.

[0041] Alternatively, and also preferably, mixing the cooling passageway air with (additional) ambient air, before supplying to the compressor.

[0042] According to a third aspect, the present invention provides a gas turbine apparatus comprising a rotor according to the first aspect.

[0043] Preferably, the apparatus comprises an ambient intake, located around the rotor, for supplying ambient air to the cooling passageway of the one or more turbine blades. Alternatively, and also preferably, the apparatus comprises a coolant intake, located around the rotor, in fluid-communication with the compressor so as to receive substantially all the core engine flow.

[0044] Preferably, the ambient intake I coolant intake is annular, having internal surfaces operatively receiving the shroud of the rotor, to allow the shroud and rotor to rotate within the annular intake.

[0045] Preferably, the apparatus comprises an intake channel, extending from the rotor to the compressor or combustor.

[0046] Preferably, a radial outlet of the cooling passageway of the rotor is configured to provide air to the intake channel.

[0047] Preferably, the intake channel comprises a further ambient intake, configured to provide a mixture of cooling passageway air and ambient air to the compressor.

[0048] Preferably, in one version, air supplied to the compressor is only air which has passed through one or more cooling passageways of the one or more rotor blades.

[0049] In an alternative version, but also preferably, air supplied to the compressor is a mixture of air which has already passed through one or more cooling passageways of the one or more rotor blades and ambient air from a further ambient intake.

[0050] Preferably, the gas turbine apparatus is a micro gas turbine (MGT).

[0051] According to a fourth aspect, the invention provides a method for cooling a rotor of a gas turbine apparatus, comprising the method according to the second aspect.

[0052] Preferably, the method comprising supplying ambient air to the cooling passageway of the one or more turbine blades from an ambient intake, located around the rotor. Alternatively, and also preferably, supplying cooling air to the cooling passageway of the one or more turbine blades from a coolant intake located around the rotor in receipt of substantially all the core engine flow.

[0053] Preferably, subsequently supplying air from the cooling passageway of the one or more turbine blades to an intake channel, extending from the rotor to the compressor. Alternatively, and also preferably, subsequently supplying cooling passageway air, from the one or more turbine blades, to an intake channel, extending from the rotor to the combustor.

[0054] Preferably, supplying cooling passageway air, from the one or more turbine blades, to an intake channel, extending from the rotor to the compressor.

[0055] Preferably, the method comprising: supplying only cooling passageway air to the compressor; or mixing cooling passageway air and ambient air, and supplying to the compressor.

[0056] Alternatively, and also preferably, the method comprising: supplying only cooling passageway air to the combustor as an air intake thereof; or mixing cooling passageway air and ambient air, and supplying to the combustor as a mixed intake thereof.

[0057] The present invention may also relate to a rotor for a radial turbine, comprising: a hub; and a plurality of turbine blades, extending from or affixed to at least part of the hub, providing a plurality of combustion gas passageways between respective turbine blades along which combustion gases may pass so as to rotate the rotor, wherein one or more of the plurality of turbine blades comprise at least part of a cooling passageway, the combustion gas passageways and cooling passageway being separate passageways configured to provide movement of combustion and ambient gases in counter-flow conditions acting upon at least part of the turbine blade. Preferably, the rotor further comprises a shroud, the plurality of turbine blades extending between at least part of the hub and at least part of the shroud, providing a plurality of enclosed combustion gas passageways through which combustion gases may pass so as to rotate the rotor. Preferably, the rotor may comprise any one or more features of the first embodiment.

[0058] The present invention may also relate to a method for cooling a rotor of a radial turbine, comprising: passing combustion gases along passageways formed between respective turbine blades of the rotor, so as to rotate the rotor; passing ambient air through a cooling passageway formed in one or more turbine blades of the rotor, so as to cool at least part of the turbine blade, wherein combustion gas passageways and the cooling passageway are separate passageways providing movement of combustion and ambient gases in counter-flow conditions acting upon at least part of the turbine blade. Preferably, the method comprising a rotor according to the first aspect. Preferably, the method comprising any one or more features of the second embodiment.

[0059] Advantageously, the invention proposes merging the recuperator into the turbine to form a heat-exchanging radial turbine.

[0060] In simple terms, the entire mass flow of air of the engine is passed through cooling channels on the inside of the turbine, flu idical ly separated from the combustion gases on the outside via fine tolerances and labyrinth seals. Existing methods of cooling effectively lose the gases from cooling by ejecting them into the exhaust stream; however, in the present invention, the cooling gas is retained and its temperature is raised with waste energy the turbine otherwise cannot consume. Doing this decreases the temperature the turbine is actively experiencing, allowing the TET to be higher and, thus, combining both heat recuperation and higher TET for higher thermal efficiency. With a higher TET, there is also no reason not to include a regular heat exchanger afterwards, since the exhaust gas temperature will still be approximately the same - further increasing the thermal efficiency of the engine.

[0061] Maximising the cooling effect allows the turbine to operate at higher TETs necessary to be compatible with higher compression ratios. Current MGT architecture is limited to a single stage 4-5:1 compression ratio with heat recuperator for a maximum thermal efficiency of about 35% compared to good piston diesel generators of around 39%-45% efficiency. Simply raising the compression ratio of MGTs to 12:1 would lift the isentropic efficiency from about 35% to about 51 % through adding a two-stage compound compression which, after losses, would be operating at a similar level of performance to a piston engine.

[0062] Accordingly, the invention provides the following: air is compressed and routed into internal cooling channels of the turbine; centrifugal force drives the air outward through these channels, absorbing heat from the turbine wall; preheated air exits the turbine and enters the combustion chamber, where fuel is injected and ignited; hot combustion gases flow over the external turbine surfaces, generating shaft power before being exhausted (or passed through a recuperator cycle); and this architecture enables simultaneous blade cooling and intake air preheating, increasing overall cycle efficiency without the usual performance penalty of air-bleed systems.

[0063] Advantageously, the invention provides heat-exchange but no mixing of ambient and combustion gases. One or more cooling channels through the rotor are entirely separate to any channel conducting combustion gases.

[0064] Advantageously, the shrouded rotor is configured to conduct combustion gases through one or more channels and (ambient) cooling air through one or more different channels in opposite directions of flow, so to improve heat exchange.

[0065] Advantageously, the hub wall and shroud provide additional surface area for cooling and, thereby, improve cooling over a similar-sized rotor without shroud. Further, the shroud draws heat energy away from the hub, improving heat transfer away from the rotor.

[0066] Advantageous, the turbine rotor of the present invention may work at higher temperatures, including optionally those above the melting point of the material, as this provides higher thermal efficiency. The rotor blade temperature is reduced through convection of cooling air within internal passageways of the rotor.

[0067] Advantageously, the invention provides larger internal volumes for the flow of cooling air within the turbine blades, increasing both the amount of flow of cooling air and the surface area of the turbine blade in fluid-contact with the cooling air.

[0068] Optionally, up to 100% of the core engine flow is passed through cooling passageways in the turbine blades, significantly improving the flow of cooling air.

[0069] Advantageously, the invention provides the following.

[0070] Optionally, large cooling channels I cooling passageways in each rotor provide significantly more cooling flow, being, say, 30 to 50% of the total engine flow.

[0071] The rotor of the present invention and the radial turbine create lower pumping losses and no additional compressor work as there is no air bypassed from the main flow. The apparatus, optionally, utilises ambient (uncompressed) cooling air, providing a greater AT.

[0072] Optionally, counter-flow heat exchange is more effective and provides improved heat exchange between the ambient cooling air and the combustion gases acting upon the rotor.

[0073] A higher exhaust gas temperature improves heat recouperation and / or combined heat and power (CHP) performance.

[0074] The rotor of the present invention is of simple design and may be forged or cast, making manufacture of the rotor of similar cost to an uncooled rotor. Most preferably, the rotor may be forged in one piece.

[0075] The invention will now be disclosed, by way of example only, with reference to the following drawings, in which:

[0076] Figure 4a is a schematic flow diagram of a first embodiment of gas turbine apparatus according to the present invention;

[0077] Figure 4b is a schematic flow diagram of a second embodiment of gas turbine apparatus according to the present invention;

[0078] Figures 5a, 5b and 5c are perspective and cross-sectional views of a rotor of the gas turbine apparatus of Figures 4a and 4b;

[0079] Figure 6 is a cross-sectional view of a micro gas turbine apparatus in accordance with Figure 4a;

[0080] Figure 7 is a schematic flow diagram of a third embodiment of gas turbine apparatus according to the present invention;

[0081] Figure 8 is a cross-sectional view of a rotor of the gas turbine apparatus of Figure 7.

[0082] Figure 4a shows a flow diagram of a turbine apparatus, identified generally by reference 1 . The apparatus 1 includes a compressor 2, a combustor 3, a turbine 4 and a generator 5.

[0083] Those skilled in the art will understand that some aspects of the turbine apparatus 1 are common with known apparatus 101 , such as the compressor, combustor and the generator. Where the present invention differs greatly, is the turbine 4. The turbine 4 includes a turbine rotor 20 - described in more detail with respect to Figures 5a to 5c - capable of providing counter-flow ambient cooling air through its rotor blades 24. The turbine 4 includes an ambient air inlet 6, capable of supplying ambient air to the turbine rotor 20, and a turbine outlet 7a, capable of supplying air to the compressor 2. The compressor 2 includes a compressor inlet 7b, for receiving air from the turbine 4, and is capable of compressing the received air to provide compressed air to a compressor outlet 8a, which forms a combustor inlet 8b of the combustor 3. The combustor 3 includes a fuel inlet 9, the combustor being capable of burning the fuel in the presence of compressed air supplied by the combustor inlet 8b, and is capable of creating combustion gases. The combustor 3 also includes a combustor outlet 10a, for conveying the combustion gases to the turbine 4, which forms a combustion gas inlet 10b of the turbine 4. The turbine 4 is capable of receiving combustion gases through the combustion gas inlet 10b, and the combustion gases are capable of working upon the rotor 20 to rotate the rotor 20, creating a mechanical, rotational output 11 . The turbine 4 also includes an exhaust outlet 12, for exhausting combustion gases. Optionally, a heat-exchanger (not shown) may be included to recover heat from the exhaust. As can be seen from Figure 4a, the combustion gas inlet 10b and ambient inlet 6, and the exhaust outlet 12 and turbine outlet 7a, are opposed and at different ends of the turbine 4 - thereby, capable of providing counter-flow movement of combustion and cooling gases.

[0084] In use, ambient cooling air is supplied to the turbine through the inlet 6, passed through the rotor 20 of the turbine 4 in a counter-flow condition with respect to combustion gases acting on the rotor 20, and is conveyed to the compressor 2. The air in the turbine outlet 7a I compressor inlet 7b is no longer ambient, since it has gained heat energy from the combustion gases in the turbine 4, and is compressed by the compressor 2, gaining further heat energy. The compressed air is conveyed to the combustor 3, through the inlet 8b, where fuel is burnt in the presence of that compressed air to generate combustion gases, which are then conveyed to the turbine 4. The turbine 4 receives the combustion gases, through the inlet 10b, and the combustion gases interact with the rotor 20 to cause it to rotate, providing the rotational output 11 . The rotational output 11 has two forms, being: a) a rotational input 11a to the generator 5, for generating electricity; and 2) a rotational input 11 b to the compressor 2, for rotating the compressor 2 (and compressing further air). In addition, spent combustion gases exit the turbine 4 through the exhaust outlet 12, which may or may not be fed to a heat-exchanger for additional cooling of the exhaust combustion gases I heating of air between the compressor 2 and combustor 3. Accordingly, counter-flow cooling is provided through the turbine 4.

[0085] It will, therefore, be understood that all air which is compressed by the compressor 2 has previously passed through the turbine 4 as ambient cooling air. Owing to the proposed cooling, turbine 4 is provided with a greater flow of cooling air through the turbine 4. In addition, as the cooling air is ambient and not derived from the compressor feed - which would be warmer owing to its compression - the present invention provides a greater AT of cooling air at the turbine.

[0086] Figure 4b provides a variant of Figure 4a and shows a flow diagram of a turbine apparatus, identified generally by reference T. The embodiment of Figure 4b is based upon the embodiment of Figure 4a and, therefore, only the differences will be discussed - being differences relating to the supply of ambient air.

[0087] The apparatus T includes a first flow adjustor 13, for splitting an ambient air feed 15, having an inlet 13a and a first outlet 13b, which forms the ambient inlet 6. The first flow adjustor 13 also includes a second outlet 13c, for supplying a bypass ambient air feed 16, which bypasses the turbine 4. The apparatus T also includes a second flow adjustor 17, for combining air from the turbine outlet 7a and the bypass ambient air feed 16, having a first inlet 17a for receiving warm(er) air from the turbine outlet 7a and a second inlet 17b for receiving the bypass ambient air feed 16. The flow adjustor 17 includes a combined output 17c, which forms the compressor inlet 7b. The first and second flow adjustors 13; 17 are capable of varying the amount of ambient air which is passed to and through the turbine 4, as a turbine coolant, and recombines that coolant air with bypass air before being passed to the compressor 2.

[0088] In use, an ambient air feed 15 is split at the first flow adjustor 13 and a proportion, x, is diverted to ambient inlet 6 of the turbine 4, to act as turbine coolant, and a proportion, y, is passed to the bypass ambient air feed 16. As described in relation to Figure 4a, cooling ambient air is, therefore, supplied to the turbine 4 through the inlet 6, passed through the rotor 20 of the turbine 4 in a counter-flow condition to combustion gases acting on the rotor 20. Air in the turbine outlet 7a - no longer ambient since it has gained heat energy from the combustion gases in the turbine 4 - is combined with bypass air in the flow adjustor 17. The combined flow of air (x and y) is passed to the compressor through compressor inlet 7b. Operation of the apparatus T from the compressor 2 onwards, being the combustor 3, turbine 4 (with respect to combustion gases) and generator 5, is the same as the apparatus 1 of Figure 4a, and will not be described again.

[0089] In this second embodiment, as the proportion of air diverted to cooling the turbine 4 may be varied, the following configurations are possible. The compressor 2 may receive air which is: a mixture of turbine cooling air and bypass air in varying proportions of x and y, solely turbine cooling air; or solely bypass air. Although the last configuration is potentially possible, as this would not involve any turbine cooling which is contrary to the proposed invention, operating the apparatus this way would be done solely to enable comparison of cooled and uncooled turbine apparatus performance.

[0090] Figures 5a to 5c show a turbine rotor 20 attached to a solid shaft 21 . Such a turbine rotor 20 may be located within the turbine 4 described in the embodiments of Figures 4a and 4b, to provide improved cooling of the turbine 4.

[0091] The rotor 20 includes a hub 22 and a shroud 23 and a plurality of rotor blades 24 extending from the hub 22 radially towards the shroud 23. The hub 22 is partially cylindrical and partially conical, having a disc-like portion 25 forming an extremity of a first end 20a of the rotor 20. The shroud 23 is a frusto-conical shell which, at its end of greatest diameter (towards the first end 20a), and together with the radial rotor blades 24 and the disc-like portion 25, form a plurality of circumferential apertures 26, which are radial combustion gas inlets 26, for receipt of combustion gases from the combustor 3. A plurality of apertures 27 are formed at the second end 20b, which is also disc-like, which apertures are axially oriented combustion gas outlets 27, for exhausting the combustion gases to the exhaust 12. Accordingly, portions of the hub 22, shroud 23 and respective rotor blades 24 form a plurality of internal passageways 28 through the rotor 20, for receipt of the combustion gases, so as to cause rotation of the rotor 20.

[0092] The shroud 23 includes a plurality of axially oriented, longitudinal slots 29, each one located on an outside surface of the shroud 23, directly in-line with a respective rotor blade 24. Each longitudinal slot 29 (cooling gas inlet 29) feeds into a hollowed portion 32 of the rotor blade 24 and, at a far end of the hollowed portion 32, a radial slot 30 (cooling gas outlet 30) is provided on an outside surface of the disc-like portion 25. The longitudinal slot 29, hollowed portion 32 and radial slot 30 provide a passageway 31 through which ambient air may be passed, so as to cool the blade 24.

[0093] Figures 5b and 5c are cross-sectional views along the line a-a of Figure 5a. Each of Figures 5b and 5c show the same cross-sectional view, but from different angles. In use, combustion gas inlets and outlets 26; 27 and internal passageways 28, provide combustion gas passageways for movement of combustion gases from a right-hand side to a left-hand side of the rotor 20, as shown in those figures. In contrast, cooling gas inlets 29, outlets 30 and passageway 31 , provide a cooling passageway for movement of ambient I cooling air from a left-hand side of those images to a right-hand side of the rotor 20, which has the effect of providing counterflow of ambient air through at least a portion of the rotor blade 24. Although just one rotor blade 24 is described in detail, each is provided with the same arrangement which, therefore, increases the flow of cooling air through the rotor 20 of the turbine 4.

[0094] Figure 6 provides a cross-sectional view of a proposed micro turbine apparatus 1 in accordance with the embodiment of Figure 4a - the generator being omitted. Common components have been given the same reference numbers and will not be described in detail again.

[0095] In this embodiment, and as shown, the turbine rotor 20 of Figures 5a to 5c is shown within an annular, ambient inlet 6, which surrounds the rotor 20. The combustor 3 is also annular, surrounding the shaft 21 . The arrangement of components is compact, as this represents a micro turbine apparatus. As would be expected, the shaft 21 , turbine rotor 20 and compressor rotor are freely rotatable within the apparatus 1 . Adjacent surfaces are designed to be extremely precise with near zero tolerance, such that a boundary layer effect is created that prevents easy transmission of gas from one side to the other. In addition, or alternatively, a pressure differential may be used to prevent gas flow going the wrong way.

[0096] The main purpose of Figure 6 is to describe the flow of ambient and combustion gases through the apparatus 1 . Accordingly, movement of air before combustion is shown by the arrow having reference 33 and movement of combustion gases is shown by the arrow having reference 34.

[0097] With reference to arrow 33, ambient air passes through the ambient inlet 6 through one or more of the cooling passageways 28, and into an intake channel 32, which extends from the rotor 20 to the compressor 2. Warm air from the intake channel 32 is passed into the compressor 2, where it is subsequently compressed and passed to the combustor 3 where the fuel is burnt in the presence of that compressed air to provide combustion gases.

[0098] With reference to arrow 34, the combustion gases are passed through combustion gas passageways 28 of the rotor 20, so as to rotate the rotor, and are then passed to the exhaust.

[0099] Importantly, the orientation of the flow of air 33 before combustion and combustion gases 34 through the rotor 20 of the turbine 4 has a counter-flow configuration, which improves heat transfer and, thereby, cooling of the blades 24 of the rotor 20.

[0100] Although in Figure 6 no generator is shown, shaft 21 may be directly or otherwise connected a generator to provide electricity.

[0101] Figure 6 may also be relatively easily modified to provide an additional ambient inlet, in a configuration along the lines of the embodiment of Figure 4b. Thereby, a proportion of cooling air may be diverted and used to cool the turbine, before being recombined with bypass air before the combined air stream is compressed.

[0102] Those skilled in the art will understand that the ambient cooling air is passed through the turbine in the opposite direction to the combustion gases and now leaves in a condition that is noticeably hotter than before. The effect of this is to reduce the actual temperature of the rotor I turbine, encased within the turbine housing. Heat energy is being removed from the turbine which would allow the turbine to be used at a temperature which is in excess of the melting point of the material of the turbine.

[0103] Although the turbine blades are shown as straight in the first and second embodiments, the turbine blades may be curved in all planes. A third embodiment of gas turbine apparatus is described in relation to Figures 7 and 8.

[0104] Figure 7 shows a schematic flow diagram of a turbine apparatus, identified generally by reference 301 . The apparatus 301 includes a compressor 302, a combustor 303, a turbine 304 and a generator 305. Those skilled in the art will understand that some aspects of the turbine apparatus 301 are common with apparatus 1 of Figure 4a, such as the compressor, combustor, turbine and the generator. In fact, the turbine 304 is identical to turbine 4 of Figure 4a; however, cooling air is supplied differently and flows differently through turbine 304, as described further below.

[0105] As before, where there the present embodiment differs greatly from the prior art is turbine 304. Turbine 304 includes a turbine rotor 320 - described in more detail in connection with Figure 8, being capable of providing a substantial flow of cooling air through its rotor blades 324. In fact, this third embodiment is configured so as to direct substantially all, if not all, of the core engine flow from the compressor through cooling passageways of the turbine rotor 320, significantly improving cooling.

[0106] The compressor 302 includes a compressor inlet 306, for receiving ambient cooling air, and is capable of compressing the received air to provide compressed air to a compressor outlet 307, which forms an inlet 308 to the turbine 304. The turbine 304 is capable of supplying cooling air to the turbine rotor 320, and includes a turbine outlet 309, capable of supplying air to the combustor 303.

[0107] The combustor 303 includes a fuel inlet 310 and air inlet 311 , the combustor 303 being capable of burning the fuel in the presence of compressed air supplied by the turbine 304, and is capable of creating combustion gases. The combustor 303 also includes a combustor outlet 312, for conveying the combustion gases to the turbine 304, which forms a turbine combustion gas inlet 313 of the turbine 304. The turbine 304 is capable of receiving combustion gases through the turbine combustion gas inlet 313, and the combustion gases are capable of working upon the rotor 320 to rotate the rotor 320, creating a mechanical, rotational output 314. The turbine 304 also includes an exhaust outlet 315, for exhausting combustion gases. Optionally, a heat-exchanger (not shown) may be included to recover heat from the exhaust.

[0108] In use, ambient cooling air is supplied to the compressor 302 through compressor inlet 306, compressed by the compressor 302, gaining heat energy. The compressed air is conveyed to the turbine 304, via compressor outlet 307 and turbine inlet 308, as a cooling flow of air being substantially all, if not all, the core engine flow, and is passed through the rotor 320 of the turbine 304. After exiting the turbine 304 through turbine outlet 309, the compressed air, having gained additional heat energy from the turbine, is conveyed to the combustor 303, via air inlet 311 , where fuel is burnt in the presence of that compressed air to generate combustion gases, which are then conveyed to the turbine 304. The turbine 304 receives the combustion gases, via combustor outlet 312 and turbine combustion gas inlet 313, and the combustion gases interact with the rotor 320 to cause it to rotate, providing the rotational output 314. The rotational output 314 has two forms, being: a) a rotational input 314a to the generator 305, for generating electricity; and 2) a rotational input 314b to the compressor 302, for rotating the compressor 302 (and compressing further air). In addition, spent combustion gases exit the turbine 304 through the exhaust 315, which may or may not be fed to a heat-exchanger for additional cooling of the exhaust combustion gases I heating of air between the compressor 302 and combustor 303. Although the turbine blades are shown as straight in the present embodiment, the turbine blades may be curved in all planes.

[0109] Figure 8 is a cross-sectional view of a turbine rotor 320 attached to a solid shaft 321 - being similar to Figure 5c. Such a turbine rotor 320 may be located within the turbine 304 described in the embodiment of Figure 7, to provide improved cooling of the turbine 4.

[0110] The rotor 320 includes a hub 322 and a shroud 323 and a plurality of rotor blades 324 extending from the hub 322 radially towards the shroud 323. The hub 322 is partially cylindrical and partially conical, having a disc-like portion 325 forming an extremity of a first end 320a of the rotor 320. The shroud 323 is a frusto-conical shell which, at its end of greatest diameter (towards the first end 320a), and together with the radial rotor blades 324 and the disc-like portion 325, form a plurality of circumferential apertures 326, which are radial combustion gas inlets 326, for receipt of combustion gases from the combustor 303. A plurality of apertures 327 are formed at the second end 320b, which is also disc-like, which apertures are axially oriented combustion gas outlets 327, for exhausting the combustion gases to the exhaust 312. Accordingly, portions of the hub 322, shroud 323 and respective rotor blades 324 form a plurality of internal passageways 328 through the rotor 320, for receipt of the combustion gases, so as to cause rotation of the rotor 320.

[0111] The shroud 323 includes a plurality of axially oriented, longitudinal slots 329, each one located on an outside surface of the shroud 323, directly in-line with a respective rotor blade 324. Each longitudinal slot 329 (cooling gas outlet 329) is in fluid communication with a hollowed portion 332 of the rotor blade 324 and, at a far end of the hollowed portion 332, a radial slot 330 (cooling gas inlet 330) is provided on an outside surface of the disc-like portion 325 for receipt of cooling gas flow. The radial slot 330, hollowed portion 332 and longitudinal slot 329 provide a passageway 331 through which cooling air may be passed, so as to cool the blade 324.

[0112] In use, combustion gas inlets and outlets 326; 327 and internal passageways 328, provide combustion gas passageways for movement of combustion gases from a right-hand side to a left-hand side of the rotor 320, as shown in Figure 8. In contrast to Figure 4a, cooling gas outlets 329, inlets 330 and passageways 331 of this third embodiment provide a cooling passageway for movement of cooling air in a similar direction to the combustion gases, being also from the right-hand side of that image to the left-hand side of the rotor 320, which has the effect of providing cooling of at least a portion of the rotor blade 324.

[0113] Although just one rotor blade 324 is described in detail, each is provided with the same arrangement which, therefore, increases the flow of cooling air through the rotor 320 of the turbine 304. Accordingly, substantially all, if not all, of the core engine flow is passed through the cooling passageways 331 .

Claims

Claims:1 .) A rotor for a radial turbine, comprising: a hub; and a plurality of turbine blades, extending from or affixed to at least part of the hub, providing a plurality of combustion gas passageways between respective turbine blades along which combustion gases may pass so as to rotate the rotor, wherein one or more of the plurality of turbine blades comprise at least part of a cooling passageway, the combustion gas passageways and cooling passageway being separate passageways configured to provide separate movement of combustion and ambient gases, wherein the rotor further comprises a shroud, the plurality of turbine blades extending between at least part of the hub and at least part of the shroud, providing a plurality of enclosed combustion gas passageways through which combustion gases may pass so as to rotate the rotor.2.) A rotor as claimed in claim 1 , wherein the separate passageways are configured to provide movement of combustion and ambient gases in opposite directions acting upon at least part of the turbine blade.3.) A rotor as claimed in claim 1 or claim 2, wherein the separate passageways are configured to provide movement of combustion and ambient gases in opposite directions along at least part of the turbine blade.4.) A rotor as claimed in any preceding claim, wherein the separate passageways are configured to prevent mixing of combustion and ambient gases.5.) A rotor as claimed in any preceding claim, wherein the cooling passageway comprises an internal shape of the turbine blade being an approximate inversion to an external shape of the turbine blade at or through the same cross-section.6.) A rotor as claimed in any preceding claim, wherein an external surface of the shroud comprises a cooling air inlet, or outlet, operatively connected to an internal channel of the one or more blades, which is operatively connected to a cooling air outlet, or inlet, respectively, provided on an external surface of the hub and / or a further surface or part of the shroud.7.) A rotor as claimed in any preceding claim, wherein the shroud is a frusto- conical shell which, together with the hub and plurality of turbine blades, provide enclosed combustion gas passageways having radial inlets, and axial outlets.8.) A rotor as claimed in any preceding claims, wherein the cooling passageway comprises: an inlet or outlet on an external surface of the shroud; and / or an outlet or inlet on an external surface of the hub, or further surface or part of the shroud.9.) A rotor as claimed in claim 8, or anyone of claims 6 to 7, wherein:the inlet or outlet to the cooling passageway is formed on a / the frusto-conical shell surface of the shroud; and / or the outlet or inlet from the cooling passageway is formed on a radial surface of the hub.10.) A rotor as claimed in any preceding claim, wherein the cooling passageway comprises: an aperture or slot, extending axially along the shroud so as to provide a hollow region of the blade; and / or an aperture or slot, located radially on a disc portion of the hub and substantially in-line with at least part of the blade.11.) A rotor as claimed in any preceding claim, wherein the turbine blades comprise radially arranged blades of substantially arcuate, sinuous or straight profile, or combinations thereof.12.) A rotor as claimed in any preceding claim, wherein the enclosed combustion gas passageways comprise arcuate, sinuous, straight and / or partially conical surfaces, or combinations thereof.13.) A rotor as claimed in any preceding claim, wherein an internal cooling surface area of one or more of the plurality of turbine blades is about 50% to about 85% of an external surface area of the same turbine blade exposed to heat.14.) A rotor as claimed in any preceding claim, wherein a cross-sectional surface area of the cooling passageway of the turbine blade is about 40% to about 60% of the total cross-sectional surface area of the turbine blade at the same cross-section.15.) A method for cooling a rotor of a radial turbine, comprising: passing combustion gases along passageways formed between respective turbine blades of the rotor, so as to rotate the rotor; passing cooling air through a cooling passageway formed in one or more turbine blades of the rotor, so as to cool at least part of the turbine blade, the combustion gas passageways and the cooling passageway are separate passageways providing separate movement of combustion and ambient gases, wherein the rotor comprises a shroud, the turbine blades extending between at least part of a hub of the rotor and at least part of the shroud, providing a plurality of enclosed combustion gas passageways along which combustion gases pass so as to rotate the rotor.16.) A method as claimed in claim 15 comprising supplying cooling air to an axial end of the cooling passageway I rotor, and passing the cooling air through the cooling passageway formed in the turbine blade towards the shroud.17.) A method as claimed in claim 15 or claim 16 comprising a rotor as claimed in any one of claims 1 to 14.18.) A gas turbine apparatus comprising a rotor as claimed in any one of claims 119.) A gas turbine apparatus as claimed in claim 18, wherein the apparatus comprises: an ambient intake, located around the rotor, for supplying ambient air to the cooling passageway of the one or more turbine blades; or a coolant intake, located around the rotor, in fluid-communication with the compressor so as to receive substantially all the core engine flow.20.) A gas turbine apparatus as claimed in claim 18 or claim 19, wherein the ambient intake I coolant intake is annular, having internal surfaces operatively receiving the shroud of the rotor, to allow the shroud and rotor to rotate within the annular intake.21 .) A gas turbine apparatus as claimed in any one of claims 18 to 20, wherein the apparatus comprises a further ambient intake, configured to provide a mixture of cooling passageway air and ambient air to the compressor or combustor.22.) A method for cooling a rotor of a gas turbine apparatus, comprising the method of any one of claims 15 to 17.23.) A method as claimed in claim 22 comprising: supplying ambient air to the cooling passageway of the one or more turbine blades from an ambient intake, located around the rotor; or supplying cooling air to the cooling passageway of the one or more turbine blades from a coolant intake located around the rotor in receipt of substantially all the core engine flow.24.) A method as claimed in claim 22 or claim 23 comprising: subsequently supplying cooling passageway air, from the one or more turbine blades, to an intake channel, extending from the rotor to the compressor; or subsequently supplying cooling passageway air, from the one or more turbine blades, to an intake channel, extending from the rotor to the combustor.25.) A method as claimed in claim 23 or claim 24 comprising: supplying only cooling passageway air to the compressor; or mixing cooling passageway air and ambient air, and supplying to the compressor.26.) A method as claimed in claim 23 or claim 24 comprising: supplying only cooling passageway air to the combustor as an air intake thereof; or mixing cooling passageway air and ambient air, and supplying to the combustor as a mixed intake thereof.

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