WO2011054760A1

WO2011054760A1 - A cooling scheme for an increased gas turbine efficiency

Info

Publication number: WO2011054760A1
Application number: PCT/EP2010/066513
Authority: WO
Inventors: Anton Winkler; Urs Benz; Andre Theuer; Diane Lauffer; Madhavan Poyyapakkam
Original assignee: Alstom Technology Ltd
Priority date: 2009-11-07
Filing date: 2010-10-29
Publication date: 2011-05-12
Also published as: US20120324863A1; US8572980B2; EP2496885A1; EP2496885B1

Abstract

A burner for a combustion chamber of a turbine, with an injection device for the introduction of at least one fuel into the burner. The injection device has at least one body (22) arranged in the burner with at least two nozzles for introducing the fuel into the chamber, the body having a streamlined cross - sectional profile extending with a longitudinal direction perpendicularly to a main flow direction prevailing in the burner and two lateral surfaces (33) essentially parallel to the main flow direction joined at their upstream side by a leading edge and joined at their downstream side forming a trailing edge, the nozzles (15) being distributed along said trailing edge. The body comprises an enclosing outer wall (37) defining said streamlined cross - sectional profile, wherein within this outer wall, there is provided a longitudinal inner air plenum (51) for the introduction of air into the injection device. The air plenum is provided with holes (56) such that air exiting through this holes impinges the inner side of the leading edge.

Description

A cooling scheme for an increased gas turbine efficiency TECHNICAL FIELD

The present invention relates to a novel fuel lance for a burner for a primary combustion chamber of a turbine or secondary combustion chamber of a turbine with sequential combustion having a first and a secondary combustion chamber, for the introduction of at least one gaseous and/or liquid fuel into the burner. Modifications to the cooling scheme of the fuel lance are proposed to increase the GT engine efficiency as well as to simplify the design. The invention normally requires good fuel-air mixing obtained with low momentum flux ratios.

PRIOR ART

In order to achieve a high efficiency, a high turbine inlet temperature is required in standard gas turbines. As a result, there arise high NOx emission levels and high life cycle costs. These problems can be mitigated with a sequential combustion cycle, wherein the compressor delivers nearly double the pressure ratio of a conventional one. The main flow passes the first combustion chamber (e.g. using a burner of the general type as disclosed in EP 1 257 809 or as in US 4,932,861, also called EV combustor, where the EV stands for environmental), wherein a part of the fuel is combusted. After expanding at the high- pressure turbine stage, the remaining fuel is added and combusted (e.g. using a burner of the type as disclosed in US 5,431,018 or US 5,626,017 or in US 2002/0187448, also called SEV combustor, where the S stands for sequential). Both combustors contain premixing burners, as low NOx emissions require high mixing quality of the fuel and the oxidizer.

Since the second combustor is fed by expanded exhaust gas of the first combustor, the operating conditions allow self ignition (spontaneous ignition) of the fuel air mixture without additional energy being supplied to the mixture. To prevent ignition of the fuel air mixture in the mixing region, the residence time therein must not exceed the auto ignition delay time. This criterion ensures flame-free zones inside the burner. This criterion poses challenges in obtaining appropriate distribution of the fuel across the burner exit area. SEV-burners are currently designed for operation on natural gas and oil only. Therefore, the momentum flux of the fuel is adjusted relative to the momentum flux of the main flow so as to penetrate into the vortices. The subsequent mixing of the fuel and the oxidizer at the exit of the mixing zone is just sufficient to allow low NOx emissions (mixing quality) and avoid flashback (residence time), which may be caused by auto ignition of the fuel air mixture in the mixing zone. The cross flow injection concept used in the current SEV-fuel injection devices (SEV fuel lances) necessitates high-pressure carrier air supply, which reduces the overall efficiency of the power plant.

SUMMARY OF THE INVENTION

It is the object of the present invention to provide an improved fuel injection device for combustion chambers of gas turbines. In particular a new injection device shall be proposed which can be operated with low pressure (carrier) air which at the same time acts as carrier air for fuel injection as well as cooling air.

More specifically, the present invention relates to a burner for a combustion chamber of a turbine, preferably of a gas turbine, with an injection device for the introduction of at least one gaseous and/or liquid fuel into the burner. The injection device has at least one body or lance which is arranged in the burner and extends into the burner cavity, wherein the at least one body has at least two nozzles for introducing the at least one fuel into the burner. The burner may also be designed as an element comprising more than one such body located next to each other, e.g. a burner with three bodies located next to each other, normally each with a different inclination angle with respect to the main flow direction. The at least one body is preferentially configured as a streamlined body which has a streamlined cross-sectional profile and which extends with a longitudinal direction perpendicularly (or at a slight inclination) to a main flow direction prevailing in the burner. The body has two lateral surfaces normally at least for one central body essentially parallel to the main flow direction and converging, i.e. inclined for the others. These lateral surfaces are joined at their upstream side by a leading edge portion of the body (typically a rounded portion) and joined at their downstream side forming a trailing edge (typically a sharp edge). The at least two nozzles are preferably located at different longitudinal positions along the preferentially essentially straight trailing edge of the body. So they are normally distributed along said trailing edge. The body comprises an enclosing outer wall defining said streamlined cross-sectional profile. Within this outer wall (in the cavity defined thereby), there is provided a longitudinal inner carrier air plenum (typically a tubular structure) for the introduction of carrier air into the injection device. The carrier air plenum is specifically provided with holes such that carrier air exiting through these holes impinges on the inner side of the leading edge portion of the body. The sizes and distribution of these holes are preferentially designed in order to guarantee a uniform carrier air distribution.

In one burner at least one such injection device is located, preferably at least two such injection devices are located within one burner, even more preferably three such injection devices or flutes are located within one burner.

These holes in the carrier air plenum are typically distributed along the longitudinal direction and also in the direction orthogonal thereto, so along the rounded leading edge inner shape.

Such injection device can be used in a primary burner but preferably it is used in a secondary burner located downstream of a primary combustion chamber responsible for supplying a secondary combustion chamber with fuel, wherein in this secondary combustion chamber the fuel is auto igniting. A burner according to this design is typically such that upstream of the body and downstream of the last row of rotating blades of the high-pressure turbine there are no additional vortex generators necessary, and preferably also no additional flow conditioning elements.

As mentioned, according to the invention at least two nozzles are located at the trailing edge of the body. Preferably between 4 and 30 nozzles, preferentially located in equidistant distribution along the trailing edge, inject fuel and/or carrier gas essentially parallel to the main flow direction (in-line injection).

Generally the injection device can be used for gas or liquid fuel.

According to a preferred embodiment, the carrier air plenum is a tubular duct located in the upstream portion of the cavity defined by the outer wall. The expression tubular duct shall not imply a circular cross-section of the duct, the cross-section may be circular, oval, preferably the cross-section of the tubular duct has, at least in the portion facing the leading edge part of the outer wall, a similar shape as the outer wall on its inner side. Preferentially, the wall of the tubular duct is distanced from the outer wall leaving an interspace in between for circulation of carrier air, leading to impingement cooling of the inner wall and at the same time to convective cooling thereafter. Preferably the wall of the tubular duct in the region facing the outer wall is running essentially parallel thereto, such that the cooling channel formed between these two walls has an essentially constant cross- section in particular along the longitudinal direction. Further preferably the distance between the wall of the tubular duct and the outer wall is established/maintained by at least one distance keeping element. Such a distance keeping elements can be located at the outer wall and/or at the wall of the tubular duct, it may for example be in the form of protrusions and/or ridges provided on the inner side of the outer wall.

According to a further preferred embodiment, the carrier air plenum extends essentially along the full length of the body. Preferably, at the bottom end it is closed by a bottom plate, which can also be provided with holes for impingement cooling of a bottom plate of the body.

A further preferred embodiment is characterised in that air exiting from the carrier air plenum is used as carrier air of the injection devices. In other words carrier air for the fuel injection is exclusively provided by this carrier air plenum, so the carrier air for the fuel injection first takes the function of cooling of the injection device and after that takes a function of carrier air for fuel injection. Preferentially the carrier air exits at the injection devices via an annular slit enclosing a central fuel jet. The central fuel jet normally exits via an annular fuel slit, so the central fuel jet is also an annular fuel jet enclosed by the carrier air.

Yet another embodiment of the invention is characterised in that within the enclosing outer wall defining said streamlined cross-sectional profile, there is further provided a longitudinal inner fuel tubing for the introduction of liquid and/or gaseous fuel. In other words the carrier air plenum and this longitudinal inner fuel tubing run parallel within the cavity formed by the outer wall. Normally the longitudinal inner fuel tubing is provided with branching off tubing leading to the at least two nozzles. Preferably the carrier air plenum is located in the upstream portion of the cavity defined by the outer wall while the longitudinal inner fuel tubing is located in the downstream portion of the cavity defined by the outer wall. Like this, when the carrier air plenum is exclusively located in the upstream portion of the cavity while the longitudinal inner fuel tubing is exclusively located in the downstream portion of the cavity, the fuel supply parts are optimally shielded from the heat which is predominantly a problem at the leading edge of the device. Preferably the wall of the carrier air plenum is distanced from the wall of the longitudinal inner fuel tubing for circulation of carrier air. Preferentially in a cross-sectional view the distance between the wall of the inner fuel tubing and the outer wall and the distance between the wall of the carrier air plenum and the outer wall is essentially the same so the couple of the inner fuel tubing and the carrier air plenum tubing have a similar outline as the inner side of the outer wall structure leading to an optimum flow cavity for the carrier air. The wall portions of the inner fuel tubing and a carrier air plenum tubing facing each other are normally located essentially perpendicular to the main flow direction, and are preferentially distanced from each other such that carrier air may also circulate between these two walls. In other words, the longitudinal inner fuel tubing is preferably circumferentially distanced from the outer wall, defining an interspace for the delivery of carrier air to the at least one nozzle.

According to yet another preferred embodiment, air exiting from the carrier air plenum exits the injection device via effusion holes, apart from taking over the carrier air function in the fuel nozzles. Such effusion holes can for example be located at the trailing edge of the injection device and/or at the lateral surfaces of the injection device and/or at the leading edge of the injection device and/or at large scale mixing devices of the injection device. Such large scale mixing devices can for example be vortex generators located at the lateral surfaces upstream of the nozzles which are provided with perforations through which the carrier air can penetrate.

According to a further preferred embodiment, the at least two nozzles have their outlet orifices downstream of the trailing edge of the streamlined body, leading to an optimum mixing while necessitating only low pressure carrier air. Preferably the distance between the essentially straight trailing edge at the position of the nozzle, and the outlet orifice of said nozzle, measured along the main flow direction, is at least 2 mm, preferably at least 3 mm, more preferably in the range of 4-10 mm.

According to a further preferred embodiment, the streamlined body has a cross-sectional profile which is mirror symmetric (excluding the vortex generators, which may also not be mirror symmetric in their distribution on the lateral faces) with respect to the central plane of the body.

Typically and preferentially the at least one nozzle injects fuel and/or carrier gas at an inclination angle between 0-30° with respect to the main flow direction, so preferentially there is in-line injection of the fuel.

According to a further preferred embodiment, within said longitudinal inner fuel tubing provided for gaseous fuel there is provided a second inner fuel tubing for a second type of fuel, wherein preferably this second type of fuel is a liquid fuel and wherein further preferably gaseous fuel is delivered by the interspace between the walls of said longitudinal inner fuel tubing and the walls of the second inner fuel tubing. As mentioned above, according to an embodiment of the invention upstream of the at least one nozzle on at least one lateral surface there is located at least one vortex generator. The vortex generator preferentially has an attack angle in the range of 15-20° and/or a sweep angle in the range of 55-65°. Generally speaking, vortex generators as they are disclosed in US 5,80,360 to as well as in US 5,423,608 can be used in the present context, the disclosure of these two documents being specifically incorporated into this disclosure. Typically at least two nozzles are arranged at different positions along said trailing edge, and upstream of each of these nozzles at least one vortex generator is located. Vortex generators to adjacent nozzles can be located at opposite lateral surfaces, and preferably more than three, most preferably at least four, nozzles are arranged along said trailing edge and vortex generators are altematingly located at the two lateral surfaces or downstream of each vortex generator there are located at least two nozzles.

The vortex generator can, as mentioned above, be provided with cooling elements, wherein preferably these cooling elements are effusion cooling holes provided in at least one of the surfaces of the vortex generator, and wherein even more preferably the film cooling holes are fed with air from the carrier gas feed also used for the fuel injection.

According to yet another preferred embodiment, the streamlined body extends across essentially the entire flow cross section between opposite walls of the burner.

Preferably the burner is an annular burner arranged circumferentially with respect to a turbine axis, and between 10- 100 streamlined bodies, preferably between 40 - 80 streamlined bodies are arranged around the circumference, more preferably all of them being equally distributed along the circumference.

The fuel is typically injected from the nozzle together with a carrier air stream which is supplied by the carrier air plenum, and the carrier air is low pressure air with a pressure in the range of 10-22 bar, preferably in the range of 16-22 bar, and further preferably this carrier air is directly derived from a compressor stage without subsequent cooling.

The present invention furthermore relates to the use of a burner as defined above in a secondary combustion chamber. This in particular for the combustion under high reactivity conditions, preferably for the combustion at high burner inlet temperatures and/or for the combustion of MBtu fuel, normally with a calorific value of 5000-20,000 kJ/kg, preferably 7000-17,000 kJ/kg, more preferably 10,000-15,000 kJ/kg, most preferably such a fuel comprising hydrogen gas.

Further embodiments of the invention are laid down in the dependent claims. BRIEF DESCRIPTION OF THE DRAWINGS

Preferred embodiments of the invention are described in the following with reference to the drawings, which are for the purpose of illustrating the present preferred embodiments of the invention and not for the purpose of limiting the same. In the drawings,

Fig. 1 shows a secondary burner located downstream of the high-pressure turbine together with the fuel mass fraction contour (right side) at the exit of the burner;

Fig. 2 shows an aerodynamic ally optimised lance arrangement in a central axial cut through the central lance in a), in b) a cut along the line A in a), and in c) a cut along C-C in a);

Fig. 3 shows a perspective view onto the group of lance bodies and their interior structure;

Fig. 4 shows a perspective view onto one half of the lance arrangement wherein the outer wall structure on the upper part is present;

Fig. 5 shows a perspective view onto a complete lance arrangement wherein the outer wall structure on the upper part is removed;

Fig. 6 shows an aerodynamically optimised lance arrangement according to a second embodiment in a central axial cut through the central lance.

DESCRIPTION OF PREFERRED EMBODIMENTS

Several design modifications to the existing secondary burner (SEV) designs are proposed to introduce a low pressure drop complemented by rapid mixing e.g. for highly reactive fuels and operating conditions. This invention targets for a low pressure drop fuel lance system for a reheat flute lance and burner. The (50% or higher) reduced fuel pressure drop in the flute lance is due to less design complexity and the elimination of high momentum flux fuel jets required for the state of the art cross flow lance configurations. Herein, a fuel lance cooling concept for inline fuel injection is proposed which eliminates the need for high-pressure (carrier air and fuel) requirements. An injection system with lower fuel pressure drop increases the likelihood of avoiding the use of fuel compression for the SEV. The low BTU and H2 fuels require that fuel pressure drops inside the passage have to be acceptable.

The key advantages can be summarised as follows: • Low fuel momentum flux of the fuel jets in the reheat lances reduce the fuel pressure requirement.

• The lower fuel pressure drop in the lance offers the possibility for fuel staging to control emissions and pulsations.

· Lower fuel pressure drop in the inline injectors allow for injecting H2 or Syngas with a reasonable pressure.

• Flute design offers uniform fuel distribution across the injectors.

In particular, the invention relates to situations where the high-pressure carrier air/cooling air supply, which is necessary in constructions according to the state-of-the-art with pressures in the range of 25-35 bar, is to be replaced by medium pressure carrier air/cooling air supply typically in the range of 10-22 bar, i.e. air, which is not taken from the very last compressor stage but from an intermediate stage. The advantages are as follows:

• The overall GT efficiency increases. Still, the cooling air bypasses the high- pressure turbine, but at least medium pressure carrier air/cooling air is compressed to a lower pressure level compared to high-pressure carrier/cooling air and does not need to be cooled down.

• The design of the cooling air passage can be simplified.

• The fuel is shielded in order to slow down the reactivity of the fuel air mixture · Sufficient cooling is provided to the lance.

The momentum flux of the fuel needn't be increased, if the injector is designed accordingly, i.e. if the dependence of the mixing behavior on the momentum flux ratio is weak.

Problems of the state of the art solutions:

The cross flow fuel jet underlying principle of the current SEV technology incur very high- pressure drop due to complex flow features and high momentum flux of the fuel jet. The supply fuel pressure for the SEV is drawn from the EV gas compressors, which is high in order to obtain a high momentum flux ratio (typically around 8). The fuel gas pressure requirements for the reheat fuel lances should however be decreased in order to minimize the hardware costs and auxiliary power consumption by modifying the gas compressors for future engines.

With respect to performing a reasonable fuel air mixing, the following components of current burner systems are of interest: • At the entrance of the SEV combustor, the main flow must be conditioned in order to guarantee uniform inflow conditions independent of the upstream disturbances, e.g. caused by the high-pressure turbine stage.

• Then, the flow must pass four vortex generators.

· For the injection of gaseous and liquid fuels into the vortices, fuel lances are used, which extend into the mixing section of the burner and inject the fuel(s) into the vortices of the air flowing around the fuel lance.

To this end figure 1 shows a conventional secondary burner 1. The burner, which can be an annular combustion chamber or one with rectangular cross-section, is bordered by opposite walls 3. These opposite walls 3 define the flow space for the flow 14 of oxidizing medium. This flow enters as a main flow 8 from the high pressure turbine, i.e. behind the last row of rotating blades of the high pressure turbine which is located downstream of the first combustor. This main flow 8 enters the burner at the inlet side 6. First this main flow 8 passes flow conditioning elements 9, which are typically turbine outlet guide vanes which are stationary and bring the flow into the proper orientation. Downstream of these flow conditioning elements 9 vortex generators 10 are located in order to prepare for the subsequent mixing step. Downstream of the vortex generators 10 there is provided an injection device or fuel lance 7 which typically comprises a foot 16 and an axial shaft 17 extending further downstream like a rod. At the most downstream portion of the shaft 17 fuel injection takes place, in this case fuel injection takes place via orifices/nozzles which inject the fuel in a direction perpendicular to flow direction 14 (cross flow injection).

Downstream of the fuel lance 7 there is the mixing zone 2, in which the air, bordered by the two walls 3, mixes with the fuel and then at the outlet side 5 exits into the combustion space 4 where self-ignition takes place.

At the transition between the mixing zone 2 and the combustion space 4 there is typically a transition 13, which may be in the form of a step, or as indicated here, may be provided with round edges and also with stall elements for the flow. The combustion space is bordered by the combustion chamber wall 12.

This leads to a fuel mass fraction contour 11 at the burner exit 5 as indicated on the right side of figure 1.

The fuel lance is equipped with a carrier air passage, which is needed for the following reasons:

• The carrier air is slowing down the reactivity of the fuel air mixture by local effects on both, temperature and equivalence ratio.

• The carrier air is also used for cooling the lance.

• SEV-burners are currently designed for operation on natural gas and oil. The carrier air increases the momentum flux of the fuel in order to penetrate the vortices and allow a good fuel air mixing behavior.

The system due to the last requirement given above, needs carrier air, normally taken from the last compressor stage of the gas turbine and this carrier air further needs to be cooled down. This has the following drawbacks:

• The high-pressure carrier air drawn from the last compressor stage is bypassing the high pressure turbine thus resulting in efficiency losses.

• The necessary cooling down of the high-pressure carrier air result in additional efficiency losses.

• The further drawback is related to the complicated design of the current SEV system.

The cooling air of the burner for cooling the combustion chamber walls 12 as well as the walls 2 of the combustor and the lance is currently taken from a low pressure air plenum. The air is then cooling both, the burner and the front panel 13 with effusion cooling. The need for additional high-pressure cooled down carrier air for the assistance of the fuel injection process and the cooling of the lance is resulting in additional design efforts for the high-pressure carrier air supply.

With the proposed novel cooling scheme and appropriate injector design the drawbacks of using high-pressure carrier air can be avoided.

With low enough fuel pressure requirements, as made possible by using streamlined bodies as fuel injection devices combined with in-line fuel injection, a sequential burner can be fed without fuel compression i.e. it is possible to feed the sequential burner with network pressure only (typically in the range of 10-20 bar, as compared to high-pressure as conventionally necessary which is in the range of 25-35 bar). At the same time carrier air pressure can then be as low as in the range of 10-22 bar for the assistance of this in-line injection process, so cooled down high-pressure carrier air with pressures in the range of 25-35 bar is not necessary any more. However the question is how such low pressure carrier air can then still be efficiently used at the same time for cooling of the lance, as it is desirable to use the carrier air supply used for assisting the fuel injection at the same time also for cooling the lance. The proposed solution can be summarize as follows:

Flutelike injectors with an aerodynamically optimized lance body are considered as injectors. The body is designed to mitigate non-uniformities of the flow, which is coming from the high pressure turbine. The fuel injector can be equipped to allow axial injection of the fuel. In order to enhance the spreading of the jets, large scale mixing devices may be incorporated. In water channel tests, the dependence upon the momentum flux ratio was determined. It was seen that the mixing behaviour of the in-line-configuration hardly depends on the momentum flux ratio, thus not requiring high pressure carrier air for the sake of momentum flux ratio any more.

The challenge is now shifted to providing a cooling scheme for the fuel lance, which can perform the cooling as well as the fuel shielding at a reasonable pressure drop.

Herein, effusion cooling, impingement cooling and convective cooling are combined in order to yield the desired performance.

Two embodiments are shown in the following to combine the cooling to the fuel shielding. Embodiment 1 (see figures 2-5):

The cooling of the lance balcony 18 is carried out as impingement cooling. After cooling the lance balcony 18, the cooling air is entering a carrier air plenum 51. The plenum 51 is equipped with several holes 56. These are chosen in diameter as such that a uniform distribution of the carrier air along the injectors is ensured. From the carrier plenum 51, the air impinges the inner side of the leading edge of the injectors or flutes 22. The air then cools the sidewall convectively. The cooling air is leaving the injector through various passages, e.g. three passages: This may be the large scale mixing devices 23 (e.g. vortex generators), the trailing edge 24 and/or annular slits at the injector holes. The split between each of the passages vortex generators 23, trailing edge 24 and injector 15 holes is adjusted to allow sufficient cooling of the components and a combustion behaviour as desired. Within each of the passages, the cross section is designed as such that the critical area is close to the exit of the passage, thus ensuring uniform cooling air distribution.

In more detail this concept shall be discussed with reference to figures 2-5. In this first embodiment a burner arrangement is given, in which three bodies 22 or lances are elements of a burner arrangement with three such flutes or streamlined bodies 22. This burner arrangement is to be located in the wall 3 of a general burner set-up as illustrated in figure 1.

The burner arrangement comprises a burner plate 18, also called balcony, to which the three bodies 22 are attached next to each other (with slightly different inclination angles with respect to the main flow direction 14) . They extend into the mixing space or mixing zone 2.

Each of these bodies 22 has an outer wall 37 with two lateral surfaces 33 which are arranged essentially parallel to the main flow 14 of the combustion gases.

This outer wall 37 forms a cavity within the body 22 which at the leading edge 25 joins the two lateral walls 33 in a rounded manner, while at the trailing edge 24 the lateral walls form a sharp edge, similar to a wing like structure.

The leading edge 25 and the trailing edge 24 are essentially parallel to each other along a longitudinal direction and extend perpendicularly to the main flow direction 14 of the combustion gases. Such a burner arrangement is thus located in a secondary combustion chamber of a gas turbine.

In this cavity formed by the outer wall 37 there is located, in the region adjacent to the leading edge, a carrier air channel or carrier air plenum 51, which is given as a tubular or channel like structure.

In the trailing edge region of this cavity formed by the outer wall 37, there is located a longitudinal inner fuel tubing 36 for fuel supply of the nozzles 15, which are located at the trailing edge 24, and which are provided for inline injection of the fuel. The fuel, in this case gaseous fuel, is transported via the fuel gas feed 30 to the burner arrangement and then into this inner fuel tubing channel 36 and is subsequently distributed to the individual fuel nozzles 15 by means of branching off tubings 39. These branching of tubings are arranged essentially parallel to the main flow direction of the combustion gases. In the regions between the individual branching of tubings 39 between the two yet distanced opposite walls 37 there are located distancing elements 63.

The carrier air plenum 51 in the region facing the inner side of wall 37 is defined by a wall which is located essentially parallel to wall 37. Between these two walls there is an interspace 52 through which carrier air can flow. The distance between the two walls is established/maintained by distance keeping elements 53.

Also the walls of the inner fuel tubing 36, where facing the wall 37, are parallel but distanced from the outer wall structure 37 and again maintained in this distance by distance keeping element 53. Also in this interspace carrier air may flow.

The two channels 51 and 36 are also distanced from each other by interspace 55, which is also flown through by carrier air. The interspace between the walls 37 is, at the side opposite to the burner plate 18, closed by a bottom plate 59 which is arranged essentially parallel to the plate 18.

Above the burner plate 18 there is located a cavity 26, which on its bottom side faces the mixing chamber and on its upper side is bordered by an outer wall 19. The cavity 26 is furthermore circumferentially enclosed by a side wall 41.

Into this cavity 26 the fuel feed duct 30 is guided and then delivered to the inner fuel tubing, i.e. its longitudinal part 36. As three lances are combined in one such burner arrangement, there is one supply line 30 for the central lance and one further supply line 30' for the two outer lances, the gaseous fuel is distributed to the outer lances via individual distribution tubes 60. It is however also possible to have one single fuel feed which then distributes to all three fuel lances or to have individual fuel feeds for each fuel lance.

On its upper side the outer wall 19 is connected, via a flange 62, to a comparatively low pressure supply of carrier air, typically with a pressure in the range of 10-22 bar.

This carrier air, which is derived from the compressor stage of the corresponding necessary pressure without subsequent cooling, enters the cavity 26 via the carrier gas feed 31. It correspondingly cools the upper parts of the burner arrangement located within the cavity 26 so for example the fuel tubing 30 and distribution line 60. It then flows, as indicated by arrows 64, towards the burner plate 18. Distanced from the burner plate 18, according to this first embodiment, there is located a perforated plate 57 with holes 61 forming interspace 58 between the burner plate 18 and plate 57. The carrier air 65 penetrates these holes 61 and in a first cooling step cools the balcony 18 by impingement cooling and subsequent convective cooling. So after this impingement cooling it also cools the balcony by convective cooling because the carrier air is subsequently guided into the carrier air channel 51 from the top side as indicated schematically by arrows 72.

The carrier air then travels downwards towards the bottom part of the lance 22. As the wall of the carrier air plenum 51 is perforated at least where facing the leading edge 25, carrier air exits the channel 51 via these holes and cools the leading edge 25, specifically the inner side of the wall thereof, by impingement cooling.

Subsequent to this impingement cooling the carrier air travels downwards and backwards towards the trailing edge 24 of the lance and at the same time convectively cools the wall 37 as well as shields the inner fuel tubing 36 by travelling through interspaces 52, 55 and 38. One part of this carrier air (first fraction) travels towards the nozzles 15 and along the outer wall of the branching off tubings 39 to exit into the mixing chamber via the annular slots 71, such that a carrier air sleeve is enclosing the fuel jet 34 exiting, also in an annular fashion, a fuel exit slot defined by the inner side of the wall of 39 and a central element 50. So this first fraction of carrier air exits the injection device 22 taking the function of true carrier air for fuel injection.

A second fraction of this carrier air travels between the walls 37 across the distancing elements 63 and exits the injection device at its trailing edge 24 , where corresponding holes/slots are provided for effusion cooling.

A third fraction of this carrier air exits the injection device via vortex generators 23 which are located on the surface of the walls 37 upstream of the nozzles 15. To this end, these vortex generators 23 are provided with film cooling holes 32 through which, after having entered cavity 54, the carrier air penetrates into the mixing chamber.

In this case three lances 22 are combined within one burner arrangement, it is however also possible to have one burner with one lance or a burner arrangement with two lances or whichever is most appropriate for installation and/or maintenance purposes.

In somewhat more detail three bodies 22 arranged within an annular secondary combustion chamber are given in perspective view in figure 3, wherein the bodies are cut perpendicularly to the longitudinal axis 49 to show their interior structure.

In the cavity formed by the outer wall 37 of each body on the trailing side thereof there is located the longitudinal inner fuel tubing 36. It is distanced from the outer wall 37, wherein this distance is maintained by distance keeping elements 53 provided on the inner surface of the outer wall 37.

From this inner fuel tubing 36 the branching off tubing extends towards the trailing edge 29 of the body 22. The outer walls 37 at the position of these branching off tubings 39 is shaped such as to receive and enclose these branching off tubings 39 forming the actual fuel nozzles 15 with orifices located downstream of the trailing edge 29.

In the essentially cylindrically shaped interior of the branching off tubings 39 there is located a cylindrical central element 50 which leads to an annular stream of fuel gas. As between the wall of the branching off tubings 39 and the outer walls 37 at this position there is also an essentially annular interspace, this annular stream of fuel gas at the exit of the nozzle is enclosed by an essentially annular carrier gas stream.

Towards the leading edge 25 of the body 22 in the cavity formed by the outer wall 37 of the body in this embodiment there is located the carrier air tubing channel 51 extending essentially parallel to the longitudinal inner fuel tubing channel 36. Between the two channels 36 and 51 there is an interspace 55. The walls of the carrier air tubing channel 51 facing the outer walls 37 of the body 22 run essentially parallel thereto again distanced therefrom by distancing elements 53. In the walls of the carrier air tubing channel 51 there are provided cooling holes 56 through which carrier air travelling through channel 51 can penetrate. Air penetrating through these holes 56 impinges onto the inner side of the walls 37 leading to impingement cooling in addition to the convective cooling of the outer walls 37 in this region.

Within the walls 37 there are provided the vortex generators 23 in a manner such that within the vortex generators cavities 54 are formed which are fluidly connected to the carrier air feed. From this cavity the effusion/film cooling holes 32 are branching off for the cooling of the vortex generators 23. Depending on the exit point of these holes 32 they are inclined with respect to the plane of the surface at the point of exit in order to allow efficient film cooling effects.

Embodiment 2 (see figure 6):

The cooling of the lance balcony 18 is carried out as effusion cooling, which results in a lower pressure drop of the arrangement. After cooling the lance balcony 18 the cooling air is entering a carrier air plenum 51. The plenum 51 is equipped with several holes 56. These are chosen in diameter as such that a uniform distribution of the carrier air along the injectors is ensured. From the carrier plenum 51, the air impinges the leading edge 25 of the injectors. The air then cools the sidewall convectively. The cooling air is leaving the injector through various passages, e.g. three passages: This may be the large scale mixing devices 23 (e.g. vortex generators), the trailing edge 25 or annular slits at the injector holes. The split between each of the passages vortex generators, trailing edge and injector holes is adjusted to allow sufficient cooling of the components and a combustion behaviour as desired. Within each of the passages, the cross section is designed as such that the critical area is close to the exit of the passage, thus ensuring uniform cooling air distribution.

In this second embodiment there is no hole plate 57 separating the cavity 26 from the burner plate 18 and correspondingly there is no effusion/impingement cooling in the interspace 58. In this case the cavity 26 is directly adjacent to the structure of the burner plate 18, and the burner plate 18 is cooled by means of holes 66 provided in the burner plate 18, wherein typically these effusion/film cooling holes 66 are inclined with respect to the plane of the burner plate such that air exiting these effusion holes 60 is at an oblique angle with the main flow 40 leading to efficient film cooling on the surface of the plate 18. In this embodiment the cooling air 65 in the cavity 26 flows onto the inner surface of the burner plate 18 and a fraction thereof penetrates through the holes 66 for effusion cooling of the plate 18. This is normally only a minor fraction, the major fraction of the carrier air enters the carrier air plenum 51 under generation of a cooling air flow as indicated by arrow 67 in figure 6. It then penetrates through the holes 56 leading to impingement cooling of the inner side of the leading edge wall structure 25 of the lance. It then travels in the interspaces 52, 55 and 38 again towards the trailing edge and exits either as true carrier air for fuel injection as indicated by arrow 68 via the exits slots 71, or it exits via the trailing edge as indicated by arrow 69, or it exits, in a manner similar as illustrated in figure 2, via the effusion/film cooling holes 32 in the vortex generators 23.

LIST OF REFERENCE SIGNS burner 28 side surface of 23

mixing space, mixing zone 29 trailing edge of 23

burner wall 30 fuel gas feed

combustion space 31 carrier gas feed

outlet side, burner exit 32 film cooling holes

inlet side 33 lateral surface of 22 injection device, fuel lance 34 ejection direction of main flow from high -pressure fuel/carrier gas mixture turbine 35 central plane of 22 flow conditioning, turbine 36 inner fuel tubing, longitudinal outlet guide vanes part

vortex generators 37 outer wall of 22

fuel mass fraction contour at 38 interspace between 36 and 37 burner exit 5 39 branching off tubing of inner combustion chamber wall fuel tubing

transition between 3 and 12 40 transition region between 36 flow of oxidising medium and 39

fuel nozzle 41 sidewall

foot of 7 48 cross-sectional profile of 22 shaft of 7 49 longitudinal axis of 22 foot of 7 50 central element

shaft of 7 51 carrier air channel, carrier air burner plate, balcony plenum

outer wall 52 interspace between 37 and 51 tube forming 18 53 distance keeping elements streamlined body, lance 54 cavity within 23

vortex generator on 22 55 interspace between 51 and 36 trailing edge of 22 56 cooling holes

leading edge of 22 57 hole plate

cavity 58 interspace between 18 and 57 lateral surface of 23 59 bottom plate of 22 distribution tube 68 carrier air flow surrounding holes in 57 fuel jet

flange 69 cooling airflow at trailing distancing elements edge

bottom plate of 51 70 cooling airflow out of 23 cooling air in 26 71 annular slit of ejection device effusion holes in 18 72 carrier air flow entering the cooling airflow in 51 plenum 51 from interspace 58

Claims

1. Burner (1) for a combustion chamber of a turbine, with an injection device (7) for the introduction of at least one gaseous and/or liquid fuel into the burner (1), wherein the injection device (7) has at least one body (22) which is arranged in the burner (1) with at least two nozzles for introducing the at least one fuel into the burner (1), the at least one body being configured as a streamlined body (22) which has a streamlined cross-sectional profile (48) and which extends with a longitudinal direction (49) perpendicularly or at an inclination to a main flow direction (14) prevailing in the burner (1), wherein the body (22) has two lateral surfaces (33) essentially parallel to the main flow direction (14) joined at their upstream side by a leading edge portion (25) of the body (22) and joined at their downstream side forming a trailing edge (24), the at least two nozzles (15) being distributed along said trailing edge (24), wherein the body (22) comprises an enclosing outer wall (37) defining said streamlined cross-sectional profile (48), wherein within this outer wall (37), there is provided a longitudinal inner carrier air plenum (51) for the introduction of carrier air into the injection device (7), wherein the carrier air plenum (51) is provided with holes (56) such that carrier air exiting through these holes (56) impinges on the inner side of the leading edge portion (25) of the body (22).

2. Burner (1) according to claim 1, wherein the carrier air plenum (51) is a tubular duct located in the upstream portion of the cavity defined by the outer wall (37), wherein the wall of the tubular duct is distanced from the outer wall (37) leaving an interspace (52) in between for circulation of carrier air, wherein preferably the wall of the tubular duct in the region facing the outer wall (37) is running essentially parallel there to, and wherein further preferably the distance between the wall of the tubular duct and the outer wall (37) is established by at least one distance keeping elements (53) at the outer wall and/or at the wall of the tubular duct.

3. Burner (1) according to claim 2, wherein the carrier air plenum (51) extends essentially along the full length of the body (22) terminated by a bottom plate (64), which preferably is provided with holes (56) for cooling of a bottom plate (59) of the body (22).

4. Burner (1) according to any of the preceding claims, wherein air exiting from the carrier air plenum (51) is used as carrier air of the injection devices (7), wherein preferably the carrier air exits at the injection devices (7) via an annular slit (71) enclosing a central fuel jet, wherein further preferably the central fuel jet exits via an annular fuel slit.

5. Burner (1) according to any of the preceding claims, wherein within the enclosing outer wall (37) defining said streamlined cross- sectional profile (48), there is further provided a longitudinal inner fuel tubing (36) for the introduction of liquid and/or gaseous fuel, with branching off tubing (39) leading to the at least two nozzles (15), wherein preferably the carrier air plenum (51) is located in the upstream portion of the cavity defined by the outer wall (37) while the longitudinal inner fuel tubing (36) is located in the downstream portion of the cavity defined by the outer wall (37), wherein further preferably the wall of the carrier air plenum (51) is distanced from the wall of the longitudinal inner fuel tubing (36) for circulation of carrier air.

6. Burner according to claim 5, wherein the longitudinal inner fuel tubing (36) is circumferentially distanced from the outer wall (37), defining an interspace (38) for the delivery of carrier air to the at least one nozzle (15).

7. Burner (1) according to any of the preceding claims, wherein air exiting from the carrier air plenum (51) exits the injection device (7) via effusion holes, wherein preferably effusion holes are located at the trailing edge (24) of the injection device (7) and/or at the lateral surfaces (33) and/or at the leading edge (25) and/or at large scale mixing devices (23) of the injection device (7), the latter one is preferably in the form of vortex generators (23) located at the lateral surfaces (33) upstream of the nozzles (15).

8. Burner (1) according to any of the preceding claims, wherein the at least two nozzles (15) have their outlet orifices downstream of the trailing edge (24) of the streamlined body (22), wherein preferentially the distance (d) between the essentially straight trailing edge at the position of the nozzle (15), and the outlet orifice of said nozzle (15), measured along the main flow direction (14), is at least 2 mm, preferably at least 3 mm, more preferably in the range of 4-10 mm.

9. Burner as claimed in one of the preceding claims, wherein the streamlined body (22) has a cross-sectional profile (48) which is mirror symmetric with respect to the central plane (35) of the body (22).

10. Burner (1) according to any of the preceding claims, wherein at least one nozzle (15) is inclined with respect to the flow direction (14) and/or wherein the at least one nozzle (15) injects fuel and/or carrier gas at an inclination angle between 0- 30° with respect to the main flow direction (14).

11. Burner (1) according to any of the preceding claims, wherein within said longitudinal inner fuel tubing (36) provided for gaseous fuel there is provided a second inner fuel tubing for a second type of fuel, wherein preferably this second type of fuel is a liquid fuel and wherein further preferably gaseous fuel is delivered by the interspace between the walls of said longitudinal inner fuel tubing (36) and the walls of the second inner fuel tubing.

12. Burner as claimed in any one of the preceding claims, wherein upstream of the at least one nozzle (15) on at least one lateral surface (33) there is located at least one vortex generator (23), wherein preferably the vortex generator (23) has an attack angle in the range of 15-40° and/or a sweep angle in the range of 40-70°, wherein preferentially at least two nozzles (15) are arranged at different positions along said trailing edge (24), wherein upstream of each of these nozzles (15) at least one vortex generator (23) is located, and wherein preferably vortex generators (23) to adjacent nozzles (15) are located at opposite lateral surfaces (33), and wherein even more preferably more than three, most preferably at least four, nozzles (15) are arranged along said trailing edge (24) and vortex generators (23) alternatingly located at the two lateral surfaces (33) or wherein preferably downstream of each vortex generator (23) there are located at least two nozzles (15).

13. Burner (1) according to claim 12, wherein the vortex generator (23) is provided with cooling elements (32), wherein preferably these cooling elements (32) are effusion cooling holes provided in at least one of the surfaces (27, 28) of the vortex generator (23), and wherein even more preferably the film cooling holes (32) are fed with air from the carrier gas feed (31) also used for the fuel injection.

14. Burner (1) according to any of the preceding claims, wherein the streamlined body (22) extends across essentially the entire flow cross section between opposite walls (3) of the burner (1), wherein preferably the burner is an annular burner arranged circumferentially with respect to a turbine axis, and wherein between 10- 100 streamlined bodies, preferably between 40 - 80 streamlined bodies are arranged around the circumference, more preferably all of them being equally distributed along the circumference.

15. Burner (1) according to any of the preceding claims, wherein the fuel is injected from the nozzle (15) together with a carrier air stream which is supplied by the carrier air plenum (51), and wherein the carrier air is low pressure air with a pressure in the range of 10-22 bar, preferably in the range of 16-22 bar, and wherein preferably this carrier air is directly derived from a compressor stage without subsequent cooling.

16. Use of a burner (1) according to any of the preceding claims for the combustion under a high reactivity conditions, preferably for the combustion at high burner inlet temperatures and/or for the combustion of MBtu fuel with a calorific value of 5000-20,000 kJ/kg, preferably 7000-17,000 kJ/kg, more preferably 10,000- 15,000 kJ/kg, most preferably such a fuel comprising hydrogen gas.