WO2013139404A1 - Blade row for an unsteady axial flow gas turbine stage - Google Patents

Abstract

The current invention is directed to a blade row (50) for an axial flow gas turbine stage (100), comprising a plurality of blades (68) positioned in an annular array around an blade row axis (52), wherein a radial direction (56) runs perpendicular to said blade row axis (52), said plurality of blades (68) each radially extending from an inner diameter (62) to an outer diameter (64), wherein an annulus height (65, 66) of each blade (58) is defined between said inner diameter (62) and said outer diameter (64), wherein each blade (58) has a leading edge (86) and a trailing edge (88), said leading edge (86) of each blade (58) of the plurality of blades (68) being positioned in an inlet plane (90), and said trailing edge (88) of each blade (58) of the plurality of blades (68) being positioned in an outlet plane (92). Further, said inner diameter (62) and/or said outer diameter (64) of each blade (58) has a kink (72) so that said annulus height (65, 66) of each blade (58) increases between said inlet plane (90) and said outlet plane (92).

Description

Blade row for an unsteady axial flow gas turbine stage

[0001] The present invention relates to a blade row for an axial flow gas turbine stage, comprising a plurality of blades positioned in an annular array around a blade row axis, wherein a radial direction runs perpendicular to said blade row axis, said plurality of blades each radially extending from an inner diameter to an outer diameter, wherein an annulus height is defined between said inner diameter and said outer diameter, wherein each blade has a leading edge and a trailing edge, said leading edge of each blade of the plurality of blades being positioned in an inlet plane, and said trailing edge of each blade of said plurality of blades being position in an outlet plane.

[0002] Such a blade row is, for example known from document EP 2 196 626 [0003] Blade rows for gas turbine stages are commonly known in the art. Such blade rows are used as stators and rotors of gas turbines. The gas turbines might be used stationary, for example in power plants, or might be part of a gas turbine engine, for example of an aircraft.

[0004] Usually, such gas turbines are operated in a steady mode which means the working fluid of the gas turbine flows steadily through the gas turbine and almost without any time dependence. In particular, commonly known gas turbines use steady combustors. In a steady combustor, a steady fluid flow enters the combustor, is mixed with an injected fuel and burned so that a steady flow of heated up working fluid leaves the combustor.

[0005] However, such steady combustors suffer a pressure drop in the working fluid as it runs through the combustor. Although they can be ideally modelled with a constant pressure, a pressure drop over the combustor is always present in a real gas turbine.

[0006] Hence, approaches have been made to implement unsteady combustion into gas turbines. While first approaches to building such gas turbines have been made already over 100 years ago, further efforts have been made in the past decade to implement unsteady combustion in stationary gas turbines and gas turbine engines. Such combustors are commonly referred to as "pressure-rise combustors" as they are able to even raise the pressure of the working fluid over the combustor. Types of such combustors include pulsejet combustors that apply a combustion occurring in pulses. Inlet valves and outlet valves may be present to form a combustion chamber in which the pulsed combustion occurs. However, so-called valve-less pulsejet combustors also exist that do not make use of any valves but only of the pulse combustion. Further, recent developments have turned to a so-called pulse-detonation combustor that make use of detonation waves during the combustion of the fuel and working fluid mixture. Theoretically, such pulse detonation can be used in engines also for hypersonic flight speeds. Further, the ideal thermodynamic efficiency of such a pulse-detonation engine is higher than that of commonly known turbo engines working with a steady combustion. [0007] An example for a relatively old gas turbine working with an unsteady combustion is given in document US 1 ,900,558 A. Examples for pulse detonation engines are shown in documents US 2009/0165438 A1 and US 2010/0186370 A1.

[0008] A specific issue of working with detonation shock waves in gas turbines is the propagation of the Shock wave downstream through the turbine stages. Is has been found that the shock wave impinging on the stator of the first turbine stage, the nozzle guide vanes, is reflected off the nozzle guide vanes to a high amount. Only a small portion of the working fluid travels through the nozzle guide vanes which significantly reduces the mass flow through the turbine stages, and, hence the overall efficiency of the gas turbine.

[0009] Therefore, several attempts have been made to improve the downstream travelling of the working fluid ejected from a pressure-rise combustor through the first turbine stage.

[0010] For example, document US 2008/0155959 A1 shows a transition piece for use with a gas turbine engine providing a path between the exhaust from one or more pressure-rise combustors and a downstream turbine for the extraction of work from the exhaust flow. The transition piece provides a non-expanding path for the exhaust flow through the transition piece and directs the flow so as to be effective in driving the turbine when it reaches the end of the transition piece.

[0011] Further, document US 2003/0182927 A1 shows a Shock wave reflector including a number of reflective units positioned along a longitudinal direction and separated by a gap. The detonation chamber includes a number of these reflective units formed in the wall and positioned along the longitudinal direction.

[0012] Further, document US 2007/0015099 A1 shows a pulse detonation device for dividing a pulse detonation shock wave into a primary and a control portion to reduce the strength of a propagating Shock wave and/or change its direction. The device contains a flow separator which directs a portion of the Shock wave into itself, thus reducing the Shock waves strength. [0013] Further approaches suggested altering the first gas turbine stage.

Hence, document WO 2011/150979 A1 suggest a method and device for alleviating a transonic gas turbine stage comprising a turbine nozzle with guide vanes and a rotor with rotor blades. It is suggested to direct an unsteady cooling fluid jet from an outlet at the surface of the guide vane at or near the trailing edge of the guide vanes. This unsteady cooling jet shall be synchronized with the rotor blade passing frequency so as to have a flow-rate minimum when a rotor blade is within an area around a shock originating at said guide vane.

[0014] Further, initially introduced document EP 2 196 626 A2 suggests a gas turbine engine having a pressure-rise combustor positioned upstream of a stage of a turbine nozzle guide vanes. The vanes of the stage of the turbine nozzle guide vanes form an ejector and each of the vanes has an upstream portion and downstream portion. The upstream portions of the vanes have leading edges and the upstream portions are arranged substantially straight and parallel to define mixing passages for a flow of gases therethrough. The downstream portions of the vanes are arranged at an angle to the upstream portions of the vanes to turn the flow of gases therethrough.

[0015] However, there is still a need in the art for an improved gas turbines working with an unsteady combustion and having transonic Shock waves travelling downstream from the combustor through the turbine stages.

[0016] Hence, it is an object of the current invention to improve the mass flow of a working fluid ejected from a pressure-rise combustor through the turbine stages of a gas turbine.

[0017] Therefore, according to a first aspect of the invention, there is provided a blade row for an axial flow gas turbine stage comprising a plurality of blades positioned in an annular array around a blade row axis, wherein a radial direction runs perpendicular to said blade row axis, said plurality of blades each radially extending from an inner diameter to an outer diameter, wherein an annulus height of each blade is defined between said inner diameter and said outer diameter, wherein each blade has a leading edge and a trailing edge, said leading edge of each blade of the plurality of blades being positioned in an inlet plane, and said trailing edge of each blade of the plurality of blades being positioned in an outlet plane, wherein said inner diameter and/or said outer diameter of each blade has a kink so that said annulus height of each blade increases between said inlet plane and said outlet plane.

[0018] It has been found that focusing on the shock waves only to improve the overall mass flow through a turbine stage is not sufficient to show good overall results. Instead, both the transonic shock wave and the fluid travelling with a speed of about Mach 1 directly behind the shock wave have to be taken into account.

[0019] Usually, in a steady flow, a rapid increase in the annulus height would cause a large scale flow separation to occur. This separation would provide significant aerodynamic blockage in the turbine stage plane and the swallowing capacity of the blade row will be kept low by this mechanism in steady flows. However, the rapidly increasing annulus height is introduced to better deal with the highly unsteady situation if the turbine stage is coupled a pressure-rise combustor were supersonic compression waves propagate from upstream and interact with the blade row. Such Shock waves at a speed of about Mach 1 , 5 to 2, or even detonation waves travelling at a speed of about Mach 5 to 7, however, will travel more effectively through a blade row improved according to the present invention.

[0020] In a situation with propagating shock waves, the flow behind the shock wave is at considerably higher Mach number than it would be present in a steady case. Without the increasing annulus height, this fluid cannot pass through the throat of the blade row. The result is a shock reflection passing backwards out of the turbine and into the combustion chamber. Very little of the unsteady energy is captured as available kinetic energy to drive the downstream turbine rotors. However, it has been found that with the increasing annulus height, the high Mach number flow behind the Shock waves, further accelerated by the narrowing passage will have reached the sonic condition before reaching the portion of the blade row where the annulus height increases. But as a sonic flow this part of fluid will see the increasing annulus height as a rapid expansion and the fluid will accelerate to show the increasing error. The capacity of the blade row will in- crease momentarily and more of the high-pressure air will be allowed through. This weakens the strength of the wave reflected back into the combustion system.

[0021] During the transient condition where a high Mach number fluid is flowing through the blade row with an increasing annulus height, the efficiency of the blade row will be much improved compared to the steady case. This is because what was a diffusing mixing process becomes a simple supersonic acceleration.

[0022] In the current context, the term "kink" is to be understood in a way that a function of the inner and outer diameters, respectively, is not continuously differentiable at the "kink". This means that the annulus height increases rapidly at a certain point between the inlet plane and the out plane, namely at the height of the kink. At this point, the outer diameter and/or the inner diameter of each blade of the blade row has a kink, which may also be referred to as an "edge" or "corner". In this context, "kink", "edge" or "corner" means a region of ideally an infinite wall curvature, either positive or negative. In other words, ideally a zero radius of curvature. In practice, a slight rounding of the kink is acceptable, for example with a radius of curvature of less than one millimetre. This is still to be considered a "kink" according to the current invention.

[0023] According to a second aspect of the invention, there is provided a stator for an axial flow gas turbine stage comprising a blade row according to the first aspect or one of its embodiments.

[0024] According to a third aspect of the invention, there is provided an axial flow gas turbine stage comprising a stator according to the second aspect of the invention and further comprising a rotor having a plurality of rotor blades positioned in an annular array around a rotor axis informing a rotor blade row, wherein the rotor axis runs coaxially with said blade row axis, and wherein the rotor blade row is a blade row according to the first aspect or one of its embodiments.

[0025] According to a fourth aspect of the invention, there is provided a gas turbine engine, having at least one pressure-rise combustor, said at least one pressure-rise combustor being positioned upstream of at least one axial flow gas turbine stage, wherein at least one of said at least one axial flow gas turbine stage has a stator having a blade row according to the first aspect or one of its embodiments and/or a rotor having a blade row according to the first aspect or one of its embodiments.

[0026] The stator according to a second aspect, the axial flow gas turbine stage according to the third aspect and the gas turbine engine according to the fourth aspect all comprise a blade row according to the first aspect and, therefore, provide the same advantages.

[0027] Hence, the initially laid out objective is achieved completely.

[0028] In a further embodiment of the blade row according to the first aspect, said inner diameter is constant and said outer diameter increases between said inlet plane and said outlet plane, wherein said outer diameter is kinked radially outwards.

[0029] By this, the inner diameter and hence, a shaft diameter, whereupon the blades are mounted, can be kept constant. Hence, a design process can be significantly facilitated.

[0030] Further, according to an embodiment of the blade row according to the first aspect, said outer diameter is constant and said inner diameter decreases between said inlet plane and said outlet plane, wherein said inner diameter is kinked radially inwards.

[0031] Of course, this embodiment might be more difficult to put into practice as more space at the inner circumference of the blade row is not readily available. However, under certain circumstances, an embodiment where the outer diameter of the whole turbine stage has to be kept constant may occur and, then, this embodiment may provide an advantage. [0032] According to a further embodiment of the blade row according to the first aspect of the invention, said inner diameter decreases and said outer diameter increases between said inlet plane and said outlet plane, wherein said outer diameter is kinked radially outwards, and wherein said inner diameter is kinked radially inwards.

[0033] Hence, both the foregoing embodiments may be combined to increase the overall annulus height of the blade row.

[0034] According to a further embodiment of the blade row according to the first aspect, said annulus height increases in a step.

[0035] Hence, the annulus height increases suddenly. Therefore, a very abrupt expansion of the annulus height and of the cross-sectional area the working fluid flows through occurs. This may result in a flow separation and vortices might occur right behind the step radially outwards. However, the swallowing capacity of the turbine stage is significantly increased.

[0036] In an embodiment, said kink of said step runs perpendicularly to said blade row axis.

[0037] Such an increase of the annulus height is an easy design measure which can be implemented conveniently.

[0038] However, according to a further embodiment, said kink of said step runs parallel to a throat plane of two adjacent blades of said plurality of blades.

[0039] The "throat plane" between two adjacent blades is the plane with the smallest cross-sectional area the flow has to pass through. In case the edge of the step runs parallel to the throat plane, a pressure difference between a pressure side end of the step and a suction side end of the step may be reduced. Hence, a migration of the working fluid in the separated zone behind the step towards the suction side may be reduced. [0040] According to a further embodiment of the blade row according to the first aspect of the invention, the annulus height increases linearly from a smaller annulus height to a larger annulus height over an increase portion, wherein said increase portion starts in an upstream kink. Further, the increase portion may end in a downstream kink.

[0041] Hence, the annulus height does not need to increase suddenly in a step. It may also increase linearly from a smaller annulus height to a larger annulus height. Further, other transitions from the smaller annulus height to the larger annulus height might be implemented that are non-linear.

[0042] Such a linear increase in the annulus height might be beneficial as this will reduce the losses due to the expansion of the annulus height by reducing the scale of the resulting flow separation.

[0043] In an embodiment, said annulus height increases linearly in an increase direction, wherein said increase direction runs perpendicularly to said blade row axis.

[0044] Again, such design may easily be implemented and may be calculated more easily during a layout phase of the turbine stage.

[0045] However, said annulus height might also increase linearly in an increase direction, wherein said increase direction runs perpendicularly to a throat plane of two adjacent blades.

[0046] Again, this might reduce the strength of the migration of fluid towards the suction side.

[0047] According to a further embodiment of the blade row according to the first aspect, a cross-sectional annular area of said blade row in said outlet plane is at least 1.5 times the cross-sectional annular area of the inlet plane. Preferably, the cross-sectional annular area of said blade row in said outlet plane is 2.0 times, more preferably 2.5 times, even more preferably 3.0 times the cross-sectional annular area of the inlet plane. In particular, the cross-sectional annular area of said blade row in said outlet plane is 2.0 times the cross-sectional annular area of the inlet plane.

[0048] Test results have shown that the cross-sectional annular area of the blade row should expand by at least 50% to sufficiently support the mass flow through the blade row. The corresponding increase in the annulus height, i.e. the increase of the outer diameter and/or the decrease of the inner diameter, has to be calculated accordingly based on the predetermined increase in cross-sectional annular area and the initial inner and outer diameters, i.e. the inner and outer diameters in the inlet plane.

[0049] In a further embodiment of the gas turbine engine according to the fourth aspect, the working fluid of the gas turbine engine is ejected from at least one pressure- rise combustor onto a first axial flow gas turbine stage of said at least one axial flow gas turbine stage, wherein said first axial flow gas turbine stage has a stator having a blade row according to the first aspect of the invention or one of its embodiments and/or a rotor having a blade row according to the first aspect of the invention or one of its embodiments.

[0050] In such an embodiment, the annulus height is not only increased in the stator but also in the rotor of a turbine stage. This may be useful if there is still a strong Shock wave in the outlet plane of the stator. Further, it may be in particular the first gas turbine stage, i.e. the turbine stage the working fluid ejected from the pressure-rise combustor hits onto, that is designed according to the teachings of the current invention. Hence, the stator of the first turbine stage, i.e. the nozzle guide vanes (NGV) may be designed according to the teachings of the current invention specifically.

[0051] In a further embodiment of the gas turbine engine according to the first aspect of the invention, said at least one pressure-rise combustor is a pulsejet combustor, a valveless pulsejet combustor or a pulse-detonation combustor.

[0052] Hence, the current invention may be implemented in combination with any pressure-rise combustor. [0053] Further, the gas turbine engine according to the fourth aspect of the invention may be a turbo propeller gas turbine engine, a turbofan gas turbine engine, a turbojet gas turbine engine or a turbo shaft gas turbine engine.

[0054] In a turbo propeller gas turbine engine, the turbine shaft is coupled to a propeller that drives the aircraft. In a turbojet gas turbine engine, the impulse driving the aircraft is solely generated by the working fluid flow out of the turbines. In a turbofan gas turbine engine, the impulse is generated partly by the turbojet exiting the turbine stages and partly by the flow accelerated by the fan stage. Further, the invention may be used in a turbo shaft gas turbine engine, wherein almost the whole energy of the working fluid is transmitted onto an output shaft. Such an engine may be used, for example, in a helicopter. Further, of course, the current invention may also be used in stationary gas turbines, for example, in power plants.

[0055] It goes without saying that the features mentioned above and those yet to be explained below can be used not only in the combination respectively indicated, but also in other combinations or by themselves, without departing from the scope of the present invention.

[0056] Exemplary embodiments of the invention are illustrated in the drawing and are explained in greater detail on the following description. In the figures:

Fig. 1 shows an embodiment of a gas turbine engine;

Fig. 2a shows an embodiment of a blade row in a cross-sectional view;

Fig. 2b shows a front view on a blade row designed according to the present invention;

Fig. 3 shows a schematic sectional view along the line Ill-Ill in Fig. 2b; Fig. 4 shows a cross-sectional view in a circumferential direction through adjacent blades of a blade row designed according to the current invention;

Fig. 5 shows a schematic cross-sectional view of a further embodiment of a blade row;

Fig. 6 shows a schematic cross-sectional view of a turbine stage according to the current invention;

Fig. 7 shows a schematic cross-sectional view of the third embodiment of a turbine stage;

Fig. 8 shows a cross-sectional view in a circumferential direction of two adjacent blades of a blade row designed according to a further embodiment of the current invention;

Fig. 9 shows an example embodiment underlying the results in Figs. 10 and

1 1 ;

Fig. 10 shows a comparison of the mass flow of embodiments with and without a step; and

Fig. 11 shows a corresponding comparison of the total enthalpy with and without a step.

[0057] Fig. 1 shows an embodiment of a gas turbine engine 10. The depicted gas turbine engine 10 is merely of exemplary nature and should not be considered as delimiting. The gas turbine engine 10 is shown as a turbofan gas turbine engine. However, the current invention may also be implemented in other types of gas turbine engines including, for example, a turbo propeller gas turbine engine, a turbojet gas turbine engine or a turbo shaft gas turbine engine. Further, it may also be implemented in a stationary gas turbine. [0058] The gas turbine engine 10 generates an impulse between an incoming fluid 12 and an outgoing fluid 14 and, by this, drives an aircraft to which it is attached (not shown).

[0059] The gas turbine engine 10 is rotationally symmetrical to a rotational axis 16. Hence, only a part of the gas turbine engine 10 is shown.

[0060] The gas turbine engine 10 can be divided into different stages explained in the following. The gas turbine engine 10 has a first compressor stage 18 that comprises a fan stage 20. The first compressor stage 18 is schematically shown for illustrative purposes only. The first compressor stage 18 may itself include several substages.

However, all of these substages run with a certain rotational frequency. The fan stage 20 has a substantially greater diameter than the rest of the gas turbine engine 10. This is to accelerate a part of the incoming fluid 12 directly in a propeller-like manner to generate a higher overall impulse.

[0061] Further, the gas turbine engine 10 comprises a second compressor stage 22. However, the gas turbine engine 10 implementing the current invention may also comprise only one compressor stage or even three or more compressor stages.

[0062] Further, the gas turbine engine 10 comprises a combustor stage 24. Downstream of the combustor stage 24, there is positioned a first turbine stage 26 and a second turbine stage 28. The fluid exiting the second turbine stage 28 is accelerated in an outlet nozzle 30. The outlet nozzle 30 and the depicted embodiment is merely of convergent nature. However, the outlet nozzle 30 may also be of a convergent / divergent nature.

[0063] The first compressor stage 18 and the second turbine stage 28 are connected by a first shaft 32. Hence, the rotational frequency of the second turbine stage and the first compressor stage is the same. Further, the second compressor stage 22 and the first turbine stage 44 are connected via a second shaft 34. Hence, the rotational frequency of the second compressor state 22 and the first turbine stage 44 are the same. [0064] The impulse generated by the gas turbine engine 10 is generated by heating up the incoming fluid 12 in the combustor 24 after it has been compressed by the first and second compressor stages 18 and 22. Hence, the part of the incoming fluid 12 travelling through the combustor 24 and the turbine stages 26, 28 has a significantly higher kinetic energy when leaving the outlet nozzle 30. Further, a part of the energy is used to drive the first and second compressor stages 18, 22 and, the fan stage 20. Hence, a further part of the incoming fluid 12 is directly accelerated via the fan stage 20 and, therefore, further accelerated by a fan nozzle 36. Although the incoming fluid 12 is not accelerated by the fan stage 20 as much as the fluid travelling through the combustor stage 24, it nonetheless forms a significant part of the overall impulse generated by the gas turbine engine 10 due to the significantly higher mass flow through the fan stage 20 compared to the mass flow through the combustor stage 24.

[0065] The combustor stage 24 is a so-called pressure-rise-combustor. It may have a valve arrangement 38 to operate the pressure-rise combustor. However, also a valveless arrangement without a valve arrangement 38 is possible. Hence, the pressure- rise combustor 40 might be of any type, for example a pulsejet combustor, a valveless pulsejet combustor or even a pulse detonation combustor. Usually, a plurality of combus- tors 40 are arranged around the circumference of the gas turbine engine 10. However, it may also be one annular combustor 40.

[0066] After leaving the at least one combustor 40, the unsteady working fluid may travel through a transition track 42 to have a more steady flow hitting on the nozzle guide vanes of a very first turbine stage 44 within the first turbine stage 26. This very first turbine stage 44 comprises a blade row arranged in a fixed manner relative to the rotational axis 16, i.e. a stator also called the nozzle guide vanes. It deflects the working fluid ejected from the at least one combustor 40 and travelling in a direction more or less parallel to the rotational axis 16. Downstream of the nozzle guide vanes, there is positioned a first rotor that comprises a blade row that is then driven by the working fluid deflected by the nozzle guide vanes.

[0067] Fig. 2a shows a schematical cross-sectional view through a blade row 50 that may form the nozzle guide vanes of the gas turbine engine depicted in Fig. 1. [0068] The blade row is generally indicated with reference numeral 50. The blade row 50 is arranged in an annular array around a blade row axis 52 that runs coaxi- ally with the rotational axis 16 of the gas turbine engine 10. The dimensions in Fig. 2a are for illustrative purposes only. They should not be considered as realistic proportions of the blade row according to the invention.

[0069] An axial direction is indicated by arrow 54. The axial direction 54 runs parallel to the blade row axis 52. Perpendicular to the axial direction 52 is a radial direction 56.

[0070] In the view in Fig. 2a, the cross section is cut through a single blade 58 of the blade row 50. The blade 58 has an inner diameter 62 and an outer diameter 64, 64'.

[0071] Radially inwards of the inner diameter 62, there is a support element 63 that may be part of the first shaft 32 or at least coupled to the first shaft 32. A working fluid 60 ejected from the combustor stage 24 cannot pass through this support element 63. Hence, the working fluid 60 is forced to pass through the annular array of blades 58.

[0072] Between the inner diameter 62 and the outer diameter 64, 64', there is defined an annulus height 65, 66 of the blade 58.

[0073] As it is shown in Fig. 2a, the annulus height 65, 66 of the blade 58 increases. An upstream portion of the blade 58 has a smaller outer diameter 64. A downstream part of the blade has a larger outer diameter 64'. Hence, the blade 58 has an upstream portion with a smaller annulus height 65 and a downstream portion with a larger annulus height 66. As explained above, this increases a mass flow through the first turbine stage in case of an highly unsteady flow ejected from the combustor stage 24 as in case of a pressure-rise combustor 40.

[0074] In the embodiment depicted in Fig. 2a, the smaller annulus height 65 increases to the larger annulus height 66 in a step 70. Further, a kink 72 of the step 70 runs perpendicularly to the blade row axis 52. Hence, the annulus height 65 increases with the kink 72. It is not necessary that the kink 72 is a rapid kink of the outer diameter 64 by 90° radially outwards. In case the annulus height 65 increases in a step 70 to the larger annulus height 65, however, this will be the case. Ideally, at the kink 72, there is an infinite wall curvature. Hence, the kink 72 might also be referred to as a "corner" or an "edge". However, a wall curvature of less than 1 millimetre shall still be regarded as a kink in the context of the current application. At the end of the step 70, there is a second kink 73 at the larger outer diameter 64'. However, the second kink 73 may also be omitted and substituted by a smooth transition toward the larger annulus height 65.

[0075] Fig. 2b shows a schematic front view on the blade row 50. The blade row 50 comprises a plurality of blades 58, 58', 58" etc. arranged in an annular array around the blade row axis 52. For illustrative purposes, the blades are shown in a non- overlapping manner. Of course, the individual blades 58, 58', 58" would circumferentially overlap in a real blade row. In Fig. 2b, like elements are indicated by the same reference numerals as in Fig. 2a and will not be explained again.

[0076] As the annulus height 65, 66 increases in the step 70, the working fluid 60 entering into the blade row 50 can only enter through an inner ring 74 having a smaller outer diameter 64. Only after having travelled past the step 70, the working fluid 60 may expand into an outer ring 76 of the larger outer diameter 64'. Hence, every blade 58, 58', 58" has an inner section 78 between the inner diameter 62 and the smaller outer diameter 64 and an outer section 80 between the smaller outer diameter 64 and the larger outer diameter 64'. The annular area in the radial direction 56 between the inner diameter 62 and the outer diameters 64 and 64', respectively, defines the cross-sectional annular area in the context of the current application.

[0077] Fig. 3 shows a schematic cross-sectional view along line Ill-Ill of Fig. 2b.

[0078] The blade row is not shown again in detail as in Fig. 2a but only as a schematical view. As shown, a stator 84 is provided by the plurality of blade 68 in an annular array around the blade row axis 52. Radially outwards, a housing 82 of the gas turbine stage is adapted to the increasing outer diameter of the plurality of blade 60 so that it matches the plurality of blade 68 and prevents the working fluid 60 from travelling around the blades.

[0079] Fig. 4 shows a schematic cross-sectional view in a radial direction 56 of a blade row 50. Two adjacent blade rows 58 and 58' are shown.

[0080] Each blade 58 has a leading edge 86 upstream and a trailing edge 88 downstream. All leading edges 86 lie in an inlet plane 90 and the trailing edges 88 lie in an outlet plane 92. It is part of the invention that the annulus height 65, 66 increases between this inlet plane and that outlet plane. Hence, it has to be emphasized that the annulus height does not increase between two blade rows, but the annulus height of the plurality of blades 68 of a single blade row 50 increases.

[0081] Further, Fig. 4 shows a so-called throat plane 94 between the two adjacent plates 58, 58'. The throat plane 94 is the cross-sectional area between the two adjacent plates 58, 58' which is the smallest.

[0082] In the embodiment depicted in Fig. 4, the step 70 runs perpendicularly to the axial direction 54 and blade row axis 52.

[0083] Fig. 5 shows a further embodiment of a blade row 50 and a correspondingly designed stator 84. In this embodiment, the annulus height 65, 66 increases linearly from a small annulus height 65 to a large annulus height 66. Hence, the outer diameter has a conical portion 98 between the inlet plane 90 and the outlet plane 92. An increased direction 96 along which the conical portion 98 extends runs parallel to the axial direction 54.

[0084] As laid out above, it is not necessary that the kink 72 is a sharp bend of the outer diameter 64 by 90° radially outwards. In case of the embodiment shown in Fig. 5, the kink 72 of the outer diameter 64 is less than 90°. However, it is still the case that there is - ideally - an infinite wall curvature right at the kink 72 and, hence, a "corner" or an "edge". Then, the smaller outer diameter 64 increase over an increase portion 95 to said larger outer diameter 64'. The increase in the cross-sectional annular area shall be at least 50%, preferably 100%, over the increase portion to provide a sufficiently abrupt increase in the cross-sectional annular area. The increase portion 95 then starts in the upstream kink 72 and ends in a downstream kink 73. Alternatively, there my also be a smooth transition from the increase portion 95 to the outer diameter 64'.

[0085] Fig. 6 shows a whole turbine stage comprising a stator 84 and a rotor 102. The rotor 102 rotates about a rotor axis 104 and has a plurality of rotor blades 106. In the embodiment depicted in Fig. 4, only the blade row of the stator 84 is designed according to the current invention and the annulus height of the rotor 102 is kept constant.

[0086] However, as shown in Fig. 7, a further embodiment of a turbine stage 100 may have not only the annulus height of the blade row of the stator 84 increasing but also the annulus height of the rotor blades 106. Hence, the rotor 102 may include a further rotor step 108 or linearly increasing outer diameter.

[0087] Fig. 8 shows a further embodiment of two adjacent blade rows 58, 58' of a further embodiment of a blade row 50 like elements as in Fig. 4 are indicated by the same reference numerals and will not be explained again.

[0088] Compared to Fig. 4, that shows the steps 70 extending perpendicularly to the axial direction 54, the increase in the annulus height 65, 66, for example step 70 may extend parallel to throat plane 94 between the two adjacent blades 58, 58'.

[0089] Hence, this might lead to the step 70 running more or less in a zig-zag- fashion circumferentially around the blade row 50.

[0090] Tests have been conducted to validate the effect of the increasing annulus height concerning the mass flow through the blade row. An example embodiment for the tests is shown in Fig. 9. The annulus height increases in a step running perpendicularly to the blade row axis. A start plane 1 , an end plane 5, a plane upstream of the blade 58 between the start plane 1 and the blade 58 and an approximated rotor plane 4 are observed. The rotor plane 4 is considered to be at one third of a blade chord after the trailing edge 88. The step 70 lies in a plane 3. Comparative tests have been conducted with and without the step 70.

[0091] Three combustion cycles were simulated. The post treatment is only based on the third cycle. In order to compare the two geometries, the physical values in the planes are non-dimensionalized by the mass flow. In each plane, the velocity V, the density p, the pressure P, the temperature T, the area of the plane A and the Mach number M are known. The cross sectional annular area increases by 100%. The interested values are in each planes integrated and have then been averaged in time based on the following formulas:

, ^VC*!i-

_ 1 1

Average mass flow Mach number Af =— I — > pV_xA_CeuM

1 7ft ■ '

Average mass flow Total Enthalpy H_a =

[0092] The test showed the following results:

Table 1 : Average values for geometry with step

Plane 1 Plane 5 Plane 2 Plane 4

5,1791 kg/s 5,1733 kg/s 5,0136 kg/s 5,1517 kg/s

P 50,654 bar 40,22 bar 51 ,591 bar 39,311 bar

M 0,3428 1 ,0137 0,2938 0,9723

ΊΓ_β 2,5045e+6 J/kg 2,7417e+6 J/kg 2,1593e+6 J/kg 2,2989e+6 J/kg Table 2: Average values for a nominal geometry

[0093] The tests have shown that the mass flow can be increased by about 20% compared to a geometry without any step.

[0094] Fig. 10 shows a comparison of the mass flow of embodiments with and without the step 70 over a cycle time. Fig. 1 1 shows a corresponding comparison of the total enthalpy.

[0095] As can be seen, on one cycle, the mass flow reverses during about 20% of one cycle's time. Without the step 70, the flow reverses a bit longer. Observing the values upstream the blade in plane 2, the peak of the mass flow which is characterizing for an incoming Shock wave is larger in the stepped geometry and the minimum is also larger. Hence, more mass flow can travel through the blade row 50 having a step 70. This conclusion is also visible on the exit values of mass flow, cf. the lines related to plane 4. Therefore, the step 70 clearly increases the mass flow.

[0096] As Fig. 10 shows an improved mass flow, looking on the total enthalpy plot in Fig. 1 1 shows that the efficiency of the step 70 does not decrease. The amount of energy coming in the blade row 50 is almost the same with or without step. However, the peak of energy in plane 4 is higher without step 70. But, with step 70, the peak is more diffused. The integrals of the plots show that, indeed, more total energy goes through the blade row 50 with the step, approximately in an order of 1 %.

Claims

Patent claims

1. Blade row (50) for an axial flow gas turbine stage (100), comprising a plurality of blades (68) positioned in an annular array around a blade row axis (52), wherein a radial direction (56) runs perpendicular to said blade row axis (52), said plurality of blades (68) each radially extending from an inner diameter (62) to an outer diameter (64), wherein an annulus height (65, 66) of each blade (58) is defined between said inner diameter (62) and said outer diameter (64), wherein each blade (58) has a leading edge (86) and a trailing edge (88), said leading edge (86) of each blade (58) of the plurality of blades (68) being positioned in an inlet plane (90), and said trailing edge (88) of each blade (58) of the plurality of blades (68) being positioned in an outlet plane (92), characterized in that said inner diameter (62) and/or said outer diameter (64) of each blade (58) has a kink (72) so that said annulus height (65, 66) of each blade (58) increases between said inlet plane (90) and said outlet plane (92).

2. Blade row according to claim 1 , characterized in that said inner diameter (62) is constant and said outer diameter (64) increases between said inlet plane (90) and said outlet plane (92), wherein said outer diameter (64) is kinked radially outwards.

3. Blade row according to claim 1 , characterized in that said outer diameter (64) is constant and said inner diameter (62) decreases between said inlet plane (90) and said outlet plane (92), wherein said inner diameter (62) in kinked radially inwards.

4. Blade row according to claim 1 , characterized in that said inner diameter (62)

decreases and said outer diameter (64) increases between said inlet plane (90) and said outlet plane (92), wherein said outer diameter (64) is kinked radially outwards, and wherein said inner diameter (62) is kinked radially inwards.

5. Blade row according to any of claims 1 to 4, characterized in that said annulus height (65, 66) increases in a step (70).

6. Blade row according to claim 5, characterized in that said kink (72) of said step (70) runs perpendicularly to said blade row axis (52).

7. Blade row according to claim 5, characterized in that said kink (72) of said step (70) runs parallel to a throat plane (94) of two adjacent blades (58, 58') of said plurality of blades (68).

8. Blade row according to any of claims 1 to 4, characterized in that the annulus height (65, 66) increases linearly from a smaller annulus height (65) to a larger annulus height (66) over an increase portion, wherein said increase portion (95) starts in an upstream kink (72).

9. Blade row according to claim 8, characterized in that said increase portion (95) ends in a downstream kink (73).

10. Blade row according to claim 8 or 9, characterized in that said annulus height (65, 66) increases linearly in an increase direction (99), wherein said increase direction (99) runs perpendicularly to said blade row axis (52).

1 1. Blade row according to claim 8 or 9, characterized in that said annulus height (65, 66) increases linearly in an increase direction (99), wherein said increase direction (99) runs perpendicularly to a throat plane (94) of two adjacent blades (58, 58').

12. Blade row according to any of claims 1 to 11 , characterized in that a cross- sectional annular area of said blade row (50) in said outlet plane (92) is at least 1.5 times the cross-sectional annular area of the inlet plane (90).

13. Stator (84) for an axial flow gas turbine stage (100) comprising a blade row (50) according to any of claims 1 to 12.

14. Axial flow gas turbine stage (100) comprising a stator (84) according to claim 11 and further comprising a rotor (102) having a plurality of rotor blades (106) posi- tioned in an annular array around a rotor axis (104) and forming a rotor blade row (106), wherein the rotor axis (104) runs coaxially with said blade row axis (52), characterized in that the rotor blade row (106) is a blade row according to any of claims 1 to 12.

15. Gas turbine engine (10), having at least one pressure-rise combustor (40), said at least one pressure-rise combustor (40) being positioned upstream of at least one axial flow gas turbine stage (26, 28, 100), characterized in that at least one (100) of said at least one axial flow gas turbine stage (26, 28, 100) has a stator (84) having a blade row (50) according to any of claims 1 to 1 1 and/or a rotor (102) having a blade row (50) according to any of claims 1 to 12.

16. Gas turbine engine according to claim 13, characterized in that a working fluid (60) of the gas turbine engine (10) is ejected from at least one pressure-rise combustor (40) onto a first axial flow gas turbine stage (100) of said at least one axial flow gas turbine stage (26, 28, 100), wherein said first axial flow gas turbine stage (100) has a stator (84) having a blade row (50) according to any of claims 1 to 1 1 and/or a rotor having a blade row according to any of claims 1 to 12.

17. Gas turbine engine according to claim 16, wherein said at least one pressure-rise combustor (40) is a pulsejet combustor, a valve-less pulsejet combustor or a pulse detonation combustor.

18. Gas turbine engine according to any of claims 15 to 17, characterized in that said gas turbine engine (10) is a turbo propeller gas turbine engine, a turbofan gas turbine engine, a turbojet gas turbine engine or a turbo shaft gas turbine engine.