EP2447543A1 - Axial compressor and method for operating same - Google Patents

Abstract

The compressor (2) has a set of compressor guide blades (24) attached at a guide blade carrier and combined to form a set of guide blade rows. Another set of compressor guide blades (18) attached at a compressor disk (30) of a compressor shaft (20) and combined to form another set of guide blade rows. A coolant supply channel (64) is guided through another guide blade carrier (26) and the former set of guide blades that is arranged in a flow direction of a flow medium (S). The channel opens into a hollow space (54) at an end (36) of the set of blades over an outlet opening (74). An independent claim is also included for a method for operating an axial compressor.

Description

The invention relates to an axial compressor, in particular for a gas turbine, for compressing a flow medium with a plurality of combined to Leitschaufelreihen, each attached to a guide vane compressor guide vanes and combined with a plurality of rows of blades, each attached to a compressor disk of a compressor shaft compressor rotor blades, wherein in Flow direction of the flow medium seen last compressor disc has an end face which defines a cavity. The invention further relates to a method for operating such an axial compressor.

Gas turbines are used in many areas to drive generators or work machines. In this case, the energy content of a fuel is used to generate a rotational movement of a turbine shaft. For this purpose, the fuel is burned in a combustion chamber, compressed air being supplied by an air compressor. The air compressor is usually designed as axial compressor. The working medium produced in the combustion chamber by the combustion of the fuel, under high pressure and at high temperature, is guided via a turbine unit arranged downstream of the combustion chamber, where it relaxes to perform work. The air compressor, briefly compressor, and the turbine unit are usually arranged on a common shaft, so that in operation, the turbine unit drives the compressor.

The combustion chamber of the gas turbine may be embodied as a so-called annular combustion chamber, in which a plurality of circumferentially arranged around the turbine shaft burners in a common, surrounded by a high temperature resistant surrounding wall combustion chamber space. Is to the combustion chamber in its entirety designed as an annular structure. In addition to a single combustion chamber can also be provided a plurality of combustion chambers.

To generate the rotational movement of the turbine shaft, a number of turbine blades, which are usually combined into blade groups or rows of blades, are arranged thereon. In this case, a turbine disk is usually provided for each turbine stage, to which the turbine blades are fastened by means of their blade root. For guiding the flow of the working medium in the turbine unit, moreover, turbine guide vanes are usually connected between adjacent rotor blade rows and connected to the turbine housing and combined into rows of guide blades.

The air compressor of such a gas turbine is structurally usually constructed analogous to the turbine unit and comprises in one embodiment as an axial compressor a plurality of combined to form vane rows, each attached to a vane carrier compressor vanes and a plurality of combined into rows of blades, each attached to a compressor shaft compressor blades. A row of blades and a guide vane row immediately following in the flow direction of the flow medium (here: air) form a compressor stage. As a rule, several compressor stages are provided.

The entirety of all rotating parts of the gas turbine - in particular shaft and blades - is also referred to as a rotor, the fixed parts - in particular housings and vanes - are also referred to as a stator.

The compressor shaft is usually composed of a plurality of viewed in the axial direction one behind the other and together, for example, held together by a tie rods compressor discs together. Towards the turbine unit, the compressor shaft settles over a shaft intermediate piece continued as a turbine shaft. Each of the compressor disks usually carries at its periphery the compressor blades of a blade row, which are fastened with their blade roots in corresponding fastening grooves of the compressor disk.

The last compressor disk seen in the direction of flow of the flow medium has, in a conventional design, a head or end face facing the following turbine unit, which together with other components delimits or at least separates a cavity separated from the flow channel for the flow medium, hereinafter also referred to as a cavity partially encloses. Such a configuration is for example from the EP 1 640 587 B1 known (compare there FIG. 2 ).

In the design of such gas turbines in addition to the achievable power usually a particularly high efficiency is a design target. An increase in the efficiency can be achieved for thermodynamic reasons basically by increasing the outlet temperature, with the working medium from the combustion chamber off and flows into the turbine unit. Here are temperatures of the working medium of about 1200 ° C to 1500 ° C sought for such gas turbines and also achieved.

So that such high temperatures or correspondingly high efficiencies can be achieved, the air in the compressor should be compressed as much as possible. As a result of the increasing compressor pressure in the compressor along the flow direction of the gas, the temperature at the compressor end also rises along with it. Under certain circumstances, the maximum permissible operating temperature of the material of the rear compressor disks is achieved.

Currently, the maximum allowable operating temperature for available materials is a limiting constraint for the design of gas turbines with respect to compressor end temperature If there is a risk of exceeding this limit, for example at high ambient temperatures, the operating mode of the machine must be throttled. Thus, the potential of the gas turbine can not be fully utilized.

Technical solutions have been developed in the applicant's home by which the last compressor disk seen in the flow direction of the flow medium, in particular in its head or end region, is cooled by application of cooling air. For this purpose, for example, a cooling air cooler is used, which essentially serves to supply the front turbine blading with cooled down cooling air. This cooling air is fed into the rotor by the so-called shaft cover, namely a shaft cover or sheathing arranged behind the air compressor, viewed in the flow direction of the flow medium. From there, the cooling air then enters the turbine blading. Part of this cooling air is passed to cool the head region of the last compressor disk from the Shaftcover in the cavity behind the last compressor disk.

A disadvantage of this form of cooling air flow is that the cooling air for the cavity a relatively long way through various components of the gas turbine, namely first through the supply to the Shaftcover and then through the Shaftcover itself travels, which are lapped by warm compressor air. As a result, the temperature of the cooling air rises significantly before it reaches the cavity, whereby the cooling potential for the last compressor disk is greatly reduced.

The invention is therefore based on the object to further develop an axial compressor of the type mentioned above, that with simple means held effective cooling of the last compressor disk is achieved. It should continue to be specified a corresponding operating method.

With respect to the axial compressor this object is achieved by at least one mounted through the guide vane carrier, as seen in the flow direction of the flow medium behind the last blade row compressor vane leading coolant supply channel, which at the compressor shaft facing the head of the compressor vane on a arranged there outlet opening opens into the cavity.

The invention is based on the consideration that a particularly effective cooling of the last seen in the flow direction of the flow medium compressor disk can be achieved by being acted upon at its front side with coolant of relatively low temperature. It should be avoided that the introduced into the adjacent cavity coolant heats up too much on its way into the cavity. For this purpose, the coolant is supplied to the cavity adjacent to the compressor disk to be cooled according to the invention via at least one of the compressor guide vanes of the guide vane row directly opposite the cavity in the radial direction. This is usually seen in the flow direction of the flow medium last row of vanes, or in the case of a so-called double-row design, the penultimate vane row. Since this guide blade row and the hollow space to be cooled are located substantially in the same position in the axial direction, the supply of the coolant substantially in the radial direction from outside to inside, so that comparatively long transport or supply paths in the axial direction by various components of the gas turbine, in which an undesirable heating of the coolant could take place, be avoided.

Conveniently, cooling air is used as the coolant, for example, the compressor air flow further upstream in a colder region of the compressor as a partial flow is removed. Alternatively or additionally, a cooling of the cooling air by external cooling air cooler or the like may be provided.

In order to realize the principle of the invention, it is sufficient if only a single one of the compressor guide vanes of the corresponding guide blade row is provided with a coolant supply channel of the type mentioned. For a particularly effective and uniform cooling of the last compressor disk, however, several, preferably all, of the compressor guide vanes of a row of vanes arranged behind the last row of rotor blades are preferably provided with corresponding coolant supply ducts in the circumferential direction.

In such an embodiment, it is expedient if an annular coolant distributor chamber is provided in the guide blade carrier or in a surrounding housing component, to which the sections of the coolant supply channels arranged in the compressor guide vanes are connected.

In a further advantageous embodiment, the Verdichterleitschaufeln the arranged behind the last blade row Leitschaufelreihe (s) at its head end with a common annular body which limits on the one hand with its outer surface a flow channel for the comparatively hot at this point flow medium, and the other with its inner surface the limited to cooling cavity. The executed in the manner of a shroud ring body thus seals the cavity from the flow channel and insulated both areas of space thermally from each other.

In this case, the respective coolant supply channel is expediently guided through the annular body, so that the coolant outlet opening is located on its inner surface facing the cavity.

In a further expedient embodiment, the cavity on the end face of the compressor disk opposite side by an end face of an example ring or hollow cylindrical rotor cover (Shaftcover) be limited. The end face of the rotor cover is also cooled during operation.

Particularly preferred is the use of the axial compressor described as an air compressor in a gas turbine, wherein advantageously the air compressor and the turbine unit of the gas turbine are arranged along a common shaft. Of course, it is also conceivable to operate the axial compressor as an independent device for other purposes in which a flow medium is to be compressed. The aforementioned advantages of the improved cooling of the compressor end region also apply in this case.

With regard to the method, the object mentioned at the outset is achieved by introducing a coolant into the cavity via at least one coolant supply channel, which is guided behind the last blade row by a guide vane carrier and by a guide vane carrier, viewed in the flow direction of the flow medium is initiated. The embodiments and advantages mentioned in connection with the device are analogously transferred to the method.

The advantages associated with the invention are, in particular, that in the case of an axial compressor an effective cooling of the thermally highly stressed last compressor disk and of the adjacent components is made possible by supply of coolant substantially in the radial direction via a spatially adjacent row of guide blades. This results in a lower thermal load on the compressor disk or a possible increase in the compressor discharge pressure with constant load on the compressor disk.

FIG 1

einen Halbschnitt durch eine Gasturbine,

FIG 2

einen vergrößerten Ausschnitt des Verdichters der Gasturbine gemäß FIG 1 mit einer bislang vorgesehenen Kühlvorrichtung zur Kühlung der letzten Verdichterscheibe, und

FIG 3

denselben Ausschnitt mit einer gegenüber FIG 2 verbesserten Kühlvorrichtung zur Kühlung der letzten Verdichterscheibe.

An embodiment of the invention will be explained in more detail with reference to a drawing. Show:

FIG. 1

a half-section through a gas turbine,

FIG. 2

an enlarged section of the compressor of the gas turbine according to FIG. 1 with a previously provided cooling device for cooling the last compressor disk, and

FIG. 3

same section with one opposite FIG. 2 improved cooling device for cooling the last compressor disk.

Identical parts are provided with the same reference numerals in all figures.

The gas turbine 1 according to FIG. 1 has a compressor 2 for combustion air, a combustion chamber 4 and a turbine unit 6 for driving the compressor 2 and a generator, not shown, or a working machine. For this purpose, the turbine unit 6 and the compressor 2 are arranged on a common, also called turbine rotor turbine shaft 8, with which the generator or the working machine is connected, and which is rotatably mounted about its central axis 9. The running in the manner of an annular combustion chamber 4 is equipped with a number of burners 10 for the combustion of a liquid or gaseous fuel.

The turbine unit 6 has a number of rotatable turbine blades 12 connected to the turbine shaft 8. The turbine blades 12 are annularly arranged on the turbine shaft 8 and thus form a number of blade rows. Furthermore, the turbine unit 6 comprises a number of stationary turbine vanes 14, which also annularly attach to a vane support 16 of the turbine unit 6 to form rows of vanes are. The turbine blades 12 serve to drive the turbine shaft 8 by momentum transfer from the turbine unit 6 flowing through the working medium M. The turbine vanes 14, however, serve to guide the flow of the working medium M between two seen in the flow direction of the working medium M consecutive blade rows or blade rings. A successive pair of a turbine nozzle vane 14 or vane row and a ring of turbine blades 12 or a blade row is also referred to herein as a turbine stage.

The compressor 2 of the gas turbine 1 is similar in construction to the turbine unit 6. It comprises a plurality of rotor blades combined compressor blades 18, which are fixed with their blade roots on the turbine shaft 8, referred to in this section of the gas turbine 1 as the compressor shaft 20, and in a flow channel 22 for the sucked flow medium S, here air, protrude. The compressor blades 20, which are set in rotation via the compressor shaft 20, perform compression work on the flow medium S and convey it in the direction of the turbine unit 8. The fixed compressor guide vanes 24, however, serve to guide the flow medium S between two flow directional mediums in succession blade rows. The compressor vanes 24 are attached to associated vanes 26, which in turn are connected in a manner not shown to the outer compressor housing, and - optionally together with other ring segments - form the outer boundary of the flow channel 22. The vane carriers 26 may be composed of several segments. A successive pair of a pair of compressor blades 18 or a blade row and a ring of compressor vanes 24 or a row of guide vanes is also referred to as a compressor stage.

FIG. 2 shows an enlarged view of the end or exit region of the compressor 2 and the adjoining in the flow direction 28 of the flow medium S transition region to the combustion chamber 4 and the turbine unit. 6

The compressor shaft 20 is composed of a plurality of stacked successively arranged compressor discs 30, of which in FIG. 2 only one, namely seen in the flow direction 28 of the flow medium S rear or last compressor disk 30 is visible. The respective compressor disk 30 carries at its periphery the compressor blades 18 of the associated blade row. In the intermediate areas in which the Verdichterleitschaufeln 24 are arranged, the inner boundary of the flow channel 22, however, each formed by the outer side 32 of an annular body 34, which with the head ends 36 of the compressor guide vanes 24 of the associated guide vane row is connected. The annular gap 38 located between the respective spatially fixed ring body 34 and the axially adjacent rotating disc 30 can be sealed in a conventional manner, for example by a labyrinth seal 40.

In the embodiment according to FIG. 2 The last compressor stage behind the last row of blades comprises two rows of guide vanes (so-called double row arrangement) which follow one another directly, to which a common annular body 34 is assigned. However, this detail is not relevant in the present case.

Towards the combustion chamber 4, the flow channel 22 widens for the compressed in the compressor 2 flow medium S in the manner of a diffuser. The inner boundary of the flow channel 22 is formed in this area by the peripheral surface of an annular shaft cover 42, the so-called Shaftcover. The fixed shaft cover 48 surrounds the rotating turbine shaft 8, which extends as an extension of the compressor shaft 20 to the turbine unit 6, and may be composed of individual shaft segments 44 or discs. In the axial direction towards the compressor 2, the shaft cover 42 extends almost to the ring body 34 of the last (double) vane row. At the end face directed towards the compressor 2, the shaft cover 42 has an annular flange 46 with an end face 48, which is spaced apart from the annular body 34 by an axial annular gap 50. Toward the turbine shaft 8, the annular flange 46 is spaced apart by a further, here radial annular gap 52. The annular gap 50, like the annular gap 38, can be provided with suitable sealing means to prevent overflow of the comparatively hot flow medium S from the flow channel 22 into that of the last compressor disk 30, the annular body 34, the annular flange 46 and the corresponding section of the turbine shaft 8 or enclosed cavity 54, also referred to as cavity, to prevent.

Despite these sealing measures, the components adjacent to the cavity 54, in particular the last compressor disk 30, may be exposed to considerable thermal stress during operation of the gas turbine 1 or the compressor 2. To reduce this burden is in the gas turbine 1 according FIG. 2 an introduction of a coolant K, here cooling air, provided in the cavity 54. The coolant K is guided via a coolant supply line 56 into an integrated into the shaft cover 42 coolant channel 58, for example, cylindrical contour. From there it flows through one or more introduced into the annular flange 46 overflow channels 60 in the cavity 54, so that the desired cooling of the last compressor disk 30 is realized in the manner of an impingement cooling on the end face 62. The removal of the "spent" coolant K, for example, by gap leakage at the annular gaps 38, 50 and 52nd

In this cooling concept, however, it is disadvantageous that the coolant on its way through the coolant supply line 56, which is passed for reasons of space at the combustion chamber 4, and through the coolant channel 58 within the housing of the shaft cover 42 - in countercurrent to the outside of the shaft cover 42nd flowing past, by the compressor 2 comparatively strongly heated flow medium S - is strongly warmed and thereby loses its cooling potential in part.

To avoid this problem is in the gas turbine 1 according FIG. 3 provided an alternative cooling concept. The structure of the compressor 2 and the turbine unit 6 basically corresponds to that FIG. 2 , so that at this point only the differences will be discussed.

In the variant according to FIG. 3 the coolant supply through the shaft cover 42 is dispensed with. The annular flange 46 therefore has no overflow. Instead, the coolant is supplied into the cavity 54 via a number of compressor guide vanes 24 of the last viewed in the flow direction 28 of the flow medium S (double) Leitschaufelreihe. For this purpose, at least one coolant supply channel 64 is provided, which is guided in a first section 66 through the corresponding guide blade carrier 26 to the relevant compressor guide vane 24 and continues in a second section 68 in the interior of its airfoil 70 substantially in the radial direction. At the top end 36 of the compressor vane 24, the coolant supply channel 64 is in a in FIG. 3 For example, obliquely extending end portion 72 within the connected to the compressor guide vane 24 ring body 34 and ends at the cavity 54 facing inside 73 of the ring body 34 in an outlet opening 74. Due to the comparatively short and direct Zuströmweg the cavity 54 supplied coolant K only weak heated and thus can a comparatively large cooling potential, especially in the cooling of the end face 62 of the last compressor disk 30 unfold.

In a particularly advantageous variant, an approximately uniform in the circumferential direction coolant supply takes place in the annular space designed as an annular space 54 over several of the compressor guide vanes 24 of the corresponding row of vanes, for example, in the circumferential direction every, every second or every third, etc. of the compressor guide vanes 24 of this row Leitschaufelreihe with a corresponding portion of a coolant supply channel 64 and with a corresponding outlet opening 74 may be provided for coolant K. These channel sections are thus connected in parallel on the coolant side and, for example, via an in FIG. 3 only schematically indicated, arranged in the guide blade carrier 26 or in an adjacent housing component, circumferentially circulating coolant distribution chamber 76 fed simultaneously with fresh coolant K.

It is understood that some of the in FIG. 3 Details shown have only exemplary character. Notwithstanding the representation chosen here, for example, the orientation of the end portion 72 of the coolant supply channel 64 and the position of the outlet opening 74 vary, so as to realize a baffle cooling directly on the end face 62 of the compressor disk 30. Also, the coolant supply could alternatively or additionally take place via the rear row of guide vanes of the double row, or it could be provided only a simple row of guide vanes.

Furthermore, it is conceivable that the two cooling concepts according to FIG. 2 and FIG. 3 to combine with each other.

Claims (8)

Axialcompressor (2), in particular for a gas turbine (1), for compressing a flow medium (S) having a plurality of compressor guide vanes (24) combined into guide vane carriers (16) and combined with a plurality of rotor blade rows compressor rotor blades (18) attached to a compressor disk (30) of a compressor shaft (20),
wherein in the flow direction of the flow medium (S) seen last compressor disk (30) has an end face (62) which defines a cavity (54), characterized by
at least one coolant supply channel (64) which is guided by a guide blade carrier (26) and by a compressor guide vane (24) arranged behind the last row of blades in the flow direction of the flow medium (S) and which is guided on the compressor shaft (20) ) facing head end (36) of the compressor guide vane (24) via a discharge opening (74) arranged there opens into the cavity (54). Axial compressor (2) according to claim 1,
wherein, seen in the circumferential direction, a plurality of the compressor guide vanes (24) of a row of vanes arranged behind the last row of blades are equipped with corresponding coolant supply channels (64). Axial compressor (2) according to claim 2,
wherein an annular coolant distribution chamber (76) is provided in the vane support (16) or in a surrounding housing component to which the portions of the coolant supply passages (64) disposed in the compressor vanes (24) are connected. Axialcompressor (2) according to one of claims 1 to 3, wherein the compressor guide vanes (24) of the row of vanes (n) arranged behind the last row of blades are connected at their head end (36) to an annular body (34) which on the one hand comprises a flow channel (22). for the flow medium (S) and on the other hand, the cavity (54) limited. Axial compressor (2) according to claim 4,
wherein the respective coolant supply channel (64) is guided through the annular body (34). Axial compressor (2) according to one of claims 1 to 5, wherein the cavity (54) on the end face (62) of the compressor disk (30) opposite side by an end face (48) of a rotor cover (62) is limited. Gas turbine (1) with an axial compressor (2) according to one of claims 1 to 6. Method for operating an axial compressor (2) having a plurality of compressor guide vanes (24) combined to form guide vane carriers (16) and having a plurality of rotor vanes combined, each attached to a compressor disk (30) of a compressor shaft (20) Compressor blades (24),
wherein the last compressor disk (30) seen in the flow direction of the flow medium (S) has an end face (62) which at least partially surrounds a cavity (54),
characterized in that
via at least one coolant supply channel (64), which is guided through a guide vane carrier (26) and through a compressor guide vane (24) attached to the vane carrier (26), as seen behind the last row of rotor blades in the flow direction of the flow medium (S); K) is introduced into the cavity (54).

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