US20100098530A1 - Compressor for a gas turbine - Google Patents
Compressor for a gas turbine Download PDFInfo
- Publication number
- US20100098530A1 US20100098530A1 US12/576,762 US57676209A US2010098530A1 US 20100098530 A1 US20100098530 A1 US 20100098530A1 US 57676209 A US57676209 A US 57676209A US 2010098530 A1 US2010098530 A1 US 2010098530A1
- Authority
- US
- United States
- Prior art keywords
- compressor
- shroud
- exhaust cavity
- cavity
- arrangement
- Prior art date
- Legal status (The legal status is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the status listed.)
- Abandoned
Links
Images
Classifications
-
- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F02—COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
- F02C—GAS-TURBINE PLANTS; AIR INTAKES FOR JET-PROPULSION PLANTS; CONTROLLING FUEL SUPPLY IN AIR-BREATHING JET-PROPULSION PLANTS
- F02C9/00—Controlling gas-turbine plants; Controlling fuel supply in air- breathing jet-propulsion plants
- F02C9/16—Control of working fluid flow
- F02C9/18—Control of working fluid flow by bleeding, bypassing or acting on variable working fluid interconnections between turbines or compressors or their stages
-
- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
- F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
- F01D11/00—Preventing or minimising internal leakage of working-fluid, e.g. between stages
- F01D11/001—Preventing or minimising internal leakage of working-fluid, e.g. between stages for sealing space between stator blade and rotor
-
- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F04—POSITIVE - DISPLACEMENT MACHINES FOR LIQUIDS; PUMPS FOR LIQUIDS OR ELASTIC FLUIDS
- F04D—NON-POSITIVE-DISPLACEMENT PUMPS
- F04D29/00—Details, component parts, or accessories
- F04D29/40—Casings; Connections of working fluid
- F04D29/52—Casings; Connections of working fluid for axial pumps
- F04D29/54—Fluid-guiding means, e.g. diffusers
- F04D29/541—Specially adapted for elastic fluid pumps
- F04D29/542—Bladed diffusers
-
- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F04—POSITIVE - DISPLACEMENT MACHINES FOR LIQUIDS; PUMPS FOR LIQUIDS OR ELASTIC FLUIDS
- F04D—NON-POSITIVE-DISPLACEMENT PUMPS
- F04D29/00—Details, component parts, or accessories
- F04D29/66—Combating cavitation, whirls, noise, vibration or the like; Balancing
- F04D29/68—Combating cavitation, whirls, noise, vibration or the like; Balancing by influencing boundary layers
- F04D29/681—Combating cavitation, whirls, noise, vibration or the like; Balancing by influencing boundary layers especially adapted for elastic fluid pumps
- F04D29/682—Combating cavitation, whirls, noise, vibration or the like; Balancing by influencing boundary layers especially adapted for elastic fluid pumps by fluid extraction
-
- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
- F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
- F05D2260/00—Function
- F05D2260/60—Fluid transfer
- F05D2260/602—Drainage
- F05D2260/6022—Drainage of leakage having past a seal
Definitions
- This invention relates to a compressor for a gas turbine, in particular an aircraft gas turbine, with a rotor hub carrying rotor blades, with a stator equipped with stator vanes, with a shroud associated to the stator vanes, and with an arrangement providing sealing between the shroud and rotor hub to prevent leakage.
- the compressor in particular an axial-flow compressor, includes a rotor shaft with one or several compressor stages having a rotor and a stator.
- the compressor may be provided with a so-called inlet guide vane assembly upstream of the first stage.
- the rotor equipped with rotor blades is connected to the compressor or the rotor shaft, respectively, and rotates in the casing of the compressor.
- the stator of each stage is equipped with stator vanes and is not connected to the compressor shaft. It is therefore stationary. Rotating and non-rotating blade rows alternate in the axial direction of the rotor shaft in a compressor.
- Suitable attachment must be provided for the stators.
- the stator has a hub gap, with the rotor shaft extending underneath the stator. In this arrangement, which is found, for example, on high-pressure compressors, the stator is connected to the compressor casing only. 2) On the radially inward side of the stator vanes the stator is provided with a so-called shroud, with the shroud being a ring connecting to the stator at the hub.
- This arrangement involves that, between rotor and stator hub upstream and downstream of the stator, an axial gap is formed which leads to flow leakage. This leakage must be minimized by a sealing arrangement in the form of seals relative to the rotor shaft. This arrangement is found on the sealing system for a gas turbine according to U.S. Pat. No. 6,932,349 B2. The smaller the gap between seal and shroud, the smaller the leakage flow between the forward and the rearward axial gap.
- a strong transverse duct flow which interacts with the leakage flow through the shroud and is disturbed, occurs on the stationary hub shroud from the pressure side to the suction side of the stator.
- the transverse duct flow tends to flow to the vane suction side.
- flow separation on the stator vane and, thus, severe losses occur.
- the leakage flow through the axial gap which is situated before and behind the shroud, is driven by a strong pressure gradient between the leading edge and the trailing edge of the stator and, as it interacts with the main flow, can consequently get large and produce severe losses.
- heating of the flow is encountered within the shroud cavity. In order to reduce these adverse aerodynamic effects, leakage must be minimized.
- the present invention in a broad aspect, provides that the sealing arrangement between shroud and rotor hub is formed by a discharge arrangement for leakage air. Provision is thereby made for an almost complete suppression or elimination of leakage air and, concurrently, a simplification of design and manufacture.
- the present invention is implemented in that the state-of-the-art seals are dispensed with and the leakage flow within the shroud is discharged to the outside through an exhaust cavity within the shroud and the stator vane.
- Discharge is accomplished in that the static pressure applied to the exhaust cavity is lower than the pressure prevailing in the shroud cavity.
- the static pressure must here be that much lower than the pressure prevailing in the shroud cavity, that the desired amount of reduced mass flow is removed.
- exhaust cavity is connected via suitable lines to either an upstream compressor duct or to a secondary air system of the engine, with the removed air being used as turbine cooling air or for the pressurization of the aircraft cabin or other engine or aircraft systems, for example.
- the amount of the removed, absolute leakage mass flow is controllable in that a defined throat is provided within the exhaust cavity or within the discharge arrangement.
- the leakage mass flow through a shroud cavity with seals amounts to approx. 0.5-1 percent of the total flow. Consequently, the exhaust cavity can be designed such that an amount of this size is exhausted, or, if used for example for pressurizing the aircraft cabin or for turbine cooling, it should be dimensioned such that the required mass flow, which normally amounts to approx. 2-4 percent, is discharged.
- both the boundary layer before the stator and the boundary layer behind the stator are ingested into the shroud cavity and finally discharged to the outside. This leads to a significant reduction of the stator losses and, thus, to an increase in compressor efficiency. With use being also made of the discharged air, further potential for lowering fuel consumption is provided.
- the exhaust cavity itself can have any cross-sectional shape and be provided, for example, as discrete round, elliptical or polygonal openings in any number and position on the shroud. Within the shroud, the cavity can either remain discrete or expand into a circumferential cavity. Vane passage can be provided by one or several tubes or by hollowing the vane, for example.
- the shroud without seals according to the present invention is significantly easier to design and manufacture than an arrangement with seals. This provides for weight and cost savings. Furthermore, the leakage flow, which entails both losses and heating of the fluid and the components, is nonexistent. Exhaustion of the air leads to ingestion of the boundary layer of the compressor side wall, which further contributes to a reduction of losses. Utilization of the exhausted air for secondary systems provides for a further increase in efficiency. This means increased compressor efficiency and reduced fuel consumption. In summary, an increase in stage efficiency of approx. 0.2 percent is obtainable.
- FIG. 1 shows a schematic longitudinal section through a compressor in accordance with the state of the art
- FIG. 2 is a schematic representation of the flow phenomena on a compressor in accordance with the state of the art as per FIG. 1 ,
- FIG. 3 shows a schematic longitudinal section through a compressor in accordance with the present invention
- FIG. 4 shows the air discharge on a compressor in accordance with the present invention as per FIG. 3 ,
- FIG. 5 is a representation of the air discharge on a compressor in accordance with the present invention.
- FIG. 6 shows a variant of the air discharge on a compressor in accordance with the present invention
- FIG. 7 is a schematic representation of the discharge sections of the leakage air on the shroud
- FIG. 8 is a representation, analogically to FIG. 7 , with other discharge sections, and
- FIGS. 9 a to c show different air passage sections for the leakage air in a stator vane.
- FIG. 1 shows, in a schematic longitudinal section through a stage of a compressor according to the state of the art, a rotor 1 with rotor shaft 2 , rotor hub 3 and a rotor blade 4 as well as a stator 5 with casing 6 and a stator vane 7 with a shroud 8 encompassing all stator vanes 7 of the compressor stage on the radially inward side. Disposed between the rotor blade 4 and the stator 5 is a sealing gap 9 .
- a radially extending shroud cavity 11 is provided in the rotor hub 3 which clears a forward axial gap 12 and a rearward axial gap 13 in the axial direction of the rotor shaft 2 and towards the rotor hub 3 thereof.
- two cross-sectionally frustum-shaped seals 16 and 17 are provided as sealing arrangement 10 , each forming a radial gap 18 to the sealing surface 15 a of the shroud 8 .
- leakage flow 22 through the axial gap 12 , 13 which is situated before and behind the shroud 8 , is driven by a strong pressure gradient between the leading and the trailing edge of the stator 5 and, as it interacts with the main flow, can consequently get large and produce severe losses. Furthermore, heating of the flow is encountered within the shroud cavity 11 .
- leakage must be minimized. This can be accomplished by either increasing the number of seals 16 , 17 or reducing the radial gap 18 between seal 16 , 17 and shroud 8 . In either case, the geometry is complicated in terms of design and manufacture. As a consequence, increased costs are incurred. Since control of the radial gap 18 is difficult to achieve, the operational risk for the compressor is increased as well. Summarizing, then, leakage should be completely prevented as it adversely affects the efficiency of the compressor and, consequently, the fuel consumption of the engine.
- FIGS. 3 to 9 show the aircraft gas-turbine compressor according to the present invention with rotor 1 , rotor shaft 2 , rotor hub 3 , rotor blade 4 , stator 5 , casing 6 , stator vane 7 , shroud 8 , sealing gap 9 , shroud cavity 11 , forward and rearward axial gap 12 , 13 and clearance 14 with hub-side bottom 15 b.
- a discharge arrangement 23 for the leakage air which is removed from the shroud cavity 11 via an air discharge opening 24 and an air discharge duct 25 through the shroud 8 and the stator vane 7 .
- the air discharge duct 25 here forms an exhaust cavity 30 for leakage air.
- the embodiment according to the present invention is implemented in that the state-of-the-art seals 16 , 17 in accordance with FIG. 1 are dispensed with and the leakage flow 22 from the shroud cavity 11 is discharged to the outside through the exhaust cavity 30 of the discharge arrangement 23 within the shroud 8 and the stator vane 7 .
- Discharge is accomplished as per FIG. 5 in that the static pressure p a applied to the exhaust cavity 30 in the form of the air discharge duct 25 is lower than the pressure p d prevailing in the shroud cavity 11 .
- the static pressure p a must here be sufficiently lower than the pressure pa prevailing in the shroud cavity 11 , that the desired amount of reduced mass flow is removed.
- the amount of the removed, absolute leakage mass flow is controllable in that a defined throat 29 ( FIG. 4 ) is provided within the exhaust cavity 30 or within the discharge arrangement 23 .
- the required area 40 ( FIG. 3 ) can be determined by way of the pressures p 1 and p 2 of the ingested boundary layers 31 or 32 , respectively, before and behind the shroud 8 and the exhaust mass flow ( FIG. 5 ).
- the leakage mass flow through the shroud cavity 11 with seals 16 , 17 generally amounts to approx. 0.5-1 percent of the total flow. Consequently, the exhaust cavity 30 can be designed such that an amount of this size is exhausted, or, if used for example for pressurizing the aircraft cabin or for turbine cooling, it should be dimensioned such that the required mass flow, which normally amounts to approx. 2-4 percent, is discharged.
- both the boundary layer 31 before the stator 7 and the boundary layer 32 behind the stator 7 are ingested into the shroud cavity 30 and finally discharged to the outside ( FIG. 5 ).
- With use being also made of the discharged air further potential for lowering fuel consumption of the engine, is provided.
- FIG. 6 A variant of the air discharge is shown in FIG. 6 . While in the arrangement according to FIGS. 3 to 5 the air discharge opening 24 is disposed in the radial end face of the shroud 8 , FIG. 6 shows two air discharge openings 41 , 42 which, arranged in the axial end faces of the shroud 8 , form the exhaust cavity 30 connected to the air discharge ducts 25 in the shroud 8 .
- the exhaust cavity 30 itself can have any cross-sectional shape. See FIGS. 7 and 8 , for example, showing discrete polygonal, round or elliptical openings 33 , 34 or 35 , respectively, in any number and position on the shroud 8 . Within the shroud 8 , the exhaust cavity 30 can either remain discrete or expand into a circumferential cavity. Passage through the stator vane 7 can be provided by one or several tubes as round hole 36 , oblong hole 37 or slot 38 , or by hollowing the stator vane 7 in the form of a hollow chamber 39 ( FIGS. 9 a - c ), for example.
- the shroud 8 without seals according to the present invention is significantly easier to design and manufacture than an arrangement with seals 16 , 17 in accordance with the state of the art as per FIGS. 1 and 2 .
- This provides for weight and cost savings.
- the leakage flow which entails both losses and heating of the fluid and the components, is nonexistent.
- Exhaustion of the air leads to ingestion of the boundary layer of the compressor side wall, which further contributes to a reduction of losses.
- Utilization of the exhausted air for secondary systems provides for a further increase in efficiency. This means increased compressor efficiency and reduced fuel consumption of the engine. In summary, an increase in stage efficiency of approx. 0.2 percent is obtainable.
Landscapes
- Engineering & Computer Science (AREA)
- Mechanical Engineering (AREA)
- General Engineering & Computer Science (AREA)
- Chemical & Material Sciences (AREA)
- Combustion & Propulsion (AREA)
- Physics & Mathematics (AREA)
- Fluid Mechanics (AREA)
- Structures Of Non-Positive Displacement Pumps (AREA)
Abstract
A compressor for a gas turbine, in particular an aircraft gas turbine, has a rotor hub carrying rotor blades, a stator equipped with stator vanes, a shroud associated to the stator vanes, and an arrangement providing sealing between the shroud and rotor hub to prevent leakage. To achieve almost complete suppression of leakage air and, concurrently, simplification of design and manufacture, the sealing arrangement (20) between the shroud (8) and rotor hub (3) is formed by a discharge arrangement (23) for leakage air.
Description
- This application claims priority to German
Patent Application DE 10 2008 052 101.9 filed Oct. 20, 2008, the entirety of which is incorporated by reference herein. - This invention relates to a compressor for a gas turbine, in particular an aircraft gas turbine, with a rotor hub carrying rotor blades, with a stator equipped with stator vanes, with a shroud associated to the stator vanes, and with an arrangement providing sealing between the shroud and rotor hub to prevent leakage.
- The compressor, in particular an axial-flow compressor, includes a rotor shaft with one or several compressor stages having a rotor and a stator. The compressor may be provided with a so-called inlet guide vane assembly upstream of the first stage. The rotor equipped with rotor blades is connected to the compressor or the rotor shaft, respectively, and rotates in the casing of the compressor. The stator of each stage is equipped with stator vanes and is not connected to the compressor shaft. It is therefore stationary. Rotating and non-rotating blade rows alternate in the axial direction of the rotor shaft in a compressor.
- Suitable attachment must be provided for the stators. For this, two options exist: 1) The stator has a hub gap, with the rotor shaft extending underneath the stator. In this arrangement, which is found, for example, on high-pressure compressors, the stator is connected to the compressor casing only. 2) On the radially inward side of the stator vanes the stator is provided with a so-called shroud, with the shroud being a ring connecting to the stator at the hub. This arrangement involves that, between rotor and stator hub upstream and downstream of the stator, an axial gap is formed which leads to flow leakage. This leakage must be minimized by a sealing arrangement in the form of seals relative to the rotor shaft. This arrangement is found on the sealing system for a gas turbine according to U.S. Pat. No. 6,932,349 B2. The smaller the gap between seal and shroud, the smaller the leakage flow between the forward and the rearward axial gap.
- The arrangement with a hub shroud is mechanically complex, heavy and expensive. Also, the replacement of worn seals incurs considerable assembly effort and cost investment.
- A strong transverse duct flow, which interacts with the leakage flow through the shroud and is disturbed, occurs on the stationary hub shroud from the pressure side to the suction side of the stator. The transverse duct flow tends to flow to the vane suction side. As a consequence, flow separation on the stator vane and, thus, severe losses occur. The leakage flow through the axial gap, which is situated before and behind the shroud, is driven by a strong pressure gradient between the leading edge and the trailing edge of the stator and, as it interacts with the main flow, can consequently get large and produce severe losses. Furthermore, heating of the flow is encountered within the shroud cavity. In order to reduce these adverse aerodynamic effects, leakage must be minimized. This can be accomplished by either increasing the number of seals or reducing the gap between seal and shroud. In either case, the geometry is complicated in terms of design and manufacture. As a consequence, increased costs are incurred. Since control of the sealing gap is only difficult to achieve, the operational risk for the compressor is increased as well. Summarizing, then, leakage should be completely prevented as it adversely affects the efficiency of the compressor and, consequently, the fuel consumption of the engine.
- It is a particular object of the present invention to avoid these disadvantages of the state of the art.
- The present invention, in a broad aspect, provides that the sealing arrangement between shroud and rotor hub is formed by a discharge arrangement for leakage air. Provision is thereby made for an almost complete suppression or elimination of leakage air and, concurrently, a simplification of design and manufacture.
- Since special seals are dispensed with in this arrangement, no assembly effort is required for the replacement thereof in the case of wear. This provides for reduced costs and risks as well as improved efficiency and lower fuel consumption.
- The present invention is implemented in that the state-of-the-art seals are dispensed with and the leakage flow within the shroud is discharged to the outside through an exhaust cavity within the shroud and the stator vane.
- Discharge is accomplished in that the static pressure applied to the exhaust cavity is lower than the pressure prevailing in the shroud cavity. The static pressure must here be that much lower than the pressure prevailing in the shroud cavity, that the desired amount of reduced mass flow is removed.
- This can be accomplished, for example, in that the exhaust cavity is connected via suitable lines to either an upstream compressor duct or to a secondary air system of the engine, with the removed air being used as turbine cooling air or for the pressurization of the aircraft cabin or other engine or aircraft systems, for example.
- The amount of the removed, absolute leakage mass flow is controllable in that a defined throat is provided within the exhaust cavity or within the discharge arrangement.
- Complete suppression of the leakage flow between the rearward and the forward axial gap requires that the shroud geometry, more precisely the flow-wetted area of the shroud, be appropriately attuned. The required area can be determined by way of the pressures before and behind the shroud and the exhaust mass flow.
- Generally, the leakage mass flow through a shroud cavity with seals amounts to approx. 0.5-1 percent of the total flow. Consequently, the exhaust cavity can be designed such that an amount of this size is exhausted, or, if used for example for pressurizing the aircraft cabin or for turbine cooling, it should be dimensioned such that the required mass flow, which normally amounts to approx. 2-4 percent, is discharged.
- By way of the exhaust cavity, both the boundary layer before the stator and the boundary layer behind the stator are ingested into the shroud cavity and finally discharged to the outside. This leads to a significant reduction of the stator losses and, thus, to an increase in compressor efficiency. With use being also made of the discharged air, further potential for lowering fuel consumption is provided.
- The exhaust cavity itself can have any cross-sectional shape and be provided, for example, as discrete round, elliptical or polygonal openings in any number and position on the shroud. Within the shroud, the cavity can either remain discrete or expand into a circumferential cavity. Vane passage can be provided by one or several tubes or by hollowing the vane, for example.
- The shroud without seals according to the present invention is significantly easier to design and manufacture than an arrangement with seals. This provides for weight and cost savings. Furthermore, the leakage flow, which entails both losses and heating of the fluid and the components, is nonexistent. Exhaustion of the air leads to ingestion of the boundary layer of the compressor side wall, which further contributes to a reduction of losses. Utilization of the exhausted air for secondary systems provides for a further increase in efficiency. This means increased compressor efficiency and reduced fuel consumption. In summary, an increase in stage efficiency of approx. 0.2 percent is obtainable.
- The present invention is more fully described below, in light of the accompanying figures showing an embodiment of a compressor for a gas turbine, here an aircraft gas turbine:
-
FIG. 1 shows a schematic longitudinal section through a compressor in accordance with the state of the art, -
FIG. 2 is a schematic representation of the flow phenomena on a compressor in accordance with the state of the art as perFIG. 1 , -
FIG. 3 shows a schematic longitudinal section through a compressor in accordance with the present invention, -
FIG. 4 shows the air discharge on a compressor in accordance with the present invention as perFIG. 3 , -
FIG. 5 is a representation of the air discharge on a compressor in accordance with the present invention, -
FIG. 6 shows a variant of the air discharge on a compressor in accordance with the present invention, -
FIG. 7 is a schematic representation of the discharge sections of the leakage air on the shroud, -
FIG. 8 is a representation, analogically toFIG. 7 , with other discharge sections, and -
FIGS. 9 a to c show different air passage sections for the leakage air in a stator vane. -
FIG. 1 shows, in a schematic longitudinal section through a stage of a compressor according to the state of the art, arotor 1 withrotor shaft 2,rotor hub 3 and arotor blade 4 as well as astator 5 withcasing 6 and astator vane 7 with ashroud 8 encompassing allstator vanes 7 of the compressor stage on the radially inward side. Disposed between therotor blade 4 and thestator 5 is asealing gap 9. For accommodating theshroud 8, a radially extendingshroud cavity 11 is provided in therotor hub 3 which clears a forwardaxial gap 12 and a rearwardaxial gap 13 in the axial direction of therotor shaft 2 and towards therotor hub 3 thereof. In theradial clearance 14 between the radiallyinner sealing surface 15 a of theshroud 8 and the hub-side bottom 15 b of theclearance 14, two cross-sectionally frustum-shaped seals 16 and 17 are provided as sealingarrangement 10, each forming aradial gap 18 to the sealingsurface 15 a of theshroud 8. - In this arrangement, a leakage flow occurs between the
rotor hub 3 and theshroud 8 of thestator 5 in the forward and the rearwardaxial gap rotor shaft 2. The smaller theradial gap 18 between the seals 16, 17 and theshroud 8, the lower the flow leakage between the forward and the rearwardaxial gap - As illustrated in
FIG. 2 , with therotor 1 moving inrotary direction 19, a strongtransverse duct flow 21 occurs on thestationary shroud 8 of thestator vanes 7 from the pressure side D to the suction side S of thestator 5 which interacts with theleakage flow 22 through theshroud 8 and is disturbed. Thetransverse duct flow 21 tends to flow to the suction side S of thestator vane 7. This leads to flow separation on thestator vane 7 and consequentially to severe losses. Theleakage flow 22 through theaxial gap shroud 8, is driven by a strong pressure gradient between the leading and the trailing edge of thestator 5 and, as it interacts with the main flow, can consequently get large and produce severe losses. Furthermore, heating of the flow is encountered within theshroud cavity 11. In order to reduce these adverse aerodynamic effects, leakage must be minimized. This can be accomplished by either increasing the number of seals 16, 17 or reducing theradial gap 18 between seal 16, 17 andshroud 8. In either case, the geometry is complicated in terms of design and manufacture. As a consequence, increased costs are incurred. Since control of theradial gap 18 is difficult to achieve, the operational risk for the compressor is increased as well. Summarizing, then, leakage should be completely prevented as it adversely affects the efficiency of the compressor and, consequently, the fuel consumption of the engine. -
FIGS. 3 to 9 show the aircraft gas-turbine compressor according to the present invention withrotor 1,rotor shaft 2,rotor hub 3,rotor blade 4,stator 5,casing 6,stator vane 7,shroud 8, sealinggap 9,shroud cavity 11, forward and rearwardaxial gap clearance 14 with hub-side bottom 15 b. - Provided between the
shroud 8 and therotor hub 3 as sealingarrangement 20 is adischarge arrangement 23 for the leakage air which is removed from theshroud cavity 11 via anair discharge opening 24 and anair discharge duct 25 through theshroud 8 and thestator vane 7. Theair discharge duct 25 here forms anexhaust cavity 30 for leakage air. - This arrangement provides for almost complete suppression of leakage air and, concurrently, simplification of design and manufacture. Since special seals 16, 17—as in the state of the art according to FIG. 1—are here dispensed with, no assembly effort for a replacement thereof is required in the case of wear. This provides for a reduction of costs and risks as well as improved efficiency and, thus, reduced fuel consumption of the aircraft gas turbine using the compressor according to the present invention.
- The embodiment according to the present invention is implemented in that the state-of-the-art seals 16, 17 in accordance with
FIG. 1 are dispensed with and theleakage flow 22 from theshroud cavity 11 is discharged to the outside through theexhaust cavity 30 of thedischarge arrangement 23 within theshroud 8 and thestator vane 7. - Discharge is accomplished as per
FIG. 5 in that the static pressure pa applied to theexhaust cavity 30 in the form of theair discharge duct 25 is lower than the pressure pd prevailing in theshroud cavity 11. The static pressure pa must here be sufficiently lower than the pressure pa prevailing in theshroud cavity 11, that the desired amount of reduced mass flow is removed. - As shown in
FIG. 4 , this can for example be accomplished in that theexhaust cavity 30 is connected viasuitable return lines 27 to either anupstream compressor duct 26 or to asecondary air system 28 of the engine, with the removed air being used as turbine cooling air or for the pressurization of the aircraft cabin or other engine or aircraft systems, for example. - The amount of the removed, absolute leakage mass flow is controllable in that a defined throat 29 (
FIG. 4 ) is provided within theexhaust cavity 30 or within thedischarge arrangement 23. - Complete suppression of the
leakage flow 22 between the rearward and the forwardaxial gap area 40 of theshroud 8, be appropriately attuned. The required area 40 (FIG. 3 ) can be determined by way of the pressures p1 and p2 of the ingestedboundary layers shroud 8 and the exhaust mass flow (FIG. 5 ). - In accordance with the state of the art, the leakage mass flow through the
shroud cavity 11 with seals 16, 17 generally amounts to approx. 0.5-1 percent of the total flow. Consequently, theexhaust cavity 30 can be designed such that an amount of this size is exhausted, or, if used for example for pressurizing the aircraft cabin or for turbine cooling, it should be dimensioned such that the required mass flow, which normally amounts to approx. 2-4 percent, is discharged. - By way of the
exhaust cavity 30, both theboundary layer 31 before thestator 7 and theboundary layer 32 behind thestator 7 are ingested into theshroud cavity 30 and finally discharged to the outside (FIG. 5 ). This leads to a significant reduction of the stator losses and, thus, to an increase in compressor efficiency. With use being also made of the discharged air, further potential for lowering fuel consumption of the engine, is provided. - A variant of the air discharge is shown in
FIG. 6 . While in the arrangement according toFIGS. 3 to 5 theair discharge opening 24 is disposed in the radial end face of theshroud 8,FIG. 6 shows twoair discharge openings shroud 8, form theexhaust cavity 30 connected to theair discharge ducts 25 in theshroud 8. - The
exhaust cavity 30 itself can have any cross-sectional shape. SeeFIGS. 7 and 8 , for example, showing discrete polygonal, round orelliptical openings shroud 8. Within theshroud 8, theexhaust cavity 30 can either remain discrete or expand into a circumferential cavity. Passage through thestator vane 7 can be provided by one or several tubes asround hole 36,oblong hole 37 orslot 38, or by hollowing thestator vane 7 in the form of a hollow chamber 39 (FIGS. 9 a-c), for example. - The
shroud 8 without seals according to the present invention is significantly easier to design and manufacture than an arrangement with seals 16, 17 in accordance with the state of the art as perFIGS. 1 and 2 . This provides for weight and cost savings. Furthermore, the leakage flow, which entails both losses and heating of the fluid and the components, is nonexistent. Exhaustion of the air leads to ingestion of the boundary layer of the compressor side wall, which further contributes to a reduction of losses. Utilization of the exhausted air for secondary systems provides for a further increase in efficiency. This means increased compressor efficiency and reduced fuel consumption of the engine. In summary, an increase in stage efficiency of approx. 0.2 percent is obtainable. -
- 1 Rotor
- 2 Rotor shaft
- 3 Rotor hub
- 4 Rotor blade
- 5 Stator
- 6 Casing
- 7 Stator vane
- 8 Shroud
- 9 Sealing gap
- 10 Sealing arrangement
- 11 Shroud cavity
- 12 Axial gap
- 13 Axial gap
- 14 Clearance
- 15 a Sealing surface of
shroud 8 - 15 b Hub-side bottom of
clearance 14 - 16 Seal
- 17 Seal
- 18 Radial gap
- 19 Rotary direction
- 20 Sealing arrangement
- 21 Transfer duct flow
- 22 Leakage flow
- 23 Discharge arrangement
- 24 Air discharge opening
- 25 Air discharge duct
- 26 Compressor duct
- 27 Return line
- 28 Secondary airflow
- 29 Throat
- 30 Exhaust cavity
- 31 Boundary layer
- 32 Boundary layer
- 33 Polygonal opening
- 34 Round opening
- 35 Elliptical opening
- 36 Round hole
- 37 Oblong hole
- 38 Slot
- 39 Hollow chamber
- 40 Flow-wetted area
- 41 Air discharge opening
- 42 Air discharge opening
- D Pressure side
- S Suction side
Claims (13)
1. A compressor for a gas turbine, comprising:
a rotor hub carrying rotor blades;
a stator having stator vanes;
a shroud associated with the stator vanes; and
an arrangement providing sealing between the shroud and rotor hub to prevent leakage, the sealing arrangement being formed by a discharge arrangement for leakage air.
2. The compressor of claim 1 , wherein the discharge arrangement is an exhaust cavity positioned within the shroud and the stator vanes, to discharge the leakage air to the outside.
3. The compressor of claim 2 , wherein a static pressure applied to the exhaust cavity is lower than a pressure prevailing within a shroud cavity.
4. The compressor of claim 3 , further comprising a return line connecting the exhaust cavity to an upstream compressor duct.
5. The compressor of claim 3 , wherein the exhaust cavity is connected to a secondary air system of the engine.
6. The compressor of claim 5 , and further comprising a defined throat positioned within at least one of the exhaust cavity and the air discharge duct of the discharge arrangement to regulate flow.
7. The compressor of claim 6 , wherein a flow-wetted area of an air discharge opening of the air discharge arrangement is determined by pressures of boundary layers ingested before and behind the shroud or in the shroud cavity, respectively.
8. The compressor of claim 7 , wherein at least one of the exhaust cavity and a cross-section of the air discharge opening in the shroud is provided as a discrete polygonal, round or elliptical opening.
9. The compressor of claim 7 , wherein the exhaust cavity in one of the stator vanes is provided as at least one of a round hole, an oblong hole, a slot and a hollow chamber.
10. The compressor of claim 1 , wherein a static pressure applied to the exhaust cavity is lower than a pressure prevailing within a shroud cavity.
11. The compressor of claim 1 , and further comprising a return line connecting the exhaust cavity to an upstream compressor duct.
12. The compressor of claim 1 , wherein the exhaust cavity is connected to a secondary air system of the engine.
13. The compressor of claim 1 , and further comprising a defined throat positioned within at least one of the exhaust cavity and the air discharge duct of the discharge arrangement to regulate flow.
Applications Claiming Priority (2)
Application Number | Priority Date | Filing Date | Title |
---|---|---|---|
DE102008052101.9 | 2008-10-20 | ||
DE102008052101A DE102008052101A1 (en) | 2008-10-20 | 2008-10-20 | Compressor for a gas turbine |
Publications (1)
Publication Number | Publication Date |
---|---|
US20100098530A1 true US20100098530A1 (en) | 2010-04-22 |
Family
ID=41268433
Family Applications (1)
Application Number | Title | Priority Date | Filing Date |
---|---|---|---|
US12/576,762 Abandoned US20100098530A1 (en) | 2008-10-20 | 2009-10-09 | Compressor for a gas turbine |
Country Status (3)
Country | Link |
---|---|
US (1) | US20100098530A1 (en) |
EP (1) | EP2177770A2 (en) |
DE (1) | DE102008052101A1 (en) |
Cited By (1)
Publication number | Priority date | Publication date | Assignee | Title |
---|---|---|---|---|
US20180080476A1 (en) * | 2016-09-19 | 2018-03-22 | United Technologies Corporation | Geared turbofan front center body thermal management |
Families Citing this family (1)
Publication number | Priority date | Publication date | Assignee | Title |
---|---|---|---|---|
DE102017222210A1 (en) * | 2017-12-07 | 2019-06-13 | MTU Aero Engines AG | Compressor module for a turbomachine |
Citations (3)
Publication number | Priority date | Publication date | Assignee | Title |
---|---|---|---|---|
US3846038A (en) * | 1971-12-27 | 1974-11-05 | Onera (Off Nat Aerospatiale) | Fixed blading of axial compressors |
US5328326A (en) * | 1991-04-19 | 1994-07-12 | Gec Alsthom Sa | Impulse turbine with a drum rotor, and improvements to such turbines |
US20060222485A1 (en) * | 2004-09-30 | 2006-10-05 | Snecma | Method for air circulation in a turbomachine compressor, compressor arrangement using this method, compression stage and compressor incorporating such a arrangement, and aircraft engine equipped with such a compressor |
Family Cites Families (6)
Publication number | Priority date | Publication date | Assignee | Title |
---|---|---|---|---|
DE3225208C1 (en) * | 1982-06-29 | 1983-12-22 | Gerhard Dipl.-Ing. 7745 Schonach Wisser | Rotor wheel arrangement of a turbomachine with a shroud band |
DE3523469A1 (en) * | 1985-07-01 | 1987-01-08 | Bbc Brown Boveri & Cie | Contact-free controlled-gap seal for turbo-machines |
US4721313A (en) * | 1986-09-12 | 1988-01-26 | Atlas Copco Comptec, Inc. | Anti-erosion labyrinth seal |
JP2001504564A (en) * | 1996-09-26 | 2001-04-03 | シーメンス アクチエンゲゼルシヤフト | Method for compensating pressure loss caused by cooling air guide in gas turbine equipment |
IT1318065B1 (en) | 2000-06-29 | 2003-07-21 | Nuovo Pignone Spa | SEALING AND PRESSURIZATION SYSTEM FOR THE BEARING OF A GAS TURBINE |
DE10355240A1 (en) * | 2003-11-26 | 2005-07-07 | Rolls-Royce Deutschland Ltd & Co Kg | Fluid flow machine with fluid removal |
-
2008
- 2008-10-20 DE DE102008052101A patent/DE102008052101A1/en not_active Withdrawn
-
2009
- 2009-09-11 EP EP09170012A patent/EP2177770A2/en not_active Withdrawn
- 2009-10-09 US US12/576,762 patent/US20100098530A1/en not_active Abandoned
Patent Citations (3)
Publication number | Priority date | Publication date | Assignee | Title |
---|---|---|---|---|
US3846038A (en) * | 1971-12-27 | 1974-11-05 | Onera (Off Nat Aerospatiale) | Fixed blading of axial compressors |
US5328326A (en) * | 1991-04-19 | 1994-07-12 | Gec Alsthom Sa | Impulse turbine with a drum rotor, and improvements to such turbines |
US20060222485A1 (en) * | 2004-09-30 | 2006-10-05 | Snecma | Method for air circulation in a turbomachine compressor, compressor arrangement using this method, compression stage and compressor incorporating such a arrangement, and aircraft engine equipped with such a compressor |
Cited By (1)
Publication number | Priority date | Publication date | Assignee | Title |
---|---|---|---|---|
US20180080476A1 (en) * | 2016-09-19 | 2018-03-22 | United Technologies Corporation | Geared turbofan front center body thermal management |
Also Published As
Publication number | Publication date |
---|---|
DE102008052101A1 (en) | 2010-04-22 |
EP2177770A2 (en) | 2010-04-21 |
Similar Documents
Publication | Publication Date | Title |
---|---|---|
CN113356936B (en) | Seal assembly for counter-rotating turbine assembly and method of operating the same | |
EP1793086B1 (en) | Aerofoil for a gas turbine comprising a particle deflector | |
US8939711B2 (en) | Outer rim seal assembly in a turbine engine | |
US6438941B1 (en) | Bifurcated splitter for variable bleed flow | |
CN107044447B (en) | De-icing device for the split nose of an axial turbomachine compressor | |
EP2964960B1 (en) | Gas turbine engine centrifugal compressor with seal between two diffuser parts | |
US10669893B2 (en) | Air bearing and thermal management nozzle arrangement for interdigitated turbine engine | |
US8235654B2 (en) | Compressor stator with partial shroud | |
US10107109B2 (en) | Gas turbine engine component cooling assembly | |
US11788415B2 (en) | Shroudless blade for a high-speed turbine stage | |
JP2007120501A (en) | Interstage seal, turbine blade, and interface seal between cooled rotor and stator of gas turbine engine | |
EP3095958B1 (en) | System for thermally shielding a portion of a gas turbine shroud assembly | |
US20180230839A1 (en) | Turbine engine shroud assembly | |
EP2551458A2 (en) | Blade Cooling and Sealing System | |
CA2964988C (en) | Assembly and method for influencing flow through a fan of a gas turbine engine | |
US11933193B2 (en) | Turbine engine with an airfoil having a set of dimples | |
US20210189883A1 (en) | Blade of steam turbine and steam turbine | |
US20100098530A1 (en) | Compressor for a gas turbine | |
EP3287605B1 (en) | Rim seal for gas turbine engine | |
EP3492706B1 (en) | Gas turbine engine having a tip clearance control system | |
US12018582B2 (en) | Turbine blade for an aircraft turbine engine, comprising a platform provided with a channel for primary flow rejection towards a purge cavity | |
US20230073422A1 (en) | Stator with depressions in gaspath wall adjacent trailing edges | |
CN117145634A (en) | Anti-icing system, compressor and turbofan engine | |
CN115885091A (en) | Turbine with pressurized cavity | |
US20200165968A1 (en) | Fan assembly having flow recirculation circuit with rotating airfoils |
Legal Events
Date | Code | Title | Description |
---|---|---|---|
AS | Assignment |
Owner name: ROLLS-ROYCE DEUTSCHLAND LTD & CO KG,GERMANY Free format text: ASSIGNMENT OF ASSIGNORS INTEREST;ASSIGNORS:CLEMEN, CARSTEN;ORTMANNS, JENS;SIGNING DATES FROM 20090923 TO 20090924;REEL/FRAME:023357/0954 |
|
STCB | Information on status: application discontinuation |
Free format text: ABANDONED -- FAILURE TO RESPOND TO AN OFFICE ACTION |