EP1828616B1 - Méthode d' amélioration de la stabilité de courant d' une turbomachine - Google Patents

Méthode d' amélioration de la stabilité de courant d' une turbomachine Download PDF

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Publication number
EP1828616B1
EP1828616B1 EP05815881.7A EP05815881A EP1828616B1 EP 1828616 B1 EP1828616 B1 EP 1828616B1 EP 05815881 A EP05815881 A EP 05815881A EP 1828616 B1 EP1828616 B1 EP 1828616B1
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EP
European Patent Office
Prior art keywords
blade
blades
additional
row
turbocompressor
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.)
Not-in-force
Application number
EP05815881.7A
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German (de)
English (en)
Other versions
EP1828616A1 (fr
Inventor
Marco Micheli
Current Assignee (The listed assignees may be inaccurate. Google has not performed a legal analysis and makes no representation or warranty as to the accuracy of the list.)
General Electric Technology GmbH
Original Assignee
Alstom Technology AG
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Filing date
Publication date
Application filed by Alstom Technology AG filed Critical Alstom Technology AG
Priority to EP05815881.7A priority Critical patent/EP1828616B1/fr
Publication of EP1828616A1 publication Critical patent/EP1828616A1/fr
Application granted granted Critical
Publication of EP1828616B1 publication Critical patent/EP1828616B1/fr
Not-in-force legal-status Critical Current
Anticipated expiration legal-status Critical

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Classifications

    • 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

Definitions

  • the invention relates to a method for the individual modification of a turbocompressor for the purpose of adaptation to specific operating conditions.
  • the method can be used in particular in multi-stage axial turbocompressors.
  • the invention relates to components of a turbo-compressor, which are modified by means of the specified method, and to a turbo-compressor, and to a gas turbine group, which include such a turbo-compressor.
  • Compressors of gas turbine groups and in particular air-breathing gas turbine groups must ensure stable operation over a very wide operating range. This is partly due to the wide range of environmental conditions.
  • single-shaft gas turbine groups which serve to generate electricity, to control the operating state of the compressor independently of the rotational speed.
  • machines that are installed in weak networks as they typically occur in third world countries, run with large set by the network frequency and thus speed fluctuations and often with stability-deteriorating underspeed.
  • aging, wear and contamination of the compressor blades can lead to a deterioration of the stability behavior.
  • the compressor When designing gas turbine groups, the compressor must be able to work with very good efficiencies. However, this may result in the compressor under extreme conditions, such as For example, is operated at close to the stability limit due to the prescribed by the turbine increasing pressure ratio close to the stability limit, or at the same time that very high compressor inlet temperatures shift the operating point of the compressor in a stable only conditionally stable at per se stability improving lower compressor inlet temperatures. The stability reserve may then no longer be sufficient to absorb further unfavorable factors mentioned above.
  • the invention makes it possible to change the division ratio of an axial blade grid without changing the geometry of the blades and in particular without changing components such as the rotor shaft or the housing.
  • the pitch ratio of the blade lattice of an axial blade row or a blade ring is defined as the mutual offset of two blade leaves in the circumferential direction, based on the chord length of an airfoil.
  • a method for improving the flow stability of a turbocompressor, in which the number of blades arranged in a blade row or in a blade ring is increased in an axial blade grid.
  • the distance between two blades is reduced, and the pitch ratio is reduced even without changing the blade geometry, that is, although to greater lattice losses, but also shifted to improved stability.
  • the proposed method can be realized in a particularly simple manner if the blades arranged in a blade ring are arranged with their blade roots in a circumferential groove of a rotor shaft or of a housing.
  • At least one spacer arranged in the circumferential direction between two blade roots of the blade ring is removed, and in its place at least one additional blade is inserted.
  • at least one arranged between two blade roots of the blade ring spacer replaced with a spacer with a smaller circumferential extent and installed at least one additional blade.
  • at least one existing spacer is removed and machined so that the circumferential extent of the spacer is reduced; the thus modified spacer is reinstalled, and at least one additional blade is inserted into the blade ring.
  • An alternative or cumulative method variant to be used is characterized in that at least one existing blade is replaced by a blade whose blade root has a smaller circumferential extent, and to insert at least one additional blade into the blade ring.
  • This method may also include removing an existing blade and machining its blade root so as to reduce the circumferential extent of the blade root. The blade is reinstalled after the modification of the blade root together with an additional blade.
  • blades are used as an additional blade and / or optionally as blades to be installed, whose blade blade has the same chord length and in particular the same blade geometry as the originally installed blades.
  • a modification of the blade root is also made on these blades, which are to be additionally integrated, and the circumferential extent of the blade root is reduced compared to the original condition.
  • the only grating characteristic that is changed is the division ratio.
  • the effort that is necessary for the modification is kept low because, depending on the specific situation, not all blades of the modified blade ring must be replaced, and a new interpretation of the blade geometry is eliminated.
  • the effects on the flow conditions of the downstream blade edge remain minimal.
  • blades with identical chord length of the blade and in particular with identical blade geometry are arranged in the entire blade ring.
  • a major advantage of the method described here is the fact that compressors can be individually and locally modified to be adapted to specific conditions during operation, without modifications to large and difficult-to-use components such as the rotor shaft and / or the Housing required.
  • the method can be carried out in such a way that only standard components with, if necessary, easy-to-carry out modifications, such as, for example, easy milling of the blade root, are required.
  • not all blades of the blade ring necessarily be replaced, but it may be necessary to re-deliver the additional blades to be installed, which greatly simplifies logistics, especially for installations in hard to reach regions.
  • the method described above also makes it possible to optimize turbocompressors for normal use, since it is very easily possible by means of the described method to modify a compressor, which is intended for exceptional and extreme operating conditions, for a more stable operating behavior.
  • the method specified here is particularly suitable for modifying at least one blade row of the rotor and / or the stator of a turbocompressor.
  • the method is suitable for modifying a turbocompressor of a gas turbine group.
  • the invention also includes a rotor of a turbocompressor and / or a stator of a turbocompressor with at least one row of blades, which are modified according to the method described above. It also includes a turbocompressor, which comprises a rotor and / or stator modified by means of the method.
  • a gas turbine group 1 which comprises a compressor 2, a combustion chamber 3 and a turbine 4.
  • the illustrated gas turbine group is used to drive a generator 5 for power generation.
  • a rotor comprising a rotor shaft with blades and a stator with usually arranged in the housing vanes.
  • any frequency change of the network results directly in a speed change of the gas turbo group.
  • FIG. 2 a part of a rotor of a compressor is shown with a row of blades.
  • the rotor comprises the rotor shaft 21.
  • a circumferential groove 22 is incorporated, in which the blades 23 of the blade row are arranged.
  • a blade 23 comprises a blade root 231 and an airfoil 232.
  • blade roots 231 and spacers 24 are alternately arranged.
  • the blade ring includes N blades. With U, the circumferential direction of the blade ring is designated.
  • blade roots 231 and spacers 24 in the circumferential groove and the airfoils 232 is in the in FIG. 3 shown settlement.
  • the extent of a blade root in the circumferential direction is denoted by I.
  • the extension of a spacer in the circumferential direction is denoted by b.
  • These masses are indicated in each case on the outer circumference of the rotor shaft.
  • the blade pitch that is, the distance between two blades in the circumferential direction, is denoted by t.
  • this measure varies over the blade height;
  • the chord length of an airfoil is not necessarily constant over the entire blade height.
  • the division ratio t / s is, as is familiar to the expert, a decisive grating characteristic. With decreasing division ratio, the wall friction losses of the flow in the blade grid increase. With a very large division ratio, the losses increase due to the increasingly inefficient flow diversion. For compressors, moreover, the tendency for flow separation increases. In between, there exists an optimum division ratio, in which the losses are minimal, and to which a blade grid is usually at least approximately designed. The optimum division ratio is a function of a grid load characteristic and can be determined by a person skilled in the art without any problem.
  • turbocompressors which are used, for example, as compressors of gas turbine groups, with regard to the losses in such a way that they operate in a frequently occurring operating range with lowest losses, for example in more than 70 percent of all applications.
  • the Possibility to improve the operational stability of the turbocompressor without constructive changes makes it possible, inter alia, to make this design more uncompromising and to have to take less account of extreme operating conditions than hitherto usual.
  • the proposed method makes it possible to individually change individual compressors of a series individually compared to the standard design and adapt to specific operating conditions. For example, problems of stable operation are created by gas turbine group compressors which operate at very high ambient temperatures and which may need to operate at low speed in weak electricity networks. This is further accentuated when the gas turbine group is operated at high ambient temperatures to maintain power with water and / or steam injection into the combustion chamber, thereby also increasing the pressure ratio against which the compressor must operate. According to the method proposed here, the stable operating range of the compressor is widened by increasing the number of blades arranged in the blade ring in at least one row of blades and thus reducing the dividing ratio. At the in FIG.
  • An embodiment of the method comprises increasing the number of blades from 41 to 45 in the third row of flights of an axial turbocompressor.
  • One embodiment of the method includes increasing the number of blades from 41 to 45 in the fourth row of an axial turbocompressor.
  • An embodiment of the method comprises increasing the number of blades in the fifth row of flights of an axial turbocompressor from 41 to 45.
  • An embodiment of the method comprises increasing the number of blades in the sixth row of an axial turbocompressor from 51 to 57.
  • An embodiment of the method comprises increasing the number of blades in the 7th row of an axial turbocompressor from 51 to 57.
  • An embodiment of the method comprises increasing the number of blades in the 8th row of an axial turbocompressor from 51 to 57.
  • An embodiment of the method comprises increasing the number of blades in the ninth row of an axial turbocompressor from 65 to 71.
  • An embodiment of the method comprises increasing the number of blades in the 10th row of an axial turbocompressor from 65 to 71.
  • One embodiment of the method comprises increasing the number of blades from 65 to 71 in the 11th row of an axial turbocompressor.
  • One embodiment of the method comprises increasing the number of blades in the 12th row of an axial turbocompressor from 65 to 71.
  • An embodiment of the method comprises, in the 13th row of an axial turbocompressor, increasing the number of blades from 65 to 71.
  • One embodiment of the method comprises increasing the number of blades from 83 to 91 in the 14th row of an axial turbocompressor.
  • One embodiment of the method includes increasing the number of blades from 83 to 91 in the 15th row of an axial turbocompressor.
  • An embodiment of the method comprises increasing the number of blades in the 16th row of an axial turbocompressor from 83 to 91.
  • One embodiment of the method comprises increasing the number of blades in the 17th row of an axial turbocompressor from 83 to 91.
  • One embodiment of the method comprises increasing the number of blades from 34 to 38 in the first guide row of an axial turbocompressor.
  • the first guide row is different from a Vorleit Herbert; the first guide row is understood to mean the row of guide vanes which is arranged immediately downstream of the first row of blades.
  • One embodiment of the method comprises increasing the number of blades from 46 to 50 in the second guide row of an axial turbocompressor.
  • One embodiment of the method comprises increasing the number of blades from 52 to 54 in the third guide row of an axial turbocompressor.
  • One embodiment of the method comprises increasing the number of blades from 52 to 54 in the fourth guide row of an axial turbocompressor.
  • An embodiment of the method comprises increasing the number of blades from 60 to 64 in the fifth guide row of an axial turbocompressor.
  • One embodiment of the method includes increasing the number of blades from 56 to 62 in the sixth row of axial turbocompressors.
  • An embodiment of the method comprises increasing the number of blades in the 7th row of an axial turbocompressor from 52 to 58.
  • One embodiment of the method comprises increasing the number of blades in the eighth guide row of an axial turbocompressor from 66 to 72.
  • An embodiment of the method comprises, in the ninth row of an axial turbocompressor, increasing the number of blades from 66 to 72.
  • An embodiment of the method includes increasing the number of blades from 66 to 72 in the 10th row of an axial turbocompressor.
  • One embodiment of the method includes increasing the number of blades from 66 to 72 in the 11th row of an axial turbocompressor.
  • One embodiment of the method comprises increasing the number of blades from 66 to 72 in the 12th row of an axial turbocompressor.
  • One embodiment of the method includes increasing the number of blades from 84 to 92 in the 13th row of an axial turbocompressor.
  • One embodiment of the method includes increasing the number of blades from 84 to 92 in the 14th row of an axial turbocompressor.
  • One embodiment of the method comprises increasing the number of blades in the 15th row of an axial turbocompressor from 84 to 92.
  • One embodiment of the method includes increasing the number of blades in the 16th order of an axial turbocompressor from 84 to 92.
  • One embodiment of the method includes increasing the number of blades in the 17th order of an axial turbocompressor from 84 to 92.
  • the blade geometry remains preferably unchanged in these modifications.
  • the originally installed blades are reused and additional blades are newly installed.
  • a 17-stage axial turbocompressor is modified according to a method characterized in the claims.
  • the top line indicates the number of the level.
  • LE denotes the guide rows
  • LA denotes the rows of runs.
  • N 0 denotes the number of blades in a blade ring before modification.
  • N 1 denotes the number of blades in a blade ring after the modification. In this way, the modification made can be read from the comparison of the second and third as well as the fourth and fifth line.

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  • Engineering & Computer Science (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Structures Of Non-Positive Displacement Pumps (AREA)

Claims (10)

  1. Procédé de modification individuelle d'un turbocompresseur (2) à des fins d'adaptation à des conditions de base spécifiques lors du fonctionnement, des aubes (23) disposées dans une couronne d'aubes étant disposées de manière à ce que leurs pieds d'aube (231) soient agencés dans une rainure (22), s'étendant dans la direction périphérique, d'un arbre de rotor ou d'un carter, caractérisé par l'augmentation, dans une grille d'aubes axiale du compresseur, du nombre (N) des aubes (23) disposées dans une couronne d'aubes, soit au moins une pièce d'écartement (24) disposée entre deux pieds d'aube (231) de la couronne d'aubes dans la direction périphérique (U) étant retirée ou remplacée par une pièce d'écartement présentant une étendue périphérique plus petite (b), soit au moins une aube présente (23) étant remplacée par une aube dont le pied d'aube (231) présente une plus petite étendue périphérique (1), et au moins une aube supplémentaire étant insérée.
  2. Procédé selon la revendication 1, caractérisé par l'enlèvement d'au moins une pièce d'écartement présente (24), l'usinage de celle-ci, de telle sorte que l'étendue périphérique (b) soit réduite, et l'installation à nouveau de la pièce d'écartement modifiée et d'au moins une aube supplémentaire.
  3. Procédé selon la revendication 1 ou 2, caractérisé par le fait de laisser les aubes installées à l'origine ou d'enlever et d'installer à nouveau de manière identique celles-ci ou de remplacer les aubes installées par des aubes identiques.
  4. Procédé selon la revendication 1, caractérisé par le remplacement d'une aube présente (23) par une aube dont le pied d'aube (231) présente une plus petite étendue périphérique (1), du fait qu'au moins une aube présente (23) est enlevée, et que son pied d'aube (231) est usiné de telle sorte que l'étendue périphérique (1) soit réduite, et que l'aube est à nouveau installée conjointement avec une aube supplémentaire.
  5. Procédé selon l'une quelconque des revendications précédentes, caractérisé par l'utilisation, en tant qu'aube supplémentaire, d'une aube dont la pale d'aube (232) présente la même longueur de corde (s) que les aubes installées à l'origine.
  6. Procédé selon l'une quelconque des revendications précédentes, caractérisé par l'utilisation, en tant qu'aube supplémentaire, d'une aube qui présente la même géométrie de pale d'aube que les aubes installées à l'origine.
  7. Procédé selon l'une quelconque des revendications précédentes, caractérisé par l'utilisation, pour l'ensemble de la couronne d'aubes, d'aubes présentant une longueur de corde identique (s) de la pale d'aube (232).
  8. Procédé selon l'une quelconque des revendications précédentes, caractérisé par l'utilisation, pour l'ensemble de la couronne d'aubes, d'aubes présentant une géométrie de pale d'aube identique.
  9. Procédé selon l'une quelconque des revendications précédentes, caractérisé par l'installation, en tant qu'aube supplémentaire, d'une aube identique à une aube installée à l'origine.
  10. Procédé selon l'une quelconque des revendications précédentes, caractérisé par l'installation, en tant qu'aube supplémentaire, d'une aube identique à une aube installée à l'origine, l'usinage de son pied d'aube de telle sorte que l'étendue périphérique soit réduite, et l'installation de l'aube modifiée de la sorte.
EP05815881.7A 2004-12-21 2005-11-29 Méthode d' amélioration de la stabilité de courant d' une turbomachine Not-in-force EP1828616B1 (fr)

Priority Applications (1)

Application Number Priority Date Filing Date Title
EP05815881.7A EP1828616B1 (fr) 2004-12-21 2005-11-29 Méthode d' amélioration de la stabilité de courant d' une turbomachine

Applications Claiming Priority (3)

Application Number Priority Date Filing Date Title
EP04106808A EP1674734A1 (fr) 2004-12-21 2004-12-21 Méthode d' amélioration de la stabilité de courant d' une turbomachine
EP05815881.7A EP1828616B1 (fr) 2004-12-21 2005-11-29 Méthode d' amélioration de la stabilité de courant d' une turbomachine
PCT/EP2005/056294 WO2006067025A1 (fr) 2004-12-21 2005-11-29 Procede de modification d'un turbocompresseur

Publications (2)

Publication Number Publication Date
EP1828616A1 EP1828616A1 (fr) 2007-09-05
EP1828616B1 true EP1828616B1 (fr) 2014-10-01

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Family Applications (2)

Application Number Title Priority Date Filing Date
EP04106808A Withdrawn EP1674734A1 (fr) 2004-12-21 2004-12-21 Méthode d' amélioration de la stabilité de courant d' une turbomachine
EP05815881.7A Not-in-force EP1828616B1 (fr) 2004-12-21 2005-11-29 Méthode d' amélioration de la stabilité de courant d' une turbomachine

Family Applications Before (1)

Application Number Title Priority Date Filing Date
EP04106808A Withdrawn EP1674734A1 (fr) 2004-12-21 2004-12-21 Méthode d' amélioration de la stabilité de courant d' une turbomachine

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US (1) US20080003098A1 (fr)
EP (2) EP1674734A1 (fr)
WO (1) WO2006067025A1 (fr)

Families Citing this family (3)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
WO2008101805A2 (fr) * 2007-02-05 2008-08-28 Az Technologies Sas Machine hydraulique modulaire
US20120099995A1 (en) 2010-10-20 2012-04-26 General Electric Company Rotary machine having spacers for control of fluid dynamics
CH704212A1 (de) * 2010-12-15 2012-06-15 Alstom Technology Ltd Axialkompressor.

Family Cites Families (16)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US1010630A (en) * 1908-05-26 1911-12-05 Colonial Trust Co Blade-holding means for turbines.
US927093A (en) * 1908-12-17 1909-07-06 Westinghouse Machine Co Elastic-fluid turbine.
AT76502B (de) * 1914-02-12 1919-05-26 Erste Bruenner Maschinen Fab Schaufelbefestigung für Dampf- und Gasturbinen.
US1156529A (en) * 1914-06-10 1915-10-12 Gen Electric Turbine bucket-wheel.
US1366605A (en) * 1919-06-27 1921-01-25 Gen Electric Blade-securing means and method of making the same
US1494781A (en) * 1921-05-31 1924-05-20 Westinghouse Electric & Mfg Co Blade fastening
US1590328A (en) * 1923-08-13 1926-06-29 Westinghouse Electric & Mfg Co Fastening means for turbine blading
US2857134A (en) * 1954-03-17 1958-10-21 Parsons C A & Co Ltd Assembly of blades for turbines and the like
GB777955A (en) * 1954-07-06 1957-07-03 Ruston & Hornsby Ltd Improvements in or relating to fluid flow machines such as hydraulic, steam or gas turbines or axial-flow compressors
US3032864A (en) * 1958-04-29 1962-05-08 Ford Motor Co Wheel manufacture
CH516731A (de) * 1969-12-12 1971-12-15 Bbc Sulzer Turbomaschinen Schaufelkranz für Turbomaschinen
IT1066506B (it) * 1973-12-17 1985-03-12 Seeber Willi Ruota a pale per ventilatori e procedimento per produrla
FR2416341A1 (fr) * 1978-02-02 1979-08-31 Messier Fa Generateur aero-hydraulique reversible et installations de recuperation de chaleur comprenant un tel generateur
US5232346A (en) * 1992-08-11 1993-08-03 General Electric Company Rotor assembly and platform spacer therefor
DE4436731A1 (de) * 1994-10-14 1996-04-18 Abb Management Ag Verdichter
US6379112B1 (en) * 2000-11-04 2002-04-30 United Technologies Corporation Quadrant rotor mistuning for decreasing vibration

Also Published As

Publication number Publication date
EP1674734A1 (fr) 2006-06-28
EP1828616A1 (fr) 2007-09-05
WO2006067025A1 (fr) 2006-06-29
US20080003098A1 (en) 2008-01-03

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