EP1828616A1 - Verfahren zur modifizierung eines turbokompressors - Google Patents
Verfahren zur modifizierung eines turbokompressorsInfo
- Publication number
- EP1828616A1 EP1828616A1 EP05815881A EP05815881A EP1828616A1 EP 1828616 A1 EP1828616 A1 EP 1828616A1 EP 05815881 A EP05815881 A EP 05815881A EP 05815881 A EP05815881 A EP 05815881A EP 1828616 A1 EP1828616 A1 EP 1828616A1
- Authority
- EP
- European Patent Office
- Prior art keywords
- blade
- blades
- turbocompressor
- additional
- row
- 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.)
- Granted
Links
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 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.
- Gas turbine groups must ensure stable operation over a very wide operating range. This is partly due to the wide range of environmental conditions. On the other hand, it is not possible with single-shaft gas turbine groups, which serve to generate electricity, to control the operating state of the compressor independently of the rotational speed. On the contrary, 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. Also, over the life of a compressor, aging, wear and contamination of the compressor blades can lead to a deterioration of the stability behavior. In installations where such unfavorable conditions are cumulative and accumulated, it is often desirable to modify the compressor in a gas turbogenet as such to provide increased stability against stalling.
- US 6379112 discloses the technical teaching to prevent the risk of vibrations in the blade grid in the design of the machine that a blade row, be it a blade row in the rotor or a row of stator blades in the stator , over their scope in different Segments, preferably in four quadrants, divided, and within these segments each have a different number of blades is installed. Due to the changing blade distances over the circumference of the blades of the respective subsequent, passing blade row are excited with constantly changing frequencies, which reduces their tendency to oscillate. Such a configuration of vibration-prone areas within the blade lattice should reduce the tendency to oscillate as a whole. With regard to the problem explained at the beginning of an adaptation of the turbomachine to specific boundary conditions of an operation on the edge of the stability limit, however, this publication does not provide any suggestions.
- the invention makes it possible, the division ratio of an axial
- the pitch ratio of the blade grid of an axial blade row or a blade ring is defined as the mutual offset of two blades in the circumferential direction, based on the chord length of an airfoil. It is known that, depending on the grating load characteristic of the blade grid, there is an optimum division ratio in which the losses are minimized. Deviations from this optimum division ratio lead to a rapid increase in lattice losses. Modern turbocompressors are by their design in the range of this optimal division ratio. It has now been recognized within the scope of the invention that a reduction in the dividing ratio of an existing compressor can improve the flow stability within the compressor. This change, however, is a profound intervention in the grid geometry of an axial blade row.
- the proposed method can be realized when arranged in a blade ring blades are arranged with their blade roots in a circumferentially extending groove of a rotor shaft or a housing.
- At least one arranged in the circumferential direction between two blade roots of the blade ring spacer is removed and used in its place at least one additional blade.
- 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.
- this modification is based only on the replacement or the omission or optionally a post-processing components relatively easy to handle components.
- this type of modification of blade rings, in which spacers are arranged in the circumferential direction between blade roots particularly economical.
- An alternatively or cumulatively to be used process variant is characterized in that at least one existing blades is replaced with a blade whose blade root has a lower circumferential extent, and at least insert an additional blade in 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 kept low because not all blades of the modified blade ring must be replaced depending on the specific circumstances, 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 modified individually and on site 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 are 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 specified here is particularly suitable for the modification of at least one blade row of the rotor and / or the stator of a turbo compressor.
- 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.
- FIG. 1 shows a gas subgroup
- Figure 2 shows a stage of an axial turbocompressor with arranged in a circumferential groove compressor blades
- FIG. 3 shows the stage from FIG. 2 in another view
- Figure 4 is a tabular summary of the modifications of a 17-stage axial turbocompressor.
- FIG. 1 shows 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 is a part of a rotor of a compressor with a
- 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.
- the detailed arrangement of blade roots 231 and spacers 24 in the circumferential groove and the blades 232 is shown in the development shown in Figure 3.
- 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. In the method proposed here, use is now made of the knowledge that by means of a reduction of the division ratio, the tendency to detach a turbocompressor can be reduced and its operating range can be broadened.
- 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.
- 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.
- the spacers 24 are removed, the blades are pushed together, and additional blades are arranged.
- An embodiment of the method comprises increasing the number of blades from 41 to 45 in the third row of an axial turbocompressor.
- An embodiment of the method comprises 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 from 41 to 45 in the fifth row of an axial turbocompressor.
- An embodiment of the method comprises increasing the number of blades in the 6th 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 from 51 to 57 in the 8th row of an axial turbocompressor.
- 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 from 65 to 71 in the 10th row of an axial turbocompressor.
- An embodiment of the method comprises increasing the number of blades from 65 to 71 in the 11th row of an axial turbocompressor.
- An 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 increasing the number of blades from 65 to 71 in the 13th row of an axial turbocompressor.
- An embodiment of the method comprises increasing the number of blades from 83 to 91 in the 14th row of an axial turbocompressor.
- An embodiment of the method comprises 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.
- An embodiment of the method comprises increasing the number of blades from 83 to 91 in the 17th row of an axial turbocompressor.
- An embodiment of the method comprises, in the first guide row of an axial turbocompressor, increasing the number of blades from 34 to 38.
- the first guide row is different from a Vorleit Herbert; under the first guide row is to be understood the guide vane row, which is arranged immediately downstream of the first blade row.
- An embodiment of the method comprises, in the second guide row of an axial turbocompressor, increasing the number of blades from 46 to 50.
- An embodiment of the method comprises, in the third guide row of an axial turbocompressor, increasing the number of blades from 52 to 54.
- An embodiment of the method comprises, in the fourth guide row of an axial turbocompressor, increasing the number of blades from 52 to 54.
- An embodiment of the method comprises, in the fifth guide row of an axial turbocompressor, increasing the number of blades from 60 to 64.
- An embodiment of the method comprises, in the sixth guide row of an axial turbocompressor, increasing the number of blades from 56 to 62.
- An embodiment of the method comprises increasing the number of blades in the seventh guide row of an axial turbocompressor from 52 to 58.
- An embodiment of the method comprises increasing the number of blades from 66 to 72 in the eighth guide row of an axial turbocompressor.
- 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 comprises increasing the number of blades from 66 to 72 in the 10th row of an axial turbocompressor.
- An embodiment of the method comprises increasing the number of blades from 66 to 72 in the 11th row of an axial turbocompressor.
- An embodiment of the method comprises increasing the number of blades from 66 to 72 in the 12th row of an axial turbocompressor.
- An embodiment of the method comprises increasing the number of blades in the 13th row of an axial turbocompressor from 84 to 92.
- An embodiment of the method comprises, in the 14th guide row of an axial turbocompressor, increasing the number of blades from 84 to 92.
- An embodiment of the method comprises, in the fifteenth row of an axial turbocompressor, increasing the number of blades from 84 to 92.
- An embodiment of the method comprises increasing the number of blades in the 16th row of an axial turbocompressor from 84 to 92.
- An embodiment of the method comprises, in the 17th guide row of an axial turbocompressor, increasing the number of blades from 84 to 92.
- the blade geometry in each case preferably remains 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.
- No denotes the number of blades in a blade ring before modification.
- Ni 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.
Landscapes
- Engineering & Computer Science (AREA)
- Mechanical Engineering (AREA)
- General Engineering & Computer Science (AREA)
- Structures Of Non-Positive Displacement Pumps (AREA)
Abstract
Description
Claims
Priority Applications (1)
Application Number | Priority Date | Filing Date | Title |
---|---|---|---|
EP05815881.7A EP1828616B1 (de) | 2004-12-21 | 2005-11-29 | Verfahren zur modifizierung eines turbokompressors |
Applications Claiming Priority (3)
Application Number | Priority Date | Filing Date | Title |
---|---|---|---|
EP04106808A EP1674734A1 (de) | 2004-12-21 | 2004-12-21 | Verfahren zur Verbesserung der Strömungsstabilität eines Turbokompressors |
EP05815881.7A EP1828616B1 (de) | 2004-12-21 | 2005-11-29 | Verfahren zur modifizierung eines turbokompressors |
PCT/EP2005/056294 WO2006067025A1 (de) | 2004-12-21 | 2005-11-29 | Verfahren zur modifizierung eines turbokompressors |
Publications (2)
Publication Number | Publication Date |
---|---|
EP1828616A1 true EP1828616A1 (de) | 2007-09-05 |
EP1828616B1 EP1828616B1 (de) | 2014-10-01 |
Family
ID=34930104
Family Applications (2)
Application Number | Title | Priority Date | Filing Date |
---|---|---|---|
EP04106808A Withdrawn EP1674734A1 (de) | 2004-12-21 | 2004-12-21 | Verfahren zur Verbesserung der Strömungsstabilität eines Turbokompressors |
EP05815881.7A Not-in-force EP1828616B1 (de) | 2004-12-21 | 2005-11-29 | Verfahren zur modifizierung eines turbokompressors |
Family Applications Before (1)
Application Number | Title | Priority Date | Filing Date |
---|---|---|---|
EP04106808A Withdrawn EP1674734A1 (de) | 2004-12-21 | 2004-12-21 | Verfahren zur Verbesserung der Strömungsstabilität eines Turbokompressors |
Country Status (3)
Country | Link |
---|---|
US (1) | US20080003098A1 (de) |
EP (2) | EP1674734A1 (de) |
WO (1) | WO2006067025A1 (de) |
Families Citing this family (3)
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)
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 |
-
2004
- 2004-12-21 EP EP04106808A patent/EP1674734A1/de not_active Withdrawn
-
2005
- 2005-11-29 WO PCT/EP2005/056294 patent/WO2006067025A1/de active Search and Examination
- 2005-11-29 EP EP05815881.7A patent/EP1828616B1/de not_active Not-in-force
-
2007
- 2007-06-14 US US11/812,029 patent/US20080003098A1/en not_active Abandoned
Non-Patent Citations (1)
Title |
---|
See references of WO2006067025A1 * |
Also Published As
Publication number | Publication date |
---|---|
EP1674734A1 (de) | 2006-06-28 |
WO2006067025A1 (de) | 2006-06-29 |
EP1828616B1 (de) | 2014-10-01 |
US20080003098A1 (en) | 2008-01-03 |
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