US3661477A - Heat shield - Google Patents

Heat shield Download PDF

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Publication number
US3661477A
US3661477A US885139A US3661477DA US3661477A US 3661477 A US3661477 A US 3661477A US 885139 A US885139 A US 885139A US 3661477D A US3661477D A US 3661477DA US 3661477 A US3661477 A US 3661477A
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US
United States
Prior art keywords
segments
segment
heat shield
blade wheel
shield
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.)
Expired - Lifetime
Application number
US885139A
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English (en)
Inventor
Thorbjorn Westrum
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.)
Kongsberg Gruppen ASA
Original Assignee
Kongsberg Vapenfabrikk AS
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Filing date
Publication date
Application filed by Kongsberg Vapenfabrikk AS filed Critical Kongsberg Vapenfabrikk AS
Application granted granted Critical
Publication of US3661477A publication Critical patent/US3661477A/en
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Expired - Lifetime legal-status Critical Current

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Classifications

    • 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
    • F01D25/00—Component parts, details, or accessories, not provided for in, or of interest apart from, other groups
    • F01D25/24—Casings; Casing parts, e.g. diaphragms, casing fastenings
    • F01D25/26—Double casings; Measures against temperature strain in casings
    • 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
    • F02C7/00—Features, components parts, details or accessories, not provided for in, or of interest apart form groups F02C1/00 - F02C6/00; Air intakes for jet-propulsion plants
    • F02C7/24—Heat or noise insulation
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
    • F05B—INDEXING SCHEME RELATING TO WIND, SPRING, WEIGHT, INERTIA OR LIKE MOTORS, TO MACHINES OR ENGINES FOR LIQUIDS COVERED BY SUBCLASSES F03B, F03D AND F03G
    • F05B2230/00—Manufacture
    • F05B2230/60—Assembly methods
    • F05B2230/604—Assembly methods using positioning or alignment devices for aligning or centering, e.g. pins
    • F05B2230/606—Assembly methods using positioning or alignment devices for aligning or centering, e.g. pins using maintaining alignment while permitting differential dilatation
    • 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
    • F05D2230/00—Manufacture
    • F05D2230/60—Assembly methods
    • F05D2230/64—Assembly methods using positioning or alignment devices for aligning or centring, e.g. pins
    • F05D2230/642—Assembly methods using positioning or alignment devices for aligning or centring, e.g. pins using maintaining alignment while permitting differential dilatation

Definitions

  • ABSTRACT A segmented annular heat shield for mounting between the turbine blade wheel and the compressor impeller of a radial gas turbine.
  • Each segment is fixedly secured to a stationary structure at one point at its outer edge and also fastened to said structure at two further fastening points at its inner edge in such a manner that at these further fastening points the segment is restrained from movement in the axial direction of the shield but can move freely in the plane of the segment.
  • the segments adjoin along joints of such a design as to admit of relative movement of the segments in the plane thereof due to heat expansion.
  • the present invention relates to an annular, stationary heat shield for mounting between two rotating members of different temperature for reducing the heat transfer between these members, which in particular may be the compressor impeller and the turbine blade wheel of a radial gas turbine.
  • this stationary structure In order to protect the compressor and prevent too large heat losses from the turbine, the insulating properties of this stationary structure can be enhanced by an annular, stationary heat shield located on the side facing the turbine, the hot gases in the latter moving along the heat shield, which is spaced from the remaining parts of the stationary structure so that a layer of comparatively stagnant gas acting as a heat insulation is provided between said structure and the heat shield,
  • the temperature gradients in the heat shield are very large both in the radial direction and in the axial direction, since on its outer side, i.e., the side facing the turbine, the heat shield is hotter at its outer periphery than at its inner periphery, and at the same time the temperature on the inner side of the heat shield is of course substantially lower than that on the outer side.
  • the object of the invention is to provide an annular heat shield of the type referred to, which requires considerably less room in the axial direction, is subjected to lower stresses than the known designs and can be used in radial gas turbines having a large diameter.
  • the shield according to the invention is primarily characterized in that it consists of segments.
  • the joints between the segments are so designed as to be substantially sealed, while at the same time admitting of heat expansion of each segment relatively to the adjacent segments.
  • each segment can move independently of the other segments and will not be subjected to stresses of the same magnitude as an integral ring.
  • each of the segments is of course much smaller relatively to the radius of curvature than an integral shield, whereby the deflection in the axial direction will be substantially smaller.
  • a heat shield consisting of separate segments according to the invention further makes mounting of the shield possible without the necessity of separating the compressor impeller and turbine blade wheel.
  • the sealing of the joints between the segments can be obtained by means of thin strips inserted in grooves in the opposed edges of the segments.
  • the invention also teaches a simple manner of securing the segments to the stationary structure behind them in a way that causes a minimum of stress in the segments and the fastening means.
  • each segment' is fixedly secured at one point only and restrained from movement in the axial direction of the ring at two further points.
  • the two points at which the segment is withheld in the axial direction only, are located at the inner edge of the segment, and the fastening means are so designed that they may be mounted in advance, so that the segments may be inserted radially between the compressor impeller and the turbine blade wheel without use of tools. This allows an especially compact assembly of the impeller and .blade wheel.
  • the segment can be secured without using axially extending bolts or similar fastening means which would be difficult to insert and lock, in view of the small room available between the impeller and blade wheel.
  • the segments may at their outer edge have an axially extending flange through which the fastening means used at the outer edge extends radially.
  • a fastening means for instance a bolt, can be inserted in the radial direction and tightened by means of simple tools.
  • FIG. 1 is a section of a sectional view through the axis of a radial gas turbine and illustrates the assembly of the heat shield according to the invention
  • FIG. 2 illustrates the segments of which the heat shield of the invention consists, as viewed from the rear
  • FIG. 3 illustrates a segment viewed from the end, i.e., from the line III-III in FIG. 2,
  • FIG. 4 is a cross section through the heat shield in FIG. 2 on the line IV-IV,
  • FIG. 5 is a cross section on the line VV in FIG. 2 through the heat shield adjacent a joint and with a sealing strip inserted, and
  • FIG. 6 is a cross section similar to that in FIG. 4, but on the line VI-VI in FIG. 2.
  • FIG. 1 The outer ends of a compressor blade wheel I having vanes 2 and a turbine impeller 3 having vanes 4, respectively, are shown in FIG. 1. From the compressor impeller l the compressed air further passes through diffuser rings, which are indicated at 5 and 6 and which are secured to a diffuser plate 7. From the combustion chamber the hot combustion gases enter the turbine through an inlet nozzle ring 8.
  • the impeller 1 and 3 are clamped together and blade wheel abut each other in a position that is not shown in FIG. 1. However, between the outer portions of the impeller and blade wheel there is provided an annular groove 9 into which a stationary annular structure 10 extends.
  • the structure 10 is secured to the diffuser plate 7 by bolts 11 and at its inner end carries sealing means 12 which together with the rear face of the rotating compressor impeller 1 forms a labyrinth seal.
  • the stationary structure 10 further carries an annular heat shield made up of segments 13 and spaced from the stationary structure 10 so that between this structure and the heat shield there is formed a substantially closed chamber 14 containing virtually stagnant gas serving as a heat insulation.
  • each gap or joint 17 will have a certain width, which for a ring having a diameter of about 600 mm may be approximately 3 mm.
  • Each segment 13 is substantially flat having, however, an enlarged inner margin 15 and an axially extending flange 16 at the outer edge.
  • An aperture 18 has been drilled in the flange I6 midway between the ends of the segment.
  • recesses 21 which are best illustrated in FIGS. 2 and 5, and which open towards inner side 22 of the segment.
  • each recess 21 provides an opening for a fastening bolt 23 that is illustrated in FIG. 1.
  • the recess 21 is enlarged as illustrated at 24 (FIG. to accommodate and provide an abutment for head 25 of the fastening bolt 23.
  • grooves 26 extending in a common plane parallel to the plane of the segments are provided in the opposed edges 20.0f the segments; in these grooves there may be inserted thin strips 27 of a width that is equal to or less than twice the depth of each groove 26 and greater than the depth of one groove plus the maximum spacing of the segments when cold.
  • the strips 27 will not prevent heat expansion of the segments 13, and on the other hand they will always seal the joints 17 and thereby prevent the hot gases passing along the front of the segments from penetrating into the chamber 14 and induce undesired movements and exchanges of gas therein.
  • the mounting of the segments 13 is best illustrated in FIG. 1, in which the outer end of the stationary structure and the segments 13 are illustrated in a sectional view corresponding to the line lV-lV in FIG. 2, whereas at their inner ends they are illustrated in a sectional view corresponding to the line Ill-ll] in H0. 2.
  • the bolts 23 for fastening the segments at the inner edges thereof are common to two adjacent segments.
  • the bolts maybe mounted in advance in the stationary structure 10 while inserting intermediate plate springs 28. Thereby it willbe possible to force the segments radially into position behind the bolt head 25, the recesses 21 surrounding the shaft of the bolts and the bolt head 25 being accommodated in the enlarged recess 24.
  • each segment 13 is secured at three points, viz. partly by the bolt 29, which secures the segment 13 in such a way that at the point of attachment the segment-can neither move in the axial direction nor in the radial direction relatively to the stationary structure 10, and partly by two bolts 23, which secure the segment against movement in the axial direction relatively to the stationary structure 10, but admit of movement in the radial direction.
  • the segments 13 When the segments 13 are heated in operation they will expand radially inwards and tangentially, and this expansion is possible because of the joints l7 and the way of securing the segmentsat their inner edges.
  • the novel design of the heat shield of the invention has rendered it feasible to design radial gas turbines having a relatively large diameter and having a compressor impeller and a turbine blade wheel assembled very closely together back to back.
  • a turbo-compressor unit comprising a compressor impeller, a turbine blade wheel axially aligned with the compressor impeller and connected thereto to drive the same, and a stationary supporting structure located between the impeller and blade wheel; an annular heat shield structure located in a substantially radial plane between the impeller and blade wheel and secured to said supporting structure, one side of said shield structure being subjected to hot gases in the turbine blade wheel during operation of the unit, said shield structure comprising at least three segments, said segments being individually secured to ,said supporting structure and each of said segments being substantially smaller than the assembled at least three segments, said segments being interconnected to define a substantially gas-impervious annular shield by means of elastic and sealing connecting means allowing each segment to perform, substantially independently of the adjacent segments, the deformations due 0 the increase in temperature and the temperature gradients in the segments during operation of the unit.
  • said elastic and sealing connecting means comprises a groove provided in each of the opposed edges of adjacent segments and a thin strip accommodated within the two grooves of each two opposed edges of adjacent segments, I and each strip having a width equal to or less'than twice the depth of each groove and greater than the depth of a groove plus the maximum distance between the opposed edges of adjacent segments when cold.
  • a turbo-compressor unit comprising a compressor im peller, a turbine blade wheel axially aligned with the compressor impellerand connected thereto to drive same, and a stationary supporting structure located between the impeller and blade wheel; an annular heat shield structure, located in a substantially radial plane between the impeller and blade wheel and secured to said supporting structure, one side of such shield structure being subjected to hot gases in the turbine blade wheel during operation of the unit, said shield structure comprising segments, said segments being individually secured to said supporting structure and each of said segments being substantially smaller than the assembled segments, said segments being interconnected to define a substantially gasimpervious annular shield by means of elastic and sealing connecting means allow each segment to perform, substantially independently of the adjacent segments, the defonnations due to the increase in temperature and the temperature gradients in the segments during operation of the unit, and each segment being fixedly secured at one point only and restrained from substantial movement in the axial direction of the turbocompressor unit at two further points.
  • fastening means for fastening a segment at each of the restraining points at the radially inner edge comprises a headed bolt extending at right angles to the plane of the segment and spring biased in its axial direction, and the inner edge of the segment being provided with a recess for the bolt.
  • each segment at its radially outer edge has an axially extending flange, fastening means for fastening the segment at at the fastening point located at the radially outer edge of the segment extending radially through said flange.

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  • Engineering & Computer Science (AREA)
  • Chemical & Material Sciences (AREA)
  • Combustion & Propulsion (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Turbine Rotor Nozzle Sealing (AREA)
  • Structures Of Non-Positive Displacement Pumps (AREA)
US885139A 1968-12-17 1969-12-15 Heat shield Expired - Lifetime US3661477A (en)

Applications Claiming Priority (1)

Application Number Priority Date Filing Date Title
NO5037/68A NO123552B (no) 1968-12-17 1968-12-17

Publications (1)

Publication Number Publication Date
US3661477A true US3661477A (en) 1972-05-09

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Family Applications (1)

Application Number Title Priority Date Filing Date
US885139A Expired - Lifetime US3661477A (en) 1968-12-17 1969-12-15 Heat shield

Country Status (4)

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US (1) US3661477A (no)
JP (1) JPS4825683B1 (no)
GB (1) GB1294675A (no)
NO (1) NO123552B (no)

Cited By (8)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US4688988A (en) * 1984-12-17 1987-08-25 United Technologies Corporation Coolable stator assembly for a gas turbine engine
US4987736A (en) * 1988-12-14 1991-01-29 General Electric Company Lightweight gas turbine engine frame with free-floating heat shield
US5655876A (en) * 1996-01-02 1997-08-12 General Electric Company Low leakage turbine nozzle
US6085515A (en) * 1996-06-11 2000-07-11 Siemens Aktiengesellschaft Heat shield configuration, particularly for structural parts of gas turbine plants
US20060010880A1 (en) * 2004-07-16 2006-01-19 Kim Kyung-Heui Gas turbine engine with seal assembly
US20060239841A1 (en) * 2005-04-21 2006-10-26 Panek Edward R Turbine heat shield with ribs
US20110014036A1 (en) * 2007-11-28 2011-01-20 Continental Automotive Gmbh Heat shield and turbocharger having a heat shield
US20110236184A1 (en) * 2008-12-03 2011-09-29 Francois Benkler Axial Compressor for a Gas Turbine Having Passive Radial Gap Control

Citations (7)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
GB542197A (en) * 1940-01-25 1941-12-30 British Thomson Houston Co Ltd Improved supporting arrangement for elastic fluid turbine diaphragms
US2296701A (en) * 1939-01-21 1942-09-22 Bristol Aeroplane Co Ltd Gas turbine
DE726510C (de) * 1940-01-25 1942-10-15 Aeg Trag- und Zentriereinrichtung fuer in der Teilfuge des Gehaeuses unterteilte Leitraeder von Dampf- oder Gasturbinen
US2440069A (en) * 1944-08-26 1948-04-20 Gen Electric High-temperature elastic fluid turbine
GB679916A (en) * 1949-04-29 1952-09-24 Geoffrey Bertram Robert Feilde Improvements in gas turbines
US3263424A (en) * 1965-03-25 1966-08-02 Birmann Rudolph Turbine-compressor unit
US3352105A (en) * 1964-12-14 1967-11-14 United Aircraft Canada Firewall attachment

Patent Citations (7)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US2296701A (en) * 1939-01-21 1942-09-22 Bristol Aeroplane Co Ltd Gas turbine
GB542197A (en) * 1940-01-25 1941-12-30 British Thomson Houston Co Ltd Improved supporting arrangement for elastic fluid turbine diaphragms
DE726510C (de) * 1940-01-25 1942-10-15 Aeg Trag- und Zentriereinrichtung fuer in der Teilfuge des Gehaeuses unterteilte Leitraeder von Dampf- oder Gasturbinen
US2440069A (en) * 1944-08-26 1948-04-20 Gen Electric High-temperature elastic fluid turbine
GB679916A (en) * 1949-04-29 1952-09-24 Geoffrey Bertram Robert Feilde Improvements in gas turbines
US3352105A (en) * 1964-12-14 1967-11-14 United Aircraft Canada Firewall attachment
US3263424A (en) * 1965-03-25 1966-08-02 Birmann Rudolph Turbine-compressor unit

Cited By (10)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US4688988A (en) * 1984-12-17 1987-08-25 United Technologies Corporation Coolable stator assembly for a gas turbine engine
US4987736A (en) * 1988-12-14 1991-01-29 General Electric Company Lightweight gas turbine engine frame with free-floating heat shield
US5655876A (en) * 1996-01-02 1997-08-12 General Electric Company Low leakage turbine nozzle
US6085515A (en) * 1996-06-11 2000-07-11 Siemens Aktiengesellschaft Heat shield configuration, particularly for structural parts of gas turbine plants
US20060010880A1 (en) * 2004-07-16 2006-01-19 Kim Kyung-Heui Gas turbine engine with seal assembly
US7174719B2 (en) * 2004-07-16 2007-02-13 Samsung Techwin Co., Ltd. Gas turbine engine with seal assembly
US20060239841A1 (en) * 2005-04-21 2006-10-26 Panek Edward R Turbine heat shield with ribs
US7631497B2 (en) * 2005-04-21 2009-12-15 Borgwarner Inc. Turbine heat shield with ribs
US20110014036A1 (en) * 2007-11-28 2011-01-20 Continental Automotive Gmbh Heat shield and turbocharger having a heat shield
US20110236184A1 (en) * 2008-12-03 2011-09-29 Francois Benkler Axial Compressor for a Gas Turbine Having Passive Radial Gap Control

Also Published As

Publication number Publication date
NO123552B (no) 1971-12-06
JPS4825683B1 (no) 1973-07-31
DE1960693A1 (de) 1970-07-02
DE1960693B2 (de) 1973-02-15
GB1294675A (en) 1972-11-01

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