EP0774050B1 - Structure de stator a rainures circonferentielles interrompues - Google Patents

Structure de stator a rainures circonferentielles interrompues Download PDF

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Publication number
EP0774050B1
EP0774050B1 EP95922886A EP95922886A EP0774050B1 EP 0774050 B1 EP0774050 B1 EP 0774050B1 EP 95922886 A EP95922886 A EP 95922886A EP 95922886 A EP95922886 A EP 95922886A EP 0774050 B1 EP0774050 B1 EP 0774050B1
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EP
European Patent Office
Prior art keywords
groove
dams
casing
machine
blades
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
EP95922886A
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German (de)
English (en)
Other versions
EP0774050A1 (fr
Inventor
Robert S. Mazzawy
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.)
RTX Corp
Original Assignee
United Technologies Corp
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Filing date
Publication date
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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
    • F01D11/00—Preventing or minimising internal leakage of working-fluid, e.g. between stages
    • F01D11/02—Preventing or minimising internal leakage of working-fluid, e.g. between stages by non-contact sealings, e.g. of labyrinth type
    • 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
    • F01D11/08—Preventing or minimising internal leakage of working-fluid, e.g. between stages for sealing space between rotor blade tips and stator
    • 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
    • F04D29/522—Casings; Connections of working fluid for axial pumps especially adapted for elastic fluid pumps
    • F04D29/526—Details of the casing section radially opposing blade tips
    • 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
    • F04D29/685—Inducing localised fluid recirculation in the stator-rotor interface

Definitions

  • This invention relates to axial flow rotary machines, and more particularly to a casing extending circumferentially about a compression stage for a gas turbine engine.
  • An axial flow, rotary machine such as a gas turbine engine for powering aircraft, has a compression section having a plurality of compression stages. As the working medium gases are drawn into the engine, the gases are compressed to raise the temperature and pressure of the gases. The gases are burned with fuel and expanded through a turbine to develop useful power and work. The work is transferred from the turbine section to the compression section via a rotor assembly. The rotor assembly in the compression section does work on the incoming gases to compress the gases.
  • a compression section is the fan section of a high bypass ratio turbofan engine.
  • the fan section may have a diameter of 150 to 200 mm (six to eight feet).
  • a plurality of fan blades extend radially outward from the rotor assembly across the flowpath for working medium gases.
  • the fan blades may be rotated at speeds in excess of 2,500 revolutions per minute about an axis of rotation. The rotating fan blades drive the gases rearwardly, compressing the working medium gases as the gases enter the fan section of the engine.
  • Each fan blade has a tip.
  • the tip is shrouded by the non-rotating adjacent casing; the fan blades do not have a shroud at the tip that rotates with the blades.
  • a clearance is provided between the tips and the adjacent casing to accommodate transient movement of the blade outward with respect to the casing. This region is commonly referred to as the end wall region. The transient movement may occur in response to maneuver loads and the normal growth of the blade in response to rotational forces acting on the blade.
  • the end wall region experiences different aerodynamic conditions than does the remainder of the compression stage.
  • the aerodynamic interaction between the tip of the rotor blade and the adjacent wall causes boundary layer effects and drag effects. These effects make it more difficult to force the flow rearwardly.
  • the pressure rises by a smaller amount in the endwall region than in the mid-span region of the fan blade.
  • Leakage also occurs at the end wall region between the opposite axially extending sides of the blade in response to pressure gradients across the tip of the blade. As a result, the surge margin and aerodynamic efficiency of the compression stage are adversely effected.
  • the recess is provided with circumferentially extending grooves (Figure 3).
  • the grooved construction provides an increase in surge margin over a smooth wall ( Figure 6, curve C ) with a decrease in aerodynamic efficiency.
  • the wall might be provided with a plurality of axially extending skewed chambers ( Figure 2) or a combination of both a circumferentially extending grooves and axially extending skewed chambers (Figure 4).
  • the Figure 4 construction shows the greatest increase in surge margin with the smallest decrease in aerodynamic efficiency over a smooth wall.
  • the axially skewed chambers decrease efficiency because the rotor blades use some of the work transferred to them to pump and re-pump the working medium gases in the axially skewed chamber.
  • Pumping increases the level of pressure attainable in the end wall region in response to a sudden back pressure.
  • pumping is the mechanism by which the axially extending chambers increase surge margin.
  • the circumferentially grooved construction has a moderate gain in surge margin and a moderate loss in the aerodynamic efficiency in comparison to the smooth wall.
  • surge margin if less surge margin were required, it could be accomplished with circumferentially extending grooves with less of a loss in aerodynamic efficiency than some of the other constructions illustrated in Roberts.
  • This invention is predicated upon the recognition that interrupting circumferential grooves with a critical number of dams related to the number of adjacent fan blades results in an improvement in surge margin with a negligible effect on aerodynamic efficiency.
  • a stator structure for the compression section of an axial flow rotary machine includes a casing which extends circumferentially about the array of rotor blades, the casing having an inner wall which includes a plurality of circumferentially extending grooves radially outwardly of and axially aligned with the blades, each of the grooves having a plurality of circumferentially spaced dams which interrupt the circumferential continuity of the groove and direct high velocity flow from the grooves into the adjacent boundary layer.
  • the number of circumferentiaily spaced dams lies in a range which is a function of the number of blades of the array of rotor blades; and, the circumferentially spaced grooves are disposed radially outwardly only of the mid-chord region of the rotor blade to increase the level of the effect of the grooves and dams on the boundary layer and decrease the impact on the aerodynamic efficiency.
  • a primary feature of the present invention is a casing radially outward of an array of rotor blades in a compression section.
  • the casing has a plurality of circumferentially extending grooves which are each circumferentially interrupted by a plurality of dams.
  • the dams are circumferentially spaced and extend in grooves having a depth which is twice the axial width of the groove.
  • the lands extending between adjacent grooves are half the width of the axial width of the grooves.
  • the number of dams lies in a range between three-fourths the number of rotor blades and one and one-half times the number of rotor blades.
  • the circumferential length of the dams is approximately equal to the width.
  • the number of dams in each groove is equal to the number of rotor blades and the dams are equally spaced about the circumference of the groove.
  • the grooves are disposed in a groove region which extends over the mid-chord region of the blade such that the groove region and the adjacent surface of the case outwardly of the blades are at the same radial height.
  • the adjacent surface of the case may be either recessed with respect to the inner surface of the casing or at the same radial height as the inner surface of the casing.
  • Figure 1 is a perspective view of a turbofan gas turbine engine showing a casing which circumscribes the tips of the fan blades.
  • Figure 2 is a developed view (looking outwardly) of a portion of the fan casing shown in Figure 1, illustrating the relationship of circumferentially grooves which are interrupted by circumferentially spaced dams.
  • Figure 3 is a developed view corresponding to the view shown in Figure 2.
  • Figure 4 is an alternate embodiment of the construction shown in Figure 2 in which the grooves are staggered circumferentially.
  • Figure 5 is a side-elevation view taken along the lines 3-3 of Figure 2.
  • Figure 6 is a graphical representation of the change in percentage points of surge margin and aerodynamic efficiency for the construction shown in Figure 3 to the number of circumferentially spaced dams in the circumferential grooves expressed as a function of the number N of blades in the array of rotor blades.
  • Figure 1 is a perspective view of a turbofan gas turbine engine 10 of the type used to power commercial airliners.
  • the engine includes a core section 12 and a fan section 14.
  • a plurality of unshrouded fan blades 16 extend radially outward from a rotor 18.
  • the rotor is disposed about an axis of rotation A .
  • Each fan blade has an airfoil 22 having a tip 24.
  • the tip does not have a rotating shroud, but may have a part span shroud.
  • the fan blade has a platform 26 which adapts the fan blade to engage an attachment slot in the rotor assembly.
  • the length of the fan blade may exceed 0.61 m (two feet) and in the particular embodiment shown, is approximately 0.69 m (twenty-seven (27) inches long) as measured from the platform.
  • the outer diameter of the tip of the fan blade is about 2.39 m (ninety-four (94) inches) as measured from the axis of rotation A .
  • the fan (compression) section 14 includes a primary flow path 28 for working medium gases which extends through the core section.
  • a secondary flow path 32 for working medium gases extends through the fan section outwardly of the primary flowpath.
  • the secondary or bypass flow path is annular in shape.
  • a fan casing 34 extends circumferentially about the flow path to bound the flow path at its outermost portion.
  • the casing has an inner surface 36, formed of an abradable material which extends circumferentially about the rotor blades. The inner surface is spaced radially from the rotor blades leaving a clearance gap G therebetween.
  • the clearance gap in the present embodiment is approximately 2.29 mm (ninety-thousandths (0.090) of an inch).
  • Figure 2 is an undeveloped view and Figure 3 is a developed view of the casing 34 looking outwardly toward the inner surface 36 of the casing
  • a plurality of grooves 38 extends circumferentially about the interior of the casing.
  • a plurality of dams 42 are disposed in each groove. Each dam extends axially across the groove and radially in the groove to circumferentially interrupt the groove.
  • Each dam is spaced from the adjacent dam by a pitch distance P .
  • the pitch distance is the distance from a point on one dam to the corresponding point on an adjacent dam.
  • the pitch distance P is equal to the circumference of the inner surface divided by the number of dams.
  • the number of dams is equal, for example, to the number of blades N of the array of rotor blades.
  • Two rotor blade tips 24, each associated with an adjacent fan blade 16, are shown in Figure 2 in full and in Figure 3 in phantom.
  • the rotor blades sweep upwardly (Figure 2) and from right to left ( Figure 3) under operative conditions as shown by the arrow V which indicates the direction of motion of the fan blades.
  • Each fan blade has a leading edge 44 and a trailing edge 46.
  • a suction sidewall 48 on one side of the blade extends from the leading edge to the trailing edge.
  • a pressure sidewall 52 on the other side of the blade extends from the leading edge to the trailing edge.
  • the fan blade has an axial length L as measured parallel to the axis of rotation.
  • the fan blade has a leading edge region 54 which extends for the first twenty percent (20%) of the axial length L of the blade and a trailing edge region 56 which extends for the last twenty percent (20%) of the axial length L of the fan blade.
  • the fan blade has a mid-chord region 58 which extends between the leading edge region and the trailing edge region.
  • the casing has a groove region 62 which extends circumferentially at a location which is radially outwardly of the mid-chord region of the blade and contains the grooves.
  • Figure 4 is a view corresponding to the view shown in Figure 3 of an alternate embodiment of the casing shown in Figure 3.
  • the circumferential grooves 38 are interrupted by dams 42 which are staggered circumferentially with respect to the adjacent dam.
  • FIG 5 is an enlarged, simplified cross-sectional view of the outer casing 36 and fan blade tip 24 shown in Figure 1, Figure 2, and Figure 3.
  • the inner surface 36 of the casing outwardly of the fan blade tip is at the same radial location as the adjacent structure.
  • the fan blade 16 may rub against the inner (abradable seal) surface 36, cutting a recess 64 which would result in a relationship to the inner surface as shown by the phantom lines.
  • the surface 36 outwardly of the fan blade tip is at a different radial location than the inner surface of the adjacent structure.
  • the recess may be deliberately formed in the outer case as shown, for example, in U.S. Patent No.: 4,239,452 discussed above.
  • Figure 5 shows the plurality of grooves 38.
  • each groove 38 has an axial width W , and a radial depth D as measured from the inner surface of the casing outwardly of the rotor blade.
  • Each groove is spaced axially from the adjacent groove by a distance equal to one half of the axial width W , leaving a land 66 therebetween.
  • a plurality of dams 42 at each groove extend axially from the land to circumferentially interrupt the full depth of the groove.
  • the circumferential length L C of the dam is equal to the axial width W of the groove 38.
  • Figure 6 shows the relationship of the change in aerodynamic efficiency (curve E ) and change in surge margin (curve S ) for the interrupted groove 38 construction as a function of the number of dams (interruptions) 42 and pitch P .
  • the number of dams is expressed in terms of the number of fan blades N in the array of fan blades.
  • the changes are measured in comparison to a base line construction which does not have dams that interrupt the circumferential continuity of each groove.
  • the relationships shown in Figure 6 are based on an analysis of empirical data.
  • dams 42 increases the surge margin with little or no decrease in the aerodynamic efficiency.
  • the surge margin increases by approximately four points with no apparent decrease in aerodynamic efficiency.
  • the quantity of dams increases, there is a slight decrease in aerodynamic efficiency until the quantity of dams is about equal to one and one-half times the number of rotor blades. Beyond that point, the surge margin does not measurably increase and aerodynamic efficiency continues to decrease. Accordingly, the optimum design range for the quantity of dams is believed to be equal to three-fourths to one and one-half times the number of rotor blades.
  • the pitch P (spacing) between the dams for this range increases from about two-thirds of the circumference of the inner wall divided by the number N of rotor blades to one and one-third times the circumference of the inner wall divided by the number N of rotor blades.
  • the low velocity gases in the boundary layer can cause significant losses because the boundary layer may separate from the wall 36.
  • High velocity gases in the circumferentially extending grooves 38 are driven inwardly by the dams into the low energy boundary layer. These high-energy gases act to energize the boundary layer and avoid separation of the boundary layer. This avoids concomitantly the losses associated with separation of the boundary layer and results in an aerodynamic efficiency which remains constant even as the circumferential grooves provide an improvement in surge margin. Accordingly, adding the circumferentially spaced dams to the circumferential grooves allows the surge margin to increase without an unacceptable decrease in aerodynamic efficiency of the array provided the quantity of dams equals a number of fan blades within the optimum design range.
  • the high velocities in the grooves are made possible by the pitch P of the dams which provides sufficient length of uninterrupted groove for the velocities to develop. In embodiments having too many dams, the high velocities never develop. In embodiments having too few dams, the high velocities develop but are not injected with desirable frequency into the boundary layer.

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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. Machine rotative à écoulement axial, comportant un axe de rotation (A), une partie compression (14), qui comprend un chemin d'écoulement annulaire (28) pour des gaz constituant le fluide moteur, un ensemble formant rotor, qui comprend un groupement de pales de rotor (16), qui s'étendent radialement vers l'extérieur en travers du chemin d'écoulement, chaque pale de rotor présentant un bout de pale (24), et une structure de stator, qui comprend : un carter (34) ayant une surface interne (36), le carter s'étendant circonférentiellement autour du chemin d'écoulement (28) pour les gaz constituant le fluide moteur, pour limiter le chemin d'écoulement, et s'étendant circonférentiellement à l'extérieur des pales de rotor (16) et étant espacé radialement des pales de rotor en laissant un écartement (G) entre eux, le carter (34) ayant une zone de rainures, qui est située radialement à l'extérieur de chaque pale, la zone de rainures comportant une pluralité de rainures (38), chaque rainure s'étendant circonférentiellement sur l'intérieur du carter, et chaque rainure étant espacée axialement de la rainure adjacente, en laissant un plat (66) entre elles, au moins l'une desdites rainures ayant une largeur axiale (W), une profondeur radiale (D), mesurées depuis la surface interne,
       caractérisée en ce que ladite au moins une rainure a un rapport de la profondeur radiale (D) à la largeur axiale (W) qui est supérieur ou égal à un (D/W ≥ 1), et en ce qu'une pluralité de barrages (42) sont disposés dans ladite rainure, le nombre de barrages étant au moins égal en nombre aux trois quarts du nombre de pales (N) (0,75 N) et s'étendant axialement à partir du plat (66) pour interrompre circonférentiellement la rainure, et en ce que ladite surface interne (36) dudit carter (34) est faite en un matériau susceptible de s'user par frottement.
  2. Machine selon la revendication 1, dans laquelle le nombre de barrages (42) est égal à un nombre qui est compris dans une fourchette allant des trois quarts du nombre de pales (N) (0,75 N) à une fois et demie le nombre de pales (N) (1,5 N).
  3. Machine selon la revendication 1 ou 2, dans laquelle chaque barrage (42) est espacé circonférentiellement du barrage adjacent de la distance d'un pas (P) égal à la circonférence de la surface interne divisée par (N).
  4. Machine selon l'une quelconque des revendications précédentes, dans laquelle le nombre de barrages (42) est égal au nombre de pales (N).
  5. Machine selon l'une quelconque des revendications précédentes, dans laquelle ladite rainure (38) a un rapport de la profondeur radiale (D) à la largeur axiale (W) qui est égal à deux.
  6. Machine selon la revendication 5, dans laquelle la largeur axiale du plat (66) est égale à la moitié de la largeur axiale (W) de la rainure (38).
  7. Machine selon l'une quelconque des revendications précédentes, dans laquelle au moins cinq des rainures ont une pluralité de barrages (42).
  8. Machine selon l'une quelconque des revendications précédentes, dans laquelle au moins deux des rainures (38) ont une pluralité de barrages (42), les barrages d'une rainure étant décalés circonférentiellement par rapport à l'autre rainure.
  9. Machine selon l'une quelconque des revendications précédentes, dans laquelle ledit carter (34) présente une cavité (64) dans le carter, qui fait face aux pales de rotor, et dans laquelle la zone de rainures est disposée dans ledit carter.
  10. Machine selon l'une quelconque des revendications précédentes, dans laquelle le bout (24) de la pale de rotor (16) a une zone mi-corde qui s'étend sur soixante pour cent (60 %) de la longueur (L) du bout dans le sens axial, et qui est espacée d'une distance égale à vingt pour cent (20 %) de la longueur (L) par rapport au bord d'attaque (44) et par rapport au bord de fuite (46) de la pale, et dans laquelle la zone de rainures est située radialement à l'extérieur de la zone de mi-corde de chaque pale.
EP95922886A 1994-06-14 1995-05-24 Structure de stator a rainures circonferentielles interrompues Expired - Lifetime EP0774050B1 (fr)

Applications Claiming Priority (3)

Application Number Priority Date Filing Date Title
US25979194A 1994-06-14 1994-06-14
US259791 1994-06-14
PCT/US1995/006582 WO1995034745A1 (fr) 1994-06-14 1995-05-24 Structure de stator a rainures circonferentielles interrompues

Publications (2)

Publication Number Publication Date
EP0774050A1 EP0774050A1 (fr) 1997-05-21
EP0774050B1 true EP0774050B1 (fr) 1999-03-10

Family

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Family Applications (1)

Application Number Title Priority Date Filing Date
EP95922886A Expired - Lifetime EP0774050B1 (fr) 1994-06-14 1995-05-24 Structure de stator a rainures circonferentielles interrompues

Country Status (5)

Country Link
EP (1) EP0774050B1 (fr)
JP (1) JP3640396B2 (fr)
DE (1) DE69508256T2 (fr)
TW (1) TW268070B (fr)
WO (1) WO1995034745A1 (fr)

Cited By (2)

* Cited by examiner, † Cited by third party
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CN106968986A (zh) * 2017-05-08 2017-07-21 中国航发湖南动力机械研究所 缝式处理机匣及压气机
US10107307B2 (en) 2015-04-14 2018-10-23 Pratt & Whitney Canada Corp. Gas turbine engine rotor casing treatment

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GB2385378B (en) 2002-02-14 2005-08-31 Rolls Royce Plc Engine casing
GB2419638A (en) * 2004-10-26 2006-05-03 Rolls Royce Plc Compressor casing with an abradable lining and surge control grooves
US7861823B2 (en) * 2005-11-04 2011-01-04 United Technologies Corporation Duct for reducing shock related noise
GB2435904B (en) 2006-03-10 2008-08-27 Rolls Royce Plc Compressor Casing
DE102008009604A1 (de) 2008-02-15 2009-08-20 Rolls-Royce Deutschland Ltd & Co Kg Gehäusestrukturierung zum Stabilisieren der Strömung in einer Strömungsarbeitsmaschine
US8118548B2 (en) * 2008-09-15 2012-02-21 General Electric Company Shroud for a turbomachine
CN101483624B (zh) * 2009-02-10 2011-06-08 东南大学 Msk差分检测解调电路中频率漂移补偿装置及补偿方法
US8177494B2 (en) * 2009-03-15 2012-05-15 United Technologies Corporation Buried casing treatment strip for a gas turbine engine
GB2477745A (en) * 2010-02-11 2011-08-17 Rolls Royce Plc Compressor Casing
US8740563B2 (en) * 2010-11-22 2014-06-03 General Electric Company Sealing assembly for use in turbomachines and methods of assembling same
EP2458157B1 (fr) * 2010-11-30 2015-10-14 Techspace Aero S.A. Abradable de virole intérieure de stator
JP5579104B2 (ja) * 2011-02-28 2014-08-27 三菱重工業株式会社 回転機械の抽気構造
DE102011006273A1 (de) 2011-03-28 2012-10-04 Rolls-Royce Deutschland Ltd & Co Kg Rotor einer Axialverdichterstufe einer Turbomaschine
DE102011006275A1 (de) 2011-03-28 2012-10-04 Rolls-Royce Deutschland Ltd & Co Kg Stator einer Axialverdichterstufe einer Turbomaschine
DE102011007767A1 (de) * 2011-04-20 2012-10-25 Rolls-Royce Deutschland Ltd & Co Kg Strömungsmaschine
CN103062131B (zh) * 2011-10-20 2015-08-12 中国科学院工程热物理研究所 柔性非轴对称机匣处理流动控制方法
FR2988146B1 (fr) * 2012-03-15 2014-04-11 Snecma Carter pour roue a aubes de turbomachine ameliore et turbomachine equipee dudit carter
GB201205663D0 (en) * 2012-03-30 2012-05-16 Rolls Royce Plc Effusion cooled shroud segment with an abradable system
US10465716B2 (en) 2014-08-08 2019-11-05 Pratt & Whitney Canada Corp. Compressor casing
CA2955646A1 (fr) 2016-01-19 2017-07-19 Pratt & Whitney Canada Corp. Boitier d'aube de rotor de turbine a gaz
CN109653808B (zh) * 2018-11-30 2020-08-28 西安交通大学 一种具有内齿槽的径向轮缘密封结构
US20230151825A1 (en) * 2021-11-17 2023-05-18 Pratt & Whitney Canada Corp. Compressor shroud with swept grooves

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US4239452A (en) * 1978-06-26 1980-12-16 United Technologies Corporation Blade tip shroud for a compression stage of a gas turbine engine
JPS6318799Y2 (fr) * 1980-12-02 1988-05-26
GB2146707B (en) * 1983-09-14 1987-08-05 Rolls Royce Turbine

Cited By (2)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US10107307B2 (en) 2015-04-14 2018-10-23 Pratt & Whitney Canada Corp. Gas turbine engine rotor casing treatment
CN106968986A (zh) * 2017-05-08 2017-07-21 中国航发湖南动力机械研究所 缝式处理机匣及压气机

Also Published As

Publication number Publication date
DE69508256D1 (de) 1999-04-15
JPH10501318A (ja) 1998-02-03
TW268070B (fr) 1996-01-11
WO1995034745A1 (fr) 1995-12-21
EP0774050A1 (fr) 1997-05-21
JP3640396B2 (ja) 2005-04-20
DE69508256T2 (de) 1999-10-14

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