US3918832A - Stator construction for an axial flow compressor - Google Patents

Stator construction for an axial flow compressor Download PDF

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Publication number
US3918832A
US3918832A US492574A US49257474A US3918832A US 3918832 A US3918832 A US 3918832A US 492574 A US492574 A US 492574A US 49257474 A US49257474 A US 49257474A US 3918832 A US3918832 A US 3918832A
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US
United States
Prior art keywords
vane
case
compressor
vanes
retainer
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
US492574A
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English (en)
Inventor
Richmond G Shuttleworth
Donald L Williams
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
Priority date (The priority date 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 date listed.)
Filing date
Publication date
Application filed by United Technologies Corp filed Critical United Technologies Corp
Priority to US492574A priority Critical patent/US3918832A/en
Priority to CA230,590A priority patent/CA1027870A/en
Priority to CH927875A priority patent/CH589229A5/xx
Priority to DE2532554A priority patent/DE2532554C2/de
Priority to SE7508371A priority patent/SE419260B/xx
Priority to IT25778/75A priority patent/IT1040063B/it
Priority to GB31284/75A priority patent/GB1511019A/en
Priority to FR7523244A priority patent/FR2280812A1/fr
Priority to JP50091848A priority patent/JPS6050999B2/ja
Priority to NO75752644A priority patent/NO149010C/no
Priority to BR7504846*A priority patent/BR7504846A/pt
Application granted granted Critical
Publication of US3918832A publication Critical patent/US3918832A/en
Assigned to FIRST NATIONAL BANK OF CHICAGO, THE reassignment FIRST NATIONAL BANK OF CHICAGO, THE LICENSE (SEE DOCUMENT FOR DETAILS). Assignors: ELLIOT TURBOMACHINERY CO., INC.
Anticipated expiration legal-status Critical
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
    • F01D9/00—Stators
    • F01D9/02—Nozzles; Nozzle boxes; Stator blades; Guide conduits, e.g. individual nozzles
    • F01D9/04—Nozzles; Nozzle boxes; Stator blades; Guide conduits, e.g. individual nozzles forming ring or sector
    • F01D9/042—Nozzles; Nozzle boxes; Stator blades; Guide conduits, e.g. individual nozzles forming ring or sector fixing blades to stators

Definitions

  • Each vane assembly includes one or more stator vanes extending radially inward from the case inner wall into the axial flow path of the working medium, and an arcuate retainer which slidably engages the base portion of each vane. During operation of the engine the vanes are subject to vibration which is dissipatively transferred to both the working medium and the compressor case.
  • Gas turbine engines are designed and constructed to provide adequate structural integrity of the individual components while maintaining acceptable levels of aer odynamic performance.
  • a stator encases the rotor assembly.
  • the tips of the compressor blades which extend radially outward from the rotor into the flow path of the working medium, coopcrate with a mating surface on the compressor stator which surrounds the tips of the blades to form a gas seal between therotor and stator.
  • the tips of the compressor vanes which extend radially inward from the compressor case into the flow path of the working medium, cooperate with a corresponding surface on the rotor to form a gas seal between the rotor and the stator.
  • the aerodynamic performance of the compressor is highly dependent on the clearances between the rotor and the stator at the .blade and vane tips. Even a slight reductionin tip clearance can improve the aerodynamic performance of the compressor significantly.
  • Additional clearance between the cooperating sealing surfaces of the rotor and stator is also provided to allow for distortion of the compressor case under transient thermal conditions.
  • a compressor case havinga nonuniform mass distributed about its circumference, such as an axially split case experiences distorted thermal growth due to the concentration of mass in the flange areas..As the case is exposed to changing thermal environments the areas of the case having low mass concentration expand more rapidly than those areas of the case having high mass concentration due to the time differential required for these areas-to reach steady state thermal conditions.
  • Most axial flow gas turbines have a double compressor case comprising-an outer case which provides structural support to the engine bearings and an inner case which radially defines the flow path of the working medium and which supports the compressor vanes extending from the inner wall into the flow path.
  • the inner case comprises a series of annular rings bolted together. Alternate rings support rows of stator vanes and are joined by an intermediate ring which is the sealing surface opposing the corresponding row of rotor blade tips.
  • the joining features of each ring are formed with respect to the axis of each individual ring within limits of a concentricity tolerance. As adjacent rings are joined together concentricity tolerances can accumulate so that, potentially, the misalignmentbetween the first and the last stages of the stator assembly can be excessive. Consequently design clearance between the rotor and the stator assemblies is increased to accommodate for this potential misalignment.
  • gas turbines particularly those used in industrial applications, have a single compressor case which is split axially and joined by longitudinal flanges on opposing halves of the compressor case.
  • the mass distribution about the circumference of the compressor case is nonuniform and causes eccentric distortion of the compressor case during transient thermal conditions.
  • a single compressor case with uniform mass distribution has the potential for high aerodynamic performance. Although adequate mechanical support of the vanes from such a case is difficult. The vanes are necessarily cantilevered from the case and, therefore, tend to vibrate during operation of the engine. Vanes which are rigidly mounted to the compressor case have a limited lifetime since they tend to crack or fracture due to vibratory stresses in the area where the airfoil is joined to a supporting structure.
  • a single-wall case requires vane loading slots into which the vanes are inserted and repositioned circumferentially about the compressor case. The loading slot region frequently experiences excessive stress concentration in the areas where small radii interrupt uniform stress patterns.
  • Warnkin discloses a plurality of stator vanes which are assembled on an arcuate segment and, subsequently, mounted in a circumferential track of an axially split compressor case.
  • Each compressor vane has a wedge shaped root which is in; serted through an aperture in the arcuate segment where it is held firmly in position.
  • a primary object of the present invention is to improve the structural integrity of the compressor case and compressor vanes of a gas turbine engine.
  • a further object of the present invention is to improve the aerodynamic performance of the compressor.
  • a plurality of compressor vanes having a retaining slot in the base portion of each vane is slidably mounted on an arcuate retainer having a correspondingly shaped cross section, the vanes extending from the arcuate retainer in a radially inward direction toward the center of curvature of the arc;
  • a plurality of arcuate retainers and the vanes mounted thereupon comprise a set of vanes for a single compressor stage and are bolted into a circumferential retaining track of uniform cross section which has been machined into the inner wall of the compressor case, one or more bolts penetrating the case from the outer wall engage each arcuate retainer.
  • a primary feature of the present invention is the dissipative transfer of vibrational energy from the vanes to the working medium and to the compressor case.
  • An additional feature is the one piece compressor case having circumferential tracks of uniform cross section for retaining detachable vane assemblies.
  • Principal advantages of the present invention are the ability of the stator assembly to dissipate vibrational energy without cracking the stator vanes and the ability of the compressor case to maintain uniform blade and vane tip clearances around the circumference of the rotor.
  • An additional advantage of the present invention is the confinement of high stress concentrations to the vanes and to the arcuate retainers which are easily replaceable and low cost parts.
  • FIG. 1 is a simplified elevation view showing an axial flow gas turbine engine
  • FIG. 2 is a simplified cross-sectional view of a portion of the compressor of the gas turbine engine shown in FIG. 1;
  • FIG. 3 is a section view of the compressor taken in the direction 3 as shown in FIG. 2;
  • FIG. 4 is a section view taken along the line 4-4 as shown in FIG. 3;
  • FIG. 5 is a section view of the compressor vane attachment under gas pressure loads.
  • FIG. 6 is a section view of the compressor vane attachment under a condition of vibrational excitation in which the vibratory loads are opposite to and exceed the gas pressure loads.
  • the gas turbine 10 shown in FIG. 1 is an axial flow engine having a multistage compressor 12 joined to a multistage turbine 14 by a combustor 16. Air is compressed in the compressor, is mixed with fuel and burned in the combustion section to produce hot gases which are expanded through a series of nozzles within the turbine section. The more air that an engine can 4 compress and use the greater is the power or thrust that can be produced within the engine.
  • a rotor assembly 18 comprises a plurality of compressor wheels 20 which are separated axially by spacers 22.
  • Each compressor wheel includes a disk 24 and a plurality of blades attached thereto as represented by the single blade 26 on each wheel.
  • Each blade has a platform 28 at the base of an airfoil section 30.
  • An axial gap between the blade platforms of adjacent wheels is spanned by an inner air seal 32.
  • the rotor assembly is radially enclosed by a compressor stator 34 comprising a plurality of vane stages 36 each mounted within a circumferential track 38 in a compressor case 40.
  • each vane stage comprises a plurality of vane assemblies 42 which include one or more vanes 44, an arcuate retainer 46 and a pair of end plates 48.
  • Each vane assembly is held within the circumferential track by one or more bolts 50 which penetrate the case to engage the arcuate retainer.
  • Each compressor vane has an airfoil section 52, a base 54 including retaining slot 56 and a tip 58 as shown in FIG. 4.
  • An axial clearance 60 is provided between the vane base and the compressor case and a radial clearance 62 is provided between the vane base and the arcuate retainer.
  • each vane has a case bearing surface 64 which is opposed by a vane bearing surface 66 of the case and a retainer bearing surface 68 which is opposed by a vane bearing surface 70 of the retainer.
  • each arcuate retainer 46 During assembly of the compressor one or more vanes 44 are slidably mounted on each arcuate retainer 46, the tee shaped arcuate retainer of the preferred embodiment engaging the correspondingly shaped slot in the base of each vane.
  • An end plate 48 is affixed to each end of the arcuate retainer to trap the vanes on the retainer.
  • a plurality of vane assemblies is bolted into each circumferential track to form each compressor vane stage 36, each assembled vane extending radia'lly inward across the flow path of the working medium.
  • the end plates perform the additional function of preventing circumferential movement of the vanes about the track during operation of the engine.
  • the number of vanes mounted within each vane assembly is varied according to the size and weight of the individual components. Including a large number of vanes in each vane assembly lessens the number of steps required to assemble a complete vane stage in the compressor case. Including a smaller number of vanes in each vane assembly reduces the weight of the assembly and makes it more easily mountable within the compressor case.
  • a single vane stage comprises five vane assemblies which weigh approximately thirty pounds each and include fourteen vanes. As many as ten vane assemblies are commonly used.
  • the number of vanes on the vane assembly to be last assembled is limited by the cord length of its arcuate retainer.
  • the cord length of the last arcuate retainer must be smaller than the distance between the tips 58 of the vanes through which the vane assembly passes as it is positioned into the circumferential track from a radially inward direction.
  • the fifth vane assembly is split and includes one vane assembly having thirteen vanes and one assembly having a single vane as shown in FIG. 3.
  • the vane assemblies are bolted into the circumferential track by vane assembly comprising a'single vane is mounted within a circumferential track in the same manner as vane assemblies having a plurality of vanes.
  • One of the significant aspects of a compressor constructed in accordance with the present invention is the damping of vibratory energy in the compressor vanes during engine operation.
  • the axial clearances 60 between the base of each vane and the compressor case, and a radial clearance 62 between the base of each vane and the corresponding arcuate retainer permit limited movement of the vanes after the arcuate retainer of each vane assembly is secured to the compressor case.
  • both the radial and axial clearances are one thousandth to thirteen thousandths of an inch.
  • vanes tilt toward the front or low pressure end of the compressor until the retainer bearing surface of the vane 68 comes to rest against the vane bearing surface of the retainer 70 and simultaneously the case bearing surface of the vane 64 comes to rest against the vane bearing surface of the case 66.
  • inherent vibrational loads cyclically exceed the static pressure loads on the vane causing the vane to tilt from its forward position to a rearward position as shown in FIG. 6; Rearward movement of the vane is op-- posed by the pressure loading forces which cushion the airfoil surface to dissipate vibrational energy from the vane.
  • friction damping occurs between the side bearing surfaces of adjacent vanes and between the bearing surfaces on each vane which are in contact with the retainer or the case.
  • vibrational energy is most commonly removed from the compressor vanes of gas turbine engines through a rigidly fixed attachment joining the vane and to the case.
  • the life of the vane is shortened because vibratory stresses concentrate at the juncture between the vane airfoil and the vane base. The cumulative vibratory stresses at this juncture ultimately crack or fracture the vanes.
  • Vanes attached in accordance with the present invention are not rigidly affixed to the case and do not experience excessive vibratory stresses.
  • the compressor case has essentially u-shaped tracks machined into the circumferential inner wall.
  • the absence of vane loading slots allows case stresses to uniformly distribute around the circumference of the case thereby maximizing the case life. Stress concentrations within the stator do occur at the internal structural corners of the retainer and the vane base. However, the retainers and the vanes are easily replaced at minimal cost when structural cracks appear.
  • Movement of the vanes relative to the case and the arcuate retainer can cause wear along the surfaces of contact.
  • the bearing surfaces are treated with a hardfacing material.
  • the construction shown in the preferred embodiment has a simple geometry which permits the application of hardfacing material to the bearing surfaces.
  • an increase in the blade and vane tip clearance of ten thousandths throughout the length of the compressor decreases the compressor efficiency by one percent.
  • a common design goal is the 6 control of the tip clearances to within one percent of the 'spanwidth of the airfoil sections which is for one typical embodiment a thirty eight thousandths nominal clearance between each vane or blade tip and its corresponding sealing surface at a diameter of approximately fifty inches.
  • Control of the tip clearance requires control of the distortion of the compressor case and control of the concentricity of mating surfaces with respect to a common axis.
  • Control of the compressor case distortion is achieved by use of a nonsplit compressor having a uniform cross-sectional area about its circumference at any axial position.
  • the use of a nonsplit compressor case allows a uniform cross section by eliminating the mass concentration of the flanges joining a split compressor case.
  • the areas of a split case which have high mass, such as the flange areas exhibit a retarded rate of thermal response.
  • a nonuniform thermal response distorts the sealing surfaces of the case at the blade tips from a circular configuration, and alters the radially inward position of effected compressor vanes.
  • the tip clearances must be adjusted to compensate for the range of thermal response rather than a single thermal response as in the preferred embodiment.
  • a second principal problem in holding minimum tip clearances is the buildup of concentricity tolerances between opposing compressor parts.
  • Most gas turbines in use today utilize stator constructions of the double case type wherein the inner case supports the compressor vanes and the outer wall provides structural support to the engine bearings.
  • the inner case supporting the vanes comprises a plurality of cylindrically shaped vane supports placed in axially adjacent positions and bolted together.
  • Each cylindrical support is machined to concentricity tolerances in relation to its own axis. As the support cylinders are bolted together the potential concentricity misalignment increases from the first support to the last support.
  • each circumferential track is machined in relation to the same axis which is the axis of the compressor case.
  • the single case construction has a reduced number of pieces and, therefore, has reduced assembly complexity.
  • the single case construction is lighter than the double case construction and can be mass balanced for improved blade tip clearance.
  • the mass of the compressor case at any axial position can be increased to match the predicted thermal response of the rotor at that axial position.
  • the thermal response of a compressor having a double case construction may be similarly controlled, the geometry is more complex and accurate distortion prediction is more difficult.
  • a compressor stator including a plurality of vane assemblies detachably mounted within a circumferential track in the inner wall of the case of the compressor wherein each vane assembly comprises:
  • At least one vane having a base including a retaining slot which is slidably engaged by the arcuate retainer, the vane extending radially inward from the retainer toward the center of curvature of the retainer;
  • a first end plate attached to one end of the arcuate retainer for preventing the circumferential rotation of the vanes around the track in the inner wall.
  • the invention according to claim 1 further including a second end plate attached to the end of the arcuate retainer opposite said first end plate for trapping the vanes on the arcuate retainer.
  • a stator construction for an axial flow compressor having means for damping vane vibration including a compressor case having, in the case inner wall, a circumferential track containing a plurality of vane assemblies each of which comprises an arcuate retainer slidably engaged to a slot in the base of the stator vanes which are attached to and extend radially inward from the arcuate retainer toward the center of curvature of the retainer, and a first end plate attached to one end of the retainer to prevent circumferential movement of the vanes about the compressor case, each of said vanes having an axial and radial clearance between the vane base and the compressor case which permits movement of the vane under the influence of vibrational excitation to dissipatively transfer vibrational energy from the blade to the working medium by pressure load forces on the airfoil and to the compressor case by friction forces between the vanes, case and retainer.
  • At least one assembly includes no more than one vane 9.
  • the invention according to claim 8 further including a second end plate attached to the end of the arcuate retainer opposite said first end plate for trapping the vanes on the arcuate retainer.

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  • Engineering & Computer Science (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Structures Of Non-Positive Displacement Pumps (AREA)
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  • Polysaccharides And Polysaccharide Derivatives (AREA)
  • Turbine Rotor Nozzle Sealing (AREA)
US492574A 1974-07-29 1974-07-29 Stator construction for an axial flow compressor Expired - Lifetime US3918832A (en)

Priority Applications (11)

Application Number Priority Date Filing Date Title
US492574A US3918832A (en) 1974-07-29 1974-07-29 Stator construction for an axial flow compressor
CA230,590A CA1027870A (en) 1974-07-29 1975-07-02 Stator construction for an axial flow compressor
CH927875A CH589229A5 (enExample) 1974-07-29 1975-07-16
DE2532554A DE2532554C2 (de) 1974-07-29 1975-07-21 Verdichterstator
SE7508371A SE419260B (sv) 1974-07-29 1975-07-23 Statoranordning vid axialflodeskompressor
GB31284/75A GB1511019A (en) 1974-07-29 1975-07-25 Stator construction for an axial flow compressor
IT25778/75A IT1040063B (it) 1974-07-29 1975-07-25 Starote per un compressore a flusso assiale
FR7523244A FR2280812A1 (fr) 1974-07-29 1975-07-25 Stator pour compresseur a courant axial
JP50091848A JPS6050999B2 (ja) 1974-07-29 1975-07-28 軸流圧縮機のステ−タ構造
NO75752644A NO149010C (no) 1974-07-29 1975-07-28 Kompressor.
BR7504846*A BR7504846A (pt) 1974-07-29 1975-07-29 Estator para compressor

Applications Claiming Priority (1)

Application Number Priority Date Filing Date Title
US492574A US3918832A (en) 1974-07-29 1974-07-29 Stator construction for an axial flow compressor

Publications (1)

Publication Number Publication Date
US3918832A true US3918832A (en) 1975-11-11

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US492574A Expired - Lifetime US3918832A (en) 1974-07-29 1974-07-29 Stator construction for an axial flow compressor

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US (1) US3918832A (enExample)
JP (1) JPS6050999B2 (enExample)
BR (1) BR7504846A (enExample)
CA (1) CA1027870A (enExample)
CH (1) CH589229A5 (enExample)
DE (1) DE2532554C2 (enExample)
FR (1) FR2280812A1 (enExample)
GB (1) GB1511019A (enExample)
IT (1) IT1040063B (enExample)
NO (1) NO149010C (enExample)
SE (1) SE419260B (enExample)

Cited By (11)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US5167488A (en) * 1991-07-03 1992-12-01 General Electric Company Clearance control assembly having a thermally-controlled one-piece cylindrical housing for radially positioning shroud segments
EP0763653A1 (en) * 1995-09-13 1997-03-19 SOCIETE DE CONSTRUCTION DES AVIONS HUREL-DUBOIS (société anonyme) Thrust reverser door with jet deflection cascade
US6234750B1 (en) 1999-03-12 2001-05-22 General Electric Company Interlocked compressor stator
US20070079506A1 (en) * 2005-10-06 2007-04-12 General Electric Company Method of providing non-uniform stator vane spacing in a compressor
US20180066673A1 (en) * 2016-09-02 2018-03-08 United Technologies Corporation Repeating airfoil tip strong pressure profile
US9920652B2 (en) 2015-02-09 2018-03-20 United Technologies Corporation Gas turbine engine having section with thermally isolated area
US20180230856A1 (en) * 2016-10-19 2018-08-16 United Technologies Corporation Engine cases and associated flange
US10280779B2 (en) 2013-09-10 2019-05-07 United Technologies Corporation Plug seal for gas turbine engine
US10287905B2 (en) 2013-11-11 2019-05-14 United Technologies Corporation Segmented seal for gas turbine engine
US10634055B2 (en) 2015-02-05 2020-04-28 United Technologies Corporation Gas turbine engine having section with thermally isolated area
US20220381150A1 (en) * 2021-05-26 2022-12-01 General Electric Company Split-line stator vane assembly

Families Citing this family (1)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
RU2232303C1 (ru) * 2002-11-10 2004-07-10 Открытое акционерное общество "Научно-производственное объединение "Сатурн" Силовой корпус компрессора гтд

Citations (4)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US2997275A (en) * 1959-03-23 1961-08-22 Westinghouse Electric Corp Stator structure for axial-flow fluid machine
US3302926A (en) * 1965-12-06 1967-02-07 Gen Electric Segmented nozzle diaphragm for high temperature turbine
US3326523A (en) * 1965-12-06 1967-06-20 Gen Electric Stator vane assembly having composite sectors
US3773430A (en) * 1972-03-17 1973-11-20 Ingersoll Rand Co Gas compressor

Family Cites Families (2)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US2857093A (en) * 1954-12-02 1958-10-21 Cincinnati Testing & Res Lab Stator casing and blade assembly
GB800098A (en) * 1955-10-31 1958-08-20 Rolls Royce Improvements in or relating to multi-stage axial-flow compressors

Patent Citations (4)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US2997275A (en) * 1959-03-23 1961-08-22 Westinghouse Electric Corp Stator structure for axial-flow fluid machine
US3302926A (en) * 1965-12-06 1967-02-07 Gen Electric Segmented nozzle diaphragm for high temperature turbine
US3326523A (en) * 1965-12-06 1967-06-20 Gen Electric Stator vane assembly having composite sectors
US3773430A (en) * 1972-03-17 1973-11-20 Ingersoll Rand Co Gas compressor

Cited By (18)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US5167488A (en) * 1991-07-03 1992-12-01 General Electric Company Clearance control assembly having a thermally-controlled one-piece cylindrical housing for radially positioning shroud segments
EP0763653A1 (en) * 1995-09-13 1997-03-19 SOCIETE DE CONSTRUCTION DES AVIONS HUREL-DUBOIS (société anonyme) Thrust reverser door with jet deflection cascade
US5671598A (en) * 1995-09-13 1997-09-30 Societe De Construction Des Avions Hurel-Dubois Forward mounted pivoting door reverser with efflux control device
US6234750B1 (en) 1999-03-12 2001-05-22 General Electric Company Interlocked compressor stator
US20070079506A1 (en) * 2005-10-06 2007-04-12 General Electric Company Method of providing non-uniform stator vane spacing in a compressor
US7743497B2 (en) * 2005-10-06 2010-06-29 General Electric Company Method of providing non-uniform stator vane spacing in a compressor
CN103982434A (zh) * 2005-10-06 2014-08-13 通用电气公司 在压气机中设置不均匀定子叶片间距的方法
US10280779B2 (en) 2013-09-10 2019-05-07 United Technologies Corporation Plug seal for gas turbine engine
US10287905B2 (en) 2013-11-11 2019-05-14 United Technologies Corporation Segmented seal for gas turbine engine
US10634055B2 (en) 2015-02-05 2020-04-28 United Technologies Corporation Gas turbine engine having section with thermally isolated area
US9920652B2 (en) 2015-02-09 2018-03-20 United Technologies Corporation Gas turbine engine having section with thermally isolated area
US20180066673A1 (en) * 2016-09-02 2018-03-08 United Technologies Corporation Repeating airfoil tip strong pressure profile
US11248622B2 (en) * 2016-09-02 2022-02-15 Raytheon Technologies Corporation Repeating airfoil tip strong pressure profile
US11773866B2 (en) 2016-09-02 2023-10-03 Rtx Corporation Repeating airfoil tip strong pressure profile
US20180230856A1 (en) * 2016-10-19 2018-08-16 United Technologies Corporation Engine cases and associated flange
US10550725B2 (en) * 2016-10-19 2020-02-04 United Technologies Corporation Engine cases and associated flange
US20220381150A1 (en) * 2021-05-26 2022-12-01 General Electric Company Split-line stator vane assembly
US11629606B2 (en) * 2021-05-26 2023-04-18 General Electric Company Split-line stator vane assembly

Also Published As

Publication number Publication date
NO149010C (no) 1984-01-25
DE2532554A1 (de) 1976-02-19
NO149010B (no) 1983-10-17
IT1040063B (it) 1979-12-20
FR2280812A1 (fr) 1976-02-27
CH589229A5 (enExample) 1977-06-30
GB1511019A (en) 1978-05-17
FR2280812B1 (enExample) 1980-09-26
DE2532554C2 (de) 1985-05-23
NO752644L (enExample) 1976-01-30
BR7504846A (pt) 1976-07-06
JPS6050999B2 (ja) 1985-11-11
SE7508371L (sv) 1976-01-30
CA1027870A (en) 1978-03-14
JPS5138112A (enExample) 1976-03-30
SE419260B (sv) 1981-07-20

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Owner name: FIRST NATIONAL BANK OF CHICAGO, THE, ONE FIRST NAT

Free format text: LICENSE;ASSIGNOR:ELLIOT TURBOMACHINERY CO., INC.;REEL/FRAME:004940/0562

Effective date: 19871109

Owner name: FIRST NATIONAL BANK OF CHICAGO, THE,ILLINOIS

Free format text: LICENSE;ASSIGNOR:ELLIOT TURBOMACHINERY CO., INC.;REEL/FRAME:004940/0562

Effective date: 19871109