US3640638A - Axial flow compressor - Google Patents

Axial flow compressor Download PDF

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US3640638A
US3640638A US50100A US3640638DA US3640638A US 3640638 A US3640638 A US 3640638A US 50100 A US50100 A US 50100A US 3640638D A US3640638D A US 3640638DA US 3640638 A US3640638 A US 3640638A
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Prior art keywords
shroud
rotor
ring
trailing edge
blades
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Expired - Lifetime
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US50100A
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Jack Britt
Alfred John Honey
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Rolls Royce PLC
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Rolls Royce PLC
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    • 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
    • F04D29/08—Sealings
    • F04D29/083—Sealings especially adapted for elastic fluid pumps
    • 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
    • F01D11/001—Preventing or minimising internal leakage of working-fluid, e.g. between stages for sealing space between stator blade and rotor
    • 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
    • F04D29/08—Sealings
    • F04D29/16—Sealings between pressure and suction sides
    • F04D29/161—Sealings between pressure and suction sides especially adapted for elastic fluid pumps
    • F04D29/164—Sealings between pressure and suction sides especially adapted for elastic fluid pumps of an axial flow wheel
    • 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
    • Y—GENERAL TAGGING OF NEW TECHNOLOGICAL DEVELOPMENTS; GENERAL TAGGING OF CROSS-SECTIONAL TECHNOLOGIES SPANNING OVER SEVERAL SECTIONS OF THE IPC; TECHNICAL SUBJECTS COVERED BY FORMER USPC CROSS-REFERENCE ART COLLECTIONS [XRACs] AND DIGESTS
    • Y10—TECHNICAL SUBJECTS COVERED BY FORMER USPC
    • Y10S—TECHNICAL SUBJECTS COVERED BY FORMER USPC CROSS-REFERENCE ART COLLECTIONS [XRACs] AND DIGESTS
    • Y10S415/00—Rotary kinetic fluid motors or pumps
    • Y10S415/914—Device to control boundary layer

Abstract

An axial flow compressor comprises a rotor having a plurality of axially spaced rings of rotor blades, a respective ring of stator blades being disposed immediately downstream of each ring of rotor blades, respective shroud surfaces being provided at the radially inner ends of the working surfaces of the stator blades by respective shroud rings carried by the rings of stator blades, running clearances being provided between the shroud rings and the rotor, and means whereby in operation a static pressure gradient is maintained in the clearance between at least one shroud ring and the rotor so that gas flow through the clearance is in a downstream direction.

Description

United States Patent Britt et al.
[ Feb. 8, 1972 [54] AXIAL FLOW COMPRESSOR [72] Inventors: Jack Britt, Ambergate; Alfred John Honey, Allestree, both of England [22] Filed: June 26,1970
[21] Appl.No.: 50,100
2,848,155 8/1958 Hausmann ..415/D1G. 1 3,300,121 l/1967 Johnson.... 3,251,601 5/1966 Harvey ..415/172 FORElGN PATENTS OR APPLICATIONS 10,179 1912 Great Britain ..415/172 Primary Examiner-Henry F. Raduazo Att0meyCushman, Darby & Cushman ABSTRACT An axial flow compressor comprises a rotor having a plurality of axially spaced rings of rotor blades, a respective ring of stator blades being disposed immediately downstream of each ring of rotor blades, respective shroud surfaces being provided at the radially inner ends of the working surfaces of the stator blades by respective shroud rings carried by the rings of stator blades, running clearances being provided between the shroud rings and the rotor, and means whereby in operation a static pressure gradient is maintained in the clearance between at least one shroud ring and the rotor so that gas flow through the clearance is in a downstream direction.
5 Claims, 2 Drawing Figures PAIENTEDFEB 81912 3.640.539
MMM W ttorneyg AXIAL FLOW COMPRESSOR This invention relates to axial flow compressors.
The compression ratio and therefore the delivery gas temperature achieved in axial flow compressors is steadily increasing and is now such that nickel base alloys have to be considered for the compressor rotors. We have now discovered that this problem is aggravated because the actual temperature reached by the structure of a stage of an axial flow compressor is greater than the mean temperature of the gas as it passes through that stage.
We believe that a reason for this anomaly is the recirculation of gas via the running clearances between the stator shroud rings and the rotor, so that the compressor stage is heated by hotter gas recirculated from a downstream stage.
Therefore, in one aspect the invention provides an axial flow compressor comprising a rotor having a plurality of axially spaced rings of rotor blades, a respective ring of stator blades being disposed immediately downstream of each ring of rotor blades, respective shroud surfaces being provided at the radially inner ends of the working surfaces of the stator blades by respective shroud rings carried by the rings of stator blades, running clearances being provided between the shroud rings and the rotor, and means whereby in operation a static pressure gradient is maintained in the clearance between at last one shroud ring and the rotor so that gas flow through the clearance is in a downstream direction.
The said means may be adapted to maintain the static pressure gradient by increasing the static pressure of a boundary layer flow passing to the leading edge of the shroud surface of the at least one shroud ring from a rotor surface at the roots of the working surfaces of the ring of rotor blades immediately upstream of the at least one shroud ring, so that said boundary layer static pressure is greater than the boundary layer static pressure at the trailing edge of the shroud surface of the at least one shroud ring.
The said means may be adapted to effect said increase in boundary layer static pressure by converting at least a portion of the dynamic pressure of the boundary layer at the said upstream edge to static pressure.
The trailing edge of the said rotor surface may be of smaller diameter than the leading edge of the shroud surface of the at least one shroud ring, so that in operation the leading edge of the said shroud surface projects radially outwardly of the said trailing edge of the rotor surface and intercepts the boundary layer flow passing therefrom.
A trailing edge of the shroud surface of the at least one shroud ring may be of greater diameter than that of a leading edge of a further rotor surface at the roots of the working surfaces of a ring of rotor blades immediately downstream of the at least one shroud ring so that in operation the trailing edge of the said shroud surface projects radially outwardly of the said leading edge of the further rotor surface.
Sealing means may be provided in said clearance between the at least one shroud ring and the rotor to reduce the gas flow therethrough.
The at least one shroud ring may be received with said clearance in an annular pocket in the rotor between the ring of rotor blades immediately upstream thereof and a ring of rotor blades immediately downstream thereof, said sealing means being provided between a wall of the pocket and the shroud ring.
In another aspect, although not so restricted, the invention provides a gas turbine engine having an axial flow compressor as set forth above.
The invention will be described, merely by way of example, with reference to the accompanying drawings, wherein:
FIG. 1 shows, partly in section, a gas turbine engine embodying a compressor according to the invention, and
FIG. 2 shows a part of the structure of FIG. 1.
Referring to FIG. 1, a gas turbine engine comprises an inlet 10, an axial flow compressor 12, a combustion section 14, a turbine section 16 and an exhaust noule 18.
Attention is now directed to FIG. 2, which although showing a compressor according to the invention will initially be used to illustrate the above-mentioned problem as it occurs in conventional compressors.
Conventionally, axial flow compressors comprise a rotor 20 having a plurality of axially spaced rings of rotor blades two of which rings are shown at 22, 24. A respective ring of stator blades is disposed downstream of each ring of rotor blades. Such a ring of stator blades is shown at 26. A shroud ring 28 is carried by the ring of stator blades 26 at the radially inner ends of the working surfaces of the stator blades. A running clearance is provided between the shroud ring 28 and the rotor 20; in the case of shroud rings 28 of intermediate stator blade rings such as 26, the running clearance typically is provided by means of an annular pocket 30 between the rings of rotor blades 22, 24 respectively immediately upstream and downstream of the ring of stator blades 26.
Surfaces (e.g., 32, 33) are provided on the rotor 20 at the roots of the working surfaces of each ring of rotor blades 22, 24, and the shroud rings 28 are provided with shroud surfaces 34, the surfaces 32, 33, 34 defining the radially inner wall of the gas glow duct through the compressor.
The function of stator blades in an axial flow compressor of course is to convert part of the dynamic head of the gas flow leaving the preceding rotor blades to a static head, whereby the pressure of the gas is increased successively from stage to stage. By virtue of the function of the stator blades, the static pressure of the gas at the trailing edges 36 of the ring of stator blades 26 is greater than the static head at the leading edges 38 thereof. The gas temperature is also greater since the conversion to static pressure of a dynamic head due to a gas velocity of V results in an increase in gas temperature of V IZC J, where C, is the specific heat at constant pressure of the gas, J is the mechanical equivalent of heat and g is the acceleration due to gravity.
We believe that the boundary layer flow across the surfaces 32, 34 suffers increases in temperature in the following way. The boundary layer flow approaching the leading edge of the surface 32 from an upstream stator shroud ring (not shown) has a velocity which is markedly different from the free stream velocity for which the blades of the rotor stage 22 are designed. Thus, the boundary layer suffers a violent change in velocity as it reaches the surface 32, resulting in an increase in temperature of the boundary layer.
The boundary layer flow then passes over the surface 32, and during this time has a large circumferential component of velocity and a small axial component of velocity. As the boundary layer flow leaves the trailing edge of the surface 32 it passes to the leading edge 40 of the stationary shroud surface 34, and since its velocity is again markedly different from the free stream velocity for which the stator blades 26 are designed, due to its small axial component, the boundary layer suffers another violent change of velocity, and its temperature is further increased.
Upon leaving the trailing edge 39 of the shroud surface 34, the boundary layer flow impinges upon the leading edge of the surface 33, as described above in relation to the surface 32, and is thus subjected to a yet further violent change in velocity and consequent rise in temperature. The rise in boundary layer temperature caused in this way may be equivalent to as many as three dynamic pressures.
The static pressure difference between the trailing 36 and leading 38 edges of the ring of stator blades 26 diverts a portion of the hot boundary layer leaving the trailing edge 39 through the clearance 30 to the leading edge 40. This recirculated portion of hot boundary layer flow is discharged at the leading edge 40 and thus is again subjected to the violent velocity changes at the edges 40, 39, and is further raised in temperature.
Some of the boundary layer flow even may be recirculated for a second time, further raising its temperature.
Considering now the novel features of FIG. 2, we provide means whereby in operation of the compressor a static presme-v sure gradient is maintained in the clearance 30 so that leakage gas flow is in a downstream direction, as indicated by the arrows 42. The static pressure gradient is maintained by increasing the static pressure of the boundary layer flow passing from the trailing edge 44 of the rotor surface 32 to the leading edge 40 of the shroud surface 34, so that the boundary layer static pressure is greater at the leading edge 40 than at the trailing edge 39 of the shroud surface 34.
To effect this increase in boundary layer static pressure, the diameter of the rotor surface 32 at its trailing edge 44 is chosen to be smaller than the diameter of the leading edge 40 of the shroud surface 34, so that when the compressor is operating at its normal speed the leading edge 40 projects radially outwardly of the trailing edge 44 and intercepts the boundary layer flow passing therefrom. The extent of the projection is indicated by the dimension referenced 46. This projection 46 converts a sufficient proportion of the dynamic pressure of the boundary layer flow to static pressure to maintain the required pressure gradient.
Thus, the boundary layer flow leaving the surface 32 still is subjected to a temperature rise due to the reduction in velocity when it impinges upon the leading edge 40, but at least some of the boundary layer flow then passes in a downstream direction through the clearance 30. While passing through the clearance 30, the flow tends to be entrained gradually by the rotating walls of the pocket defining the clearance 30. Thus, the flow emerging from the clearance 30 at the downstream end thereof has a substantial circumferential component of velocity, and the velocity change when the flow impinges on the leading edge of the surface 33 may be less violent and may cause a smaller rise in temperature.
The novel construction of FIG. 2 may result in a reduction in the rise in temperature of the boundary layer between the rings of rotor blades 22, 24 equivalent to at least one dynamic pressure.
A seal 48 is provided in the clearance 30 between the walls of the pocket defining the clearance and the shroud ring 28, since it is desirable to reduce the flow through the clearance 30. An appreciable flow through the clearance 30 in the downstream direction would disturb the gas velocities between the ring of stator blades 26 and the next ring of rotor blades 24, reducing compressor efficiency.
The trailing edge 39 of the shroud surface 34 is of greater diameter than the leading edge 50 of the rotor surface 33 so that in operation it projects radially outwardly thereof, as evidenced by the dimension 52. The boundary layer gas flow leaving the trailing edge 39 thus possibly may have an extractor of jet pump effect on the gas in the clearance 30, thus augmenting the pressure gradient in the clearance 30.
It will be appreciated that a compressor according to the invention need not have a pressure gradient maintained to produce a downstream flow in every running clearance between stator blade shroud rings and the rotor; for example, it may be sufficient to provide a pressure gradient only in one or more clearances, e.g., around the hottest stator blades shroud rings at the delivery end of the compressor.
We claim:
1. An axial flow compressor comprising a rotor having a plurality of axially spaced rings of rotor blades each having working surfaces and a root, said roots defining a rotor surface, a respective ring of stator blades being disposed immediately downstream of each ring of rotor blades, each of said stator blades having working surfaces and radially inner ends, each of said radially inner ends carrying a shroud ring,
said shroud rings defining respective shroud surfaces, running clearances being provided between said shroud rings and said rotor, a said rotor surface being disposed immediately upstream of at least one of a said shroud ring, means for maintaining a static pressure gradient in the clearance between at least said one shroud ring and said rotor whereby gas flow through said clearance will be in a downstream direction, said means including said rotor surface having a trailing edge and said shroud surface of said at least one shroud ring having a leading and trailing edge, said trailing edge of said rotor surface having a smaller diameter than said leading edge of said shroud surface, said leading edge of the said shroud surface projecting radially outwardly of the said trailing edge of said rotor surface so that the boundary layer flow of gas passing from said trailing edge of said rotor surface will be intercepted by said leading edge of said shroud surface to increase the static pressure of the boundary layer flow passing to said leading edge of said shroud surface relative to the boundary layer static pressure at said trailing edge of said shroud surface of the at least one shroud ring.
2. A compressor as claimed in claim 1 wherein a ring of rotor blades immediately downstream of the at least one shroud ring has a further rotor surface at the roots of the working surfaces of the blades, the leading edge of the rotor surface being of smaller diameter than the trailing edge of the said shroud surface, said trailing edge projecting radially outwardly of the said leading edge of the further rotor surface.
3. A compressor as claimed in claim 1 wherein sealing means are provided in said clearance between the at least one shroud ring and the rotor, reducing the gas flow therethrough.
4. A compressor as claimed in claim 3 wherein the rotor has an annular pocket between the ring of rotor blades immediately upstream of the at least one shroud ring and a ring of rotor blades immediately downstream of the at least one shroud ring, the at least one shroud ring being received with said clearance in the annular pocket, said sealing means being provided between a wall of the pocket and the shroud ring.
5. A gas turbine engine having an axial flow compressor comprising a rotor having a plurality of axially spaced rings of rotor blades each having working surfaces and a root, said roots defining a rotor surface, a respective ring of stator blades being disposed immediately downstream of each ring of rotor blades, each of said stator blades having working surfaces and radially inner ends, each of said radially inner ends carrying a shroud ring, said shroud rings defining respective shroud surfaces, running clearances being provided between said shroud rings and said rotor, a said rotor surface being disposed immediately upstream of at least one of a said shroud ring, means for maintaining a static pressure gradient in the clearance between at least said one shroud ring and said rotor whereby gas flow through said clearance will be in a downstream direction, said means including said rotor surface having a trailing edge and said shroud surface of said at least one shroud ring having a leading and trailing edge, said trailing edge of said rotor surface having a smaller diameter than said leading edge of said shroud surface, said leading edge of the said shroud surface projecting radially outwardly of the said trailing edge of said rotor surface so that the boundary layer flow of gas passing from said trailing edge of said rotor surface will be intercepted by said leading edge of said shroud surface to increase the static pressure of the boundary layer flow passing to said leading edge of said shroud surface relative to the boundary layer static pressure at said trailing edge of said shroud surface of the at least one shroud ring.

Claims (5)

1. An axial flow compressor comprising a rotor having a plurality of axially spaced rings of rotor blades each having working surfaces and a root, said roots defining a rotor surface, a respective ring of stator blades being disposed immediately downstream of each ring of rotor blades, each of said stator blades having working surfaces and radially inner ends, each of said radially inner ends carrying a shroud ring, said shroud rings defining respective shroud surfaces, running clearances being provided between said shroud rings and said rotor, a said rotor surface being disposed immediately upstream of at least one of a said shroud ring, means for maintaining a static pressure gradient in the clearance between at least said one shroud ring and said rotor whereby gas flow through said clearance will be in a downstream direction, said means including said rotor surface having a trailing edge and said shroud surface of said at least one shroud ring having a leading and trailing edge, said trailing edge of said rotor surface having a smaller diameter than said leading edge of said shroud surface, said leading edge of the said shroud surface projecting radially outwardly of the said trailing edge of said rotor surface so that the boundary layer flow of gas passing from said trailing edge of said rotor surface will be intercepted by said leading edge of said shroud surface to increase the static pressure of the boundary layer flow passing to said leading edge of said shroud surface relative to the boundary layer static pressure at said trailing edge of said shroud surface of the at least one shroud ring.
2. A compressor as claimed in claim 1 wherein a ring of rotor blades immediately downstream of the at least one shroud ring has a further rotor surface at the roots of the working surfaces of the blades, the leading edge of the rotor surface being of smaller diameter than the trailing edge of the said shroud surface, said trailing edge projecting radially outwardly of the said leading edge of the further rotor surface.
3. A compressor as claimed in claim 1 wherein sealing means are provided in said clearance between the at least one shroud ring and the rotor, reducing the gas flow therethrough.
4. A compressor as claimed in claim 3 wherein the rotor has an annular pocket between the ring of rotor blades immediately upstream of the at least one shroud ring and a ring of rotor blades immediately downstream of the at least one shroud ring, the at least one shroud ring being received with said clearance in the annular pocket, said sealing means being provided between a wall of the pocket and the shroud ring.
5. A gas turbine engine having an axial flow compressor comprising a rotor having a plurality of axially spaced rings of rotor blades each having working surfaces and a root, said roots defining a rotor surface, a respective ring of stator blades being disposed immediately downstream of each ring of rotor blades, each of said stator blades having working surfaces and radially inner ends, each of said radially inner ends carrying a shroud ring, said shroud rings defining respective shroud Surfaces, running clearances being provided between said shroud rings and said rotor, a said rotor surface being disposed immediately upstream of at least one of a said shroud ring, means for maintaining a static pressure gradient in the clearance between at least said one shroud ring and said rotor whereby gas flow through said clearance will be in a downstream direction, said means including said rotor surface having a trailing edge and said shroud surface of said at least one shroud ring having a leading and trailing edge, said trailing edge of said rotor surface having a smaller diameter than said leading edge of said shroud surface, said leading edge of the said shroud surface projecting radially outwardly of the said trailing edge of said rotor surface so that the boundary layer flow of gas passing from said trailing edge of said rotor surface will be intercepted by said leading edge of said shroud surface to increase the static pressure of the boundary layer flow passing to said leading edge of said shroud surface relative to the boundary layer static pressure at said trailing edge of said shroud surface of the at least one shroud ring.
US50100A 1969-07-02 1970-06-26 Axial flow compressor Expired - Lifetime US3640638A (en)

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Cited By (4)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US4871294A (en) * 1982-06-29 1989-10-03 Ivanov Sergei K Axial-flow fan
US5116200A (en) * 1990-06-28 1992-05-26 General Electric Company Apparatus and methods for minimizing vibrational stresses in axial flow turbines
EP2003292A3 (en) * 2007-06-14 2012-04-04 Rolls-Royce Deutschland Ltd & Co KG Blade shroud with overhang
US20160305264A1 (en) * 2013-12-05 2016-10-20 United Technologies Corporation Turbomachine rotor-stator seal

Citations (6)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
GB191210179A (en) * 1911-05-04 1912-06-20 Heinrich Holzer Arrangement for Diminishing Clearance Losses in Turbines and Pumps for Liquids and Elastic Fluids.
US2720356A (en) * 1952-06-12 1955-10-11 John R Erwin Continuous boundary layer control in compressors
US2831629A (en) * 1952-05-06 1958-04-22 Westinghouse Electric Corp Compressor apparatus
US2848155A (en) * 1950-11-22 1958-08-19 United Aircraft Corp Boundary layer control apparatus for compressors
US3251601A (en) * 1963-03-20 1966-05-17 Gen Motors Corp Labyrinth seal
US3300121A (en) * 1965-02-24 1967-01-24 Gen Motors Corp Axial-flow compressor

Patent Citations (6)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
GB191210179A (en) * 1911-05-04 1912-06-20 Heinrich Holzer Arrangement for Diminishing Clearance Losses in Turbines and Pumps for Liquids and Elastic Fluids.
US2848155A (en) * 1950-11-22 1958-08-19 United Aircraft Corp Boundary layer control apparatus for compressors
US2831629A (en) * 1952-05-06 1958-04-22 Westinghouse Electric Corp Compressor apparatus
US2720356A (en) * 1952-06-12 1955-10-11 John R Erwin Continuous boundary layer control in compressors
US3251601A (en) * 1963-03-20 1966-05-17 Gen Motors Corp Labyrinth seal
US3300121A (en) * 1965-02-24 1967-01-24 Gen Motors Corp Axial-flow compressor

Cited By (4)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US4871294A (en) * 1982-06-29 1989-10-03 Ivanov Sergei K Axial-flow fan
US5116200A (en) * 1990-06-28 1992-05-26 General Electric Company Apparatus and methods for minimizing vibrational stresses in axial flow turbines
EP2003292A3 (en) * 2007-06-14 2012-04-04 Rolls-Royce Deutschland Ltd & Co KG Blade shroud with overhang
US20160305264A1 (en) * 2013-12-05 2016-10-20 United Technologies Corporation Turbomachine rotor-stator seal

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FR2056433A5 (en) 1971-05-14
DE2032380A1 (en) 1971-02-04

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