WO2013119520A1 - Customer bleed air pressure loss reduction - Google Patents

Customer bleed air pressure loss reduction Download PDF

Info

Publication number
WO2013119520A1
WO2013119520A1 PCT/US2013/024697 US2013024697W WO2013119520A1 WO 2013119520 A1 WO2013119520 A1 WO 2013119520A1 US 2013024697 W US2013024697 W US 2013024697W WO 2013119520 A1 WO2013119520 A1 WO 2013119520A1
Authority
WO
WIPO (PCT)
Prior art keywords
duct
insert
upstream
housing
compressor
Prior art date
Application number
PCT/US2013/024697
Other languages
French (fr)
Inventor
Cheng-Zhang Wang
Thomas G. Phillips
David F. CLOUD
Peter M. MUNSELL
Original Assignee
United Technologies Corporation
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 Corporation filed Critical United Technologies Corporation
Priority to EP13747274.2A priority Critical patent/EP2812548B1/en
Priority to EP20162505.0A priority patent/EP3690209A1/en
Publication of WO2013119520A1 publication Critical patent/WO2013119520A1/en

Links

Classifications

    • 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
    • F02C6/00—Plural gas-turbine plants; Combinations of gas-turbine plants with other apparatus; Adaptations of gas-turbine plants for special use
    • F02C6/04—Gas-turbine plants providing heated or pressurised working fluid for other apparatus, e.g. without mechanical power output
    • F02C6/06—Gas-turbine plants providing heated or pressurised working fluid for other apparatus, e.g. without mechanical power output providing compressed gas
    • F02C6/08—Gas-turbine plants providing heated or pressurised working fluid for other apparatus, e.g. without mechanical power output providing compressed gas the gas being bled from the gas-turbine compressor
    • 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
    • F01D9/023—Transition ducts between combustor cans and first stage of the turbine in gas-turbine engines; their cooling or sealings
    • 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
    • F01D9/06—Fluid supply conduits to nozzles or the like
    • F01D9/065—Fluid supply or removal conduits traversing the working fluid flow, e.g. for lubrication-, cooling-, or sealing fluids
    • 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
    • F02C7/12—Cooling of plants
    • F02C7/16—Cooling of plants characterised by cooling medium
    • F02C7/18—Cooling of plants characterised by cooling medium the medium being gaseous, e.g. air
    • F02C7/185—Cooling means for reducing the temperature of the cooling air or gas
    • 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
    • F02C9/00—Controlling gas-turbine plants; Controlling fuel supply in air- breathing jet-propulsion plants
    • F02C9/16—Control of working fluid flow
    • F02C9/18—Control of working fluid flow by bleeding, bypassing or acting on variable working fluid interconnections between turbines or compressors or their stages
    • 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
    • F05D2260/00—Function
    • F05D2260/60—Fluid transfer
    • F05D2260/601—Fluid transfer using an ejector or a jet pump
    • 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
    • F05D2270/00—Control
    • F05D2270/01—Purpose of the control system
    • F05D2270/17—Purpose of the control system to control boundary layer
    • 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
    • F05D2270/00—Control
    • F05D2270/30—Control parameters, e.g. input parameters
    • F05D2270/301—Pressure
    • 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
    • Y02—TECHNOLOGIES OR APPLICATIONS FOR MITIGATION OR ADAPTATION AGAINST CLIMATE CHANGE
    • Y02T—CLIMATE CHANGE MITIGATION TECHNOLOGIES RELATED TO TRANSPORTATION
    • Y02T50/00—Aeronautics or air transport
    • Y02T50/60—Efficient propulsion technologies, e.g. for aircraft

Definitions

  • This application relates to a system for reducing pressure loss on bleed air systems for tapping air away from a gas turbine engine for use on an associated aircraft.
  • Gas turbine engines for use on aircraft typically include a fan delivering air into a compressor. Air from the compressor is directed into a combustion section where it is mixed with fuel and burned. Products of this combustion pass downstream over turbine rotors, causing them to rotate and power the fan and compressor rotors.
  • gas turbine engines When gas turbine engines are utilized on aircraft, they are also the source of air for various uses on the aircraft. As examples, cabin air, cooling air, or air for any number of other applications are tapped from the gas turbine engine.
  • air that has been at least partially compressed is utilized.
  • the air is initially taken from a port downstream of the entire compressor section, and upstream of the combustion section.
  • the air is taken from this high pressure port when the engine is at low thrust.
  • the pressure at this port will rise until a high pressure shutoff valve closes.
  • air is tapped from a port at an intermediate location in the compressor section.
  • a bleed air supply system for a gas turbine engine has a duct having an inlet end and extending to an outlet end. An inlet end of the duct is provided with a central insert at an upstream end. The insert ends within the duct upstream of the outlet end.
  • the central insert provides a venturi effect by reducing the cross-sectional flow area between the insert and an inner wall of the duct of the upstream end.
  • the insert and duct provide increased cross-sectional flow areas at downstream locations.
  • the insert has rounded axial ends.
  • the air supply system includes a plurality of ducts, each being provided with an insert.
  • the ducts have inlet ends at locations spaced by at least 90° about a cross-sectional center axis of a gas turbine engine which is to receive the air supply system.
  • the inlet ends are spaced by 180°.
  • a plurality of insert holders center the insert within the duct.
  • a gas turbine engine has a compressor delivering air into a combustion section.
  • the combustion section and the compressor are housed in a housing.
  • An air supply system communicates through the housing to deliver air from a location between an upstream end of the compressor, and an upstream end of the combustor.
  • the air supply system has a duct with an inlet end extending to an outlet end.
  • the duct is provided with a central insert at an upstream end, ending within the duct upstream of the outlet end.
  • the central insert provides a venturi effect by reducing the cross-sectional flow between the insert and an inner wall of the duct at the upstream end.
  • the insert and duct provide increased cross- sectional flow areas at downstream locations.
  • the insert has rounded axial ends.
  • the air supply system includes a plurality of ducts.
  • Each of the ducts have inlet ends at locations spaced by at least 90° about a cross-sectional center axis of the gas turbine engine.
  • the inlet ends are spaced by 180°.
  • a plurality of insert holders center the insert within the duct.
  • a diffuser is positioned downstream of the compressor.
  • An opening in the housing supplies air to the inlet end.
  • the diffuser has an outer shroud and an inner shroud with intermediate vanes. The outer shroud ends at a location upstream of a downstream end of the inner shroud at locations circumferentially aligned with the inlet end.
  • the outer shroud only ends at the upstream location at circumferential locations associated with the inlet end, but extends further downstream at other locations.
  • a flow path extends through the opening, and into the inlet end of the duct which has at least a portion formed with a part-circular radius.
  • the opening in the housing leads into the inlet end of the duct, with an inlet end ending downstream of the opening in the housing, with the part-circular radius portion formed in the housing.
  • an opening is formed through the housing and into the inlet end of the duct wherein a flow path extends through the opening, and into the inlet end of the duct which has at least a portion formed with a part-circular radius.
  • the opening in the housing leads into the inlet end of the duct.
  • the inlet end ends downstream of the opening in the housing, with the part-circular radius portion formed in the housing.
  • a gas turbine engine has a compressor delivering air into a combustion section.
  • the combustion section and compressor are housed in a housing.
  • An air supply system communicates through the housing to deliver air from a location between an upstream end of the compressor, and an upstream end of the combustor.
  • the air supply system has a duct with an inlet end extending to an outlet end.
  • the duct is provided with a central insert at an upstream end.
  • the central insert provides a venturi effect by reducing the cross-sectional flow between the insert and an inner wall of the duct at the upstream end.
  • the insert and duct provide increased cross-sectional flow areas at downstream locations.
  • the air supply system includes a plurality of ducts, each being provided with an insert.
  • the ducts inlet ends at locations spaced by at least 90° about a cross-sectional center axis of the gas turbine engine.
  • a diffuser is positioned downstream of the compressor with an opening in the housing supplying air to the inlet end.
  • the diffuser has an outer shroud and an inner shroud with intermediate vanes.
  • the outer shroud ends at a location upstream of a downstream end of the inner shroud at locations circumferentially aligned with the inlet end.
  • the outer shroud only ends at the upstream location at circumferential locations associated with the inlet end, but extends further downstream at other locations.
  • a gas turbine engine has a compressor delivering air into a combustion section.
  • the combustion section and compressor are housed in a housing.
  • An air supply system communicates through the housing to deliver air from a location between an upstream end of the compressor and an upstream end of the combustor.
  • the air supply system has a plurality of ducts to move air towards a use.
  • the ducts have inlet ends at locations spaced by at least 90° about a cross-sectional center axis of the gas turbine engine.
  • a gas turbine engine has a compressor delivering air into a combustion section.
  • the combustion section and compressor are housed in a housing.
  • An air supply system communicates through the housing to deliver air from a location between an upstream end of the compressor, and an upstream end of the combustor.
  • the air supply system has a duct with an inlet end.
  • a diffuser is positioned downstream of the compressor, and has an outer shroud and an inner shroud with intermediate vanes. The outer shroud ends at a location upstream of the downstream end of the inner shroud at locations circumferentially aligned with the inlet end.
  • the outer shroud only ends at the upstream location at circumferential locations associated with the inlet ends, but extends further downstream at other locations.
  • Figure 1 shows a prior art gas turbine engine.
  • Figure 2A shows a prior art air bleed arrangement.
  • Figure 2B shows a portion of the Figure 2A arrangement.
  • Figure 3 shows the location of one part of the prior art arrangement.
  • Figure 4 shows a tap location in the present application.
  • Figure 5 shows detail of a diffuser incorporated into this application.
  • Figure 6 A shows a first embodiment arrangement of ports.
  • Figure 6B shows a second embodiment arrangement of ports.
  • Figure 7A shows a feature of a venturi duct.
  • Figure 7B shows an alternative embodiment
  • Figure 7C shows another feature of the Figure 7A and 7B embodiments.
  • Figure 8 shows yet another feature. DETAILED DESCRIPTION
  • Figure 1 shows a prior art gas turbine engine.
  • a gas turbine engine 10 such as a turbofan gas turbine engine, circumferentially disposed about an engine centerline A, is shown in Figure 1.
  • the engine 10 includes a fan 18, a compressor 12, a combustion section 14 and turbine sections 16.
  • air compressed in the compressor 12 is mixed with fuel which is burned in the combustion section 14 and expanded across turbines 16.
  • the turbines includes rotors that rotate in response to the expansion, driving compressor rotors and fan 18.
  • This structure is shown somewhat schematically in Figure 1. While one example gas turbine engine is illustrated, it should be understood this invention extends to any other type gas turbine engine for any application.
  • FIG. 2A shows a bleed air arrangement 30.
  • a downstream tap port 34 which is associated with a location downstream of the compressor section 12, but upstream of the combustor 14 (see Figure 3).
  • air is tapped through port 34, and delivered to an outlet 32, which communicates with various uses 200 for air on the aircraft.
  • valves 142 shut off the flow from the port 34, and switch to a port 36 which is upstream of the port 34, and at an intermediate location in the compressor section 12.
  • FIG. 2B shows that there are actually two of the ports 34 in a typical prior art arrangement which are spaced circumferentially about a central line of the engine.
  • the ports 34 are spaced by a relatively small angle, and typically on the order of 25° to 30°.
  • Figure 3 shows the location of the port 34, associated with a location 40 downstream of a diffuser 44, which is the downstream end of compressor section 12. As shown, an opening 42 in a housing 41 communicates air into the duct port. Downstream of the area 40 is the combustor 14.
  • Figure 4 shows a diffuser embodiment 118 wherein an outer shroud 58 of the diffuser ends upstream of a radially inner shroud 54. As shown, vanes 56 extend between the shroud walls 54 and 58.
  • the duct 50 communicates with an opening 42. Further details of this duct will be disclosed below.
  • Figure 5 shows a portion of the diffuser 118. As shown, the outer shroud is cut upstream, at locations 58 associated with the opening 42, but otherwise extends forwardly 158 to the location of the prior art diffuser as shown, for example, in Figure 3. In embodiments, there are plural ducts 50 and openings 42, and the cutaway locations 58 are associated with each opening 42.
  • Figure 6A shows an arrangement wherein the ports 64 and 66 communicate with a common duct 62, and are spaced by approximately 180° about a centerline X.
  • Figure 6B shows an alternative arrangement wherein there are ports 68 and 72.
  • the ports 68 and 72 are spaced by an angle A, which in this embodiment would be approximately 90°. In embodiments, it is preferred that the ports are spaced by at least 90° to minimize pressure loss.
  • FIG. 300 there could be a third port incorporated at a smaller angle. While the use of ports spaced by at least 90° is a feature of this combination, it should be clear from the Figure 6B, that the other features of this application could be utilized in a system wherein the ports are spaced closer, such as shown by 300 in Figure 6B, or as shown in Figure 2B.
  • Figure 7A shows a duct 50 which includes an insert 82 at an upstream end 79.
  • the insert 82 creates a venturi with a relatively small cross-sectional flow 80 at the upstream end, and increasing to a larger flow 86.
  • Ends of the insert 84 and 184 are spherical to reduce pressure losses associated with air flowing along those surfaces.
  • a feature provided by the insert is that flow separation will be prevented since the flow would be through an annular area between the insert and the inner wall of the duct. This and the venturi effect result in the reduced pressure losses.
  • Figure 7B shows the use of the insert 182 in a somewhat alternative duct embodiment 150.
  • the insert and duct in the Figure 7 A embodiment bend at 88, while the Figure 7B embodiment extends generally linearly.
  • an upstream flow cross-sectional area Ai defined between the outer periphery of the insert 182 and the inner periphery of the duct 150 is much smaller than a downstream cross-sectional flow area A 2 again defined between the insert and the inner wall of the duct. This creates a venturi effect.
  • the insert ends at an intermediate location within the ducts 50 or 150, and will end before the outlet 400 of the duct 50, which communicates into other portions of the air supply system.
  • Figure 7C shows features of the Figure 7A or 7B embodiments wherein insert holders 90 mount the insert within the duct 50. In this embodiment there are three insert holders, each spaced by 120°.
  • Figure 8 shows another feature wherein duct 50 ends at an upstream end 250.
  • the opening 142 in the housing 143 has a lead in radius R] communicating into the duct 50.
  • the radius Ri may be .25" (0.6 cm) in one embodiment. Of course, other radii may be used. By having a part circular radius leading into the duct, the pressure losses are also reduced.

Landscapes

  • Engineering & Computer Science (AREA)
  • Chemical & Material Sciences (AREA)
  • Combustion & Propulsion (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Physics & Mathematics (AREA)
  • Fluid Mechanics (AREA)
  • Structures Of Non-Positive Displacement Pumps (AREA)

Abstract

A bleed air supply system for a gas turbine engine comprising a duct having an inlet end and extending to an outlet end. The inlet end of the duct is provided with a central insert. In another feature, there may be a plurality of ducts, and inlet ends of the plurality of ducts being spaced by at least 90. In another feature, a compressor may have a diffuser with a shroud ending upstream of the downstream end of an inner shroud, having an outer shroud ending at a location upstream of a downstream end of an inner shroud at locations circumferentially aligned with an inlet end of the duct.

Description

CUSTOMER BLEED AIR PRESSURE LOSS REDUCTION
BACKGROUND OF THE INVENTION
[0001] This application relates to a system for reducing pressure loss on bleed air systems for tapping air away from a gas turbine engine for use on an associated aircraft.
[0002] Gas turbine engines for use on aircraft typically include a fan delivering air into a compressor. Air from the compressor is directed into a combustion section where it is mixed with fuel and burned. Products of this combustion pass downstream over turbine rotors, causing them to rotate and power the fan and compressor rotors.
[0003] When gas turbine engines are utilized on aircraft, they are also the source of air for various uses on the aircraft. As examples, cabin air, cooling air, or air for any number of other applications are tapped from the gas turbine engine.
[0004] Typically, air that has been at least partially compressed is utilized. In many applications, the air is initially taken from a port downstream of the entire compressor section, and upstream of the combustion section. The air is taken from this high pressure port when the engine is at low thrust. As thrust increases, the pressure at this port will rise until a high pressure shutoff valve closes. Thereafter, air is tapped from a port at an intermediate location in the compressor section.
[0005] The pressure loss near the ports raises challenges with regard to providing sufficient air without decreasing the efficiency of the associated aircraft.
SUMMARY OF THE INVENTION
[0006] In a featured embodiment, a bleed air supply system for a gas turbine engine has a duct having an inlet end and extending to an outlet end. An inlet end of the duct is provided with a central insert at an upstream end. The insert ends within the duct upstream of the outlet end.
[0007] In another embodiment according to the foregoing embodiment, the central insert provides a venturi effect by reducing the cross-sectional flow area between the insert and an inner wall of the duct of the upstream end. The insert and duct provide increased cross-sectional flow areas at downstream locations.
[0008] In another embodiment according to the foregoing embodiment, the insert has rounded axial ends.
[0009] In another embodiment according to the foregoing embodiment, the air supply system includes a plurality of ducts, each being provided with an insert. The ducts have inlet ends at locations spaced by at least 90° about a cross-sectional center axis of a gas turbine engine which is to receive the air supply system.
[0010] In another embodiment according to the foregoing embodiment, the inlet ends are spaced by 180°.
[0011] In another embodiment according to the foregoing embodiment, a plurality of insert holders center the insert within the duct.
[0012] In another featured embodiment, a gas turbine engine has a compressor delivering air into a combustion section. The combustion section and the compressor are housed in a housing. An air supply system communicates through the housing to deliver air from a location between an upstream end of the compressor, and an upstream end of the combustor. The air supply system has a duct with an inlet end extending to an outlet end. The duct is provided with a central insert at an upstream end, ending within the duct upstream of the outlet end.
[0013] In another embodiment according to the foregoing embodiment, the central insert provides a venturi effect by reducing the cross-sectional flow between the insert and an inner wall of the duct at the upstream end. The insert and duct provide increased cross- sectional flow areas at downstream locations.
[0014] In another embodiment according to the foregoing embodiment, the insert has rounded axial ends.
[0015] In another embodiment according to the foregoing embodiment, the air supply system includes a plurality of ducts. Each of the ducts have inlet ends at locations spaced by at least 90° about a cross-sectional center axis of the gas turbine engine.
[0016] In another embodiment according to the foregoing embodiment, the inlet ends are spaced by 180°.
[0017] In another embodiment according to the foregoing embodiment, a plurality of insert holders center the insert within the duct.
[0018] In another embodiment according to the foregoing embodiment, a diffuser is positioned downstream of the compressor. An opening in the housing supplies air to the inlet end. The diffuser has an outer shroud and an inner shroud with intermediate vanes. The outer shroud ends at a location upstream of a downstream end of the inner shroud at locations circumferentially aligned with the inlet end.
[0019] In another embodiment according to the foregoing embodiment, the outer shroud only ends at the upstream location at circumferential locations associated with the inlet end, but extends further downstream at other locations. [0020] In another embodiment according to the foregoing embodiment, a flow path extends through the opening, and into the inlet end of the duct which has at least a portion formed with a part-circular radius.
[0021] In another embodiment according to the foregoing embodiment, the opening in the housing leads into the inlet end of the duct, with an inlet end ending downstream of the opening in the housing, with the part-circular radius portion formed in the housing.
[0022] In another embodiment according to the foregoing embodiment, an opening is formed through the housing and into the inlet end of the duct wherein a flow path extends through the opening, and into the inlet end of the duct which has at least a portion formed with a part-circular radius.
[0023] In another embodiment according to the foregoing embodiment, the opening in the housing leads into the inlet end of the duct. The inlet end ends downstream of the opening in the housing, with the part-circular radius portion formed in the housing.
[0024] In another featured embodiment, a gas turbine engine has a compressor delivering air into a combustion section. The combustion section and compressor are housed in a housing. An air supply system communicates through the housing to deliver air from a location between an upstream end of the compressor, and an upstream end of the combustor. The air supply system has a duct with an inlet end extending to an outlet end. The duct is provided with a central insert at an upstream end. The central insert provides a venturi effect by reducing the cross-sectional flow between the insert and an inner wall of the duct at the upstream end. The insert and duct provide increased cross-sectional flow areas at downstream locations. The insert ends within the duct upstream of the outlet end. The air supply system includes a plurality of ducts, each being provided with an insert. The ducts inlet ends at locations spaced by at least 90° about a cross-sectional center axis of the gas turbine engine. A diffuser is positioned downstream of the compressor with an opening in the housing supplying air to the inlet end. The diffuser has an outer shroud and an inner shroud with intermediate vanes. The outer shroud ends at a location upstream of a downstream end of the inner shroud at locations circumferentially aligned with the inlet end. The outer shroud only ends at the upstream location at circumferential locations associated with the inlet end, but extends further downstream at other locations. A flow path extends through the opening, and into the inlet end of the duct which has at least a portion formed with a part-circular radius. [0025] In another featured embodiment, a gas turbine engine has a compressor delivering air into a combustion section. The combustion section and compressor are housed in a housing. An air supply system communicates through the housing to deliver air from a location between an upstream end of the compressor and an upstream end of the combustor. The air supply system has a plurality of ducts to move air towards a use. The ducts have inlet ends at locations spaced by at least 90° about a cross-sectional center axis of the gas turbine engine.
[0026] In another featured embodiment, a gas turbine engine has a compressor delivering air into a combustion section. The combustion section and compressor are housed in a housing. An air supply system communicates through the housing to deliver air from a location between an upstream end of the compressor, and an upstream end of the combustor. The air supply system has a duct with an inlet end. A diffuser is positioned downstream of the compressor, and has an outer shroud and an inner shroud with intermediate vanes. The outer shroud ends at a location upstream of the downstream end of the inner shroud at locations circumferentially aligned with the inlet end.
[0027] In another embodiment according to the foregoing embodiment, the outer shroud only ends at the upstream location at circumferential locations associated with the inlet ends, but extends further downstream at other locations.
BRIEF DESCRIPTION OF THE DRAWINGS
[0028] Figure 1 shows a prior art gas turbine engine.
[0029] Figure 2A shows a prior art air bleed arrangement.
[0030] Figure 2B shows a portion of the Figure 2A arrangement.
[0031] Figure 3 shows the location of one part of the prior art arrangement.
[0032] Figure 4 shows a tap location in the present application.
[0033] Figure 5 shows detail of a diffuser incorporated into this application.
[0034] Figure 6 A shows a first embodiment arrangement of ports.
[0035] Figure 6B shows a second embodiment arrangement of ports.
[0036] Figure 7A shows a feature of a venturi duct.
[0037] Figure 7B shows an alternative embodiment.
[0038] Figure 7C shows another feature of the Figure 7A and 7B embodiments.
[0039] Figure 8 shows yet another feature. DETAILED DESCRIPTION
[0040] Figure 1 shows a prior art gas turbine engine. A gas turbine engine 10, such as a turbofan gas turbine engine, circumferentially disposed about an engine centerline A, is shown in Figure 1. The engine 10 includes a fan 18, a compressor 12, a combustion section 14 and turbine sections 16. As is well known in the art, air compressed in the compressor 12 is mixed with fuel which is burned in the combustion section 14 and expanded across turbines 16. The turbines includes rotors that rotate in response to the expansion, driving compressor rotors and fan 18. This structure is shown somewhat schematically in Figure 1. While one example gas turbine engine is illustrated, it should be understood this invention extends to any other type gas turbine engine for any application.
[0041] Figure 2A shows a bleed air arrangement 30. As shown, there is a downstream tap port 34 which is associated with a location downstream of the compressor section 12, but upstream of the combustor 14 (see Figure 3). At low thrust application, air is tapped through port 34, and delivered to an outlet 32, which communicates with various uses 200 for air on the aircraft. As thrust increases, valves 142 shut off the flow from the port 34, and switch to a port 36 which is upstream of the port 34, and at an intermediate location in the compressor section 12.
[0042] Figure 2B shows that there are actually two of the ports 34 in a typical prior art arrangement which are spaced circumferentially about a central line of the engine. In the prior art, the ports 34 are spaced by a relatively small angle, and typically on the order of 25° to 30°.
[0043] Figure 3 shows the location of the port 34, associated with a location 40 downstream of a diffuser 44, which is the downstream end of compressor section 12. As shown, an opening 42 in a housing 41 communicates air into the duct port. Downstream of the area 40 is the combustor 14.
[0044] Figure 4 shows a diffuser embodiment 118 wherein an outer shroud 58 of the diffuser ends upstream of a radially inner shroud 54. As shown, vanes 56 extend between the shroud walls 54 and 58.
[0045] The duct 50 communicates with an opening 42. Further details of this duct will be disclosed below.
[0046] Figure 5 shows a portion of the diffuser 118. As shown, the outer shroud is cut upstream, at locations 58 associated with the opening 42, but otherwise extends forwardly 158 to the location of the prior art diffuser as shown, for example, in Figure 3. In embodiments, there are plural ducts 50 and openings 42, and the cutaway locations 58 are associated with each opening 42.
[0047] Cutting away the diffuser at the areas 58 associated with the opening 42 dramatically reduces pressure loss.
[0048] Figure 6A shows an arrangement wherein the ports 64 and 66 communicate with a common duct 62, and are spaced by approximately 180° about a centerline X.
[0049] Applicant has discovered that by increasing the distance between the ports, the pressure loss across the system is dramatically reduced.
[0050] Figure 6B shows an alternative arrangement wherein there are ports 68 and 72. The ports 68 and 72 are spaced by an angle A, which in this embodiment would be approximately 90°. In embodiments, it is preferred that the ports are spaced by at least 90° to minimize pressure loss.
[0051] As shown schematically at 300, there could be a third port incorporated at a smaller angle. While the use of ports spaced by at least 90° is a feature of this combination, it should be clear from the Figure 6B, that the other features of this application could be utilized in a system wherein the ports are spaced closer, such as shown by 300 in Figure 6B, or as shown in Figure 2B.
[0052] Figure 7A shows a duct 50 which includes an insert 82 at an upstream end 79. As shown, the insert 82 creates a venturi with a relatively small cross-sectional flow 80 at the upstream end, and increasing to a larger flow 86. Ends of the insert 84 and 184 are spherical to reduce pressure losses associated with air flowing along those surfaces.
[0053] A feature provided by the insert, is that flow separation will be prevented since the flow would be through an annular area between the insert and the inner wall of the duct. This and the venturi effect result in the reduced pressure losses.
[0054] Figure 7B shows the use of the insert 182 in a somewhat alternative duct embodiment 150. As can be seen from Figure 7 A, the insert and duct in the Figure 7 A embodiment bend at 88, while the Figure 7B embodiment extends generally linearly.
[0055] As shown in Figure 7B, an upstream flow cross-sectional area Ai defined between the outer periphery of the insert 182 and the inner periphery of the duct 150 is much smaller than a downstream cross-sectional flow area A2 again defined between the insert and the inner wall of the duct. This creates a venturi effect. [0056] As can be seen from the Figures 7 A and 7B, the insert ends at an intermediate location within the ducts 50 or 150, and will end before the outlet 400 of the duct 50, which communicates into other portions of the air supply system.
[0057] Figure 7C shows features of the Figure 7A or 7B embodiments wherein insert holders 90 mount the insert within the duct 50. In this embodiment there are three insert holders, each spaced by 120°.
[0058] Figure 8 shows another feature wherein duct 50 ends at an upstream end 250. The opening 142 in the housing 143 has a lead in radius R] communicating into the duct 50. The radius Ri may be .25" (0.6 cm) in one embodiment. Of course, other radii may be used. By having a part circular radius leading into the duct, the pressure losses are also reduced.
[0059] The combination of features dramatically reduces pressure loss, and provides a more efficient system for delivering bleed air.
[0060] Although an embodiment of this invention has been disclosed, a worker of ordinary skill in this art would recognize that certain modifications would come within the scope of this invention. For that reason, the following claims should be studied to determine the true scope and content of this invention.

Claims

1. A bleed air supply system for a gas turbine engine comprising:
a duct having an inlet end and extending to an outlet end, an inlet end of said duct being provided with a central insert at an upstream end; and
said insert ending within said duct upstream of said outlet end.
2. The system as set forth in claim 1, wherein, said central insert providing a venturi effect by reducing the cross-sectional flow area between the insert and an inner wall of the duct of the upstream end, and the insert and duct providing increased cross-sectional flow areas at downstream locations.
3. The system as set forth in claim 1, wherein said insert has rounded axial ends.
4. The system as set forth in claim 1, wherein said air supply system includes a plurality of ducts, each of said ducts being provided with an insert, and said ducts having inlet ends at locations spaced by at least 90° about a cross-sectional center axis of a gas turbine engine which is to receive said air supply system.
5. The system as set forth in claim 4, wherein said inlet ends are spaced by 180°.
6. The system as set forth in claim 1, wherein a plurality of insert holders center said insert within said duct.
7. A gas turbine engine comprising:
a compressor delivering air into a combustion section, and said combustion section and said compressor being housed in a housing, an air supply system communicating through said housing to deliver air from a location between an upstream end of said compressor, and an upstream end of said combustor; and
said air supply system having a duct with an inlet end and extending to an outlet end, said duct being provided with a central insert at an upstream end, said insert ending within said duct upstream of said outlet end.
8. The engine as set forth in claim 7, and said central insert providing a venturi effect by reducing the cross-sectional flow between the insert and an inner wall of the duct at the upstream end, and said insert and duct providing increased cross-sectional flow areas at downstream locations.
9. The engine as set forth in claim 7, wherein said insert has rounded axial ends.
10. The engine as set forth in claim 7, wherein said air supply system includes a plurality of ducts, each of said ducts having inlet ends at locations spaced by at least 90° about a cross- sectional center axis of the gas turbine engine.
11. The engine as set forth in claim 10, wherein said inlet ends are spaced by 180°.
12. The engine as set forth in claim 7, wherein a plurality of insert holders center said insert within said duct.
13. The engine as set forth in claim 7, wherein a diffuser is positioned downstream of said compressor, an opening in said housing supplying air to said inlet end, and said diffuser having an outer shroud and an inner shroud with intermediate vanes, and said outer shroud ending at a location upstream of a downstream end of said inner shroud at locations circumferentially aligned with said inlet end.
14. The engine as set forth in claim 13, wherein said outer shroud only ends at the upstream location at circumferential locations associated with said inlet end, but extends further downstream at other locations.
15. The engine as set forth in claim 13, wherein a flow path through said opening, and into said inlet end of said duct has at least a portion formed with a part-circular radius.
16. The engine as set forth in claim 15, wherein said opening in said housing leads into said inlet end of said duct, with said inlet end ending downstream of said opening in said housing, with said part-circular radius portion formed in said housing.
17. The engine as set forth in claim 7, wherein an opening is formed through said housing and into said inlet end of said duct wherein a flow path through said opening, and into said inlet end of said duct has at least a portion formed with a part-circular radius.
18. The engine as set forth in claim 17, wherein said opening in said housing leads into said inlet end of said duct, with said inlet end ending downstream of said opening in said housing, with said part-circular radius portion formed in said housing.
19. A gas turbine engine comprising:
a compressor delivering air into a combustion section, and said combustion section and said compressor being housed in a housing, an air supply system communicating through said housing to deliver air from a location between an upstream end of said compressor, and an upstream end of said combustor;
said air supply system having a duct with an inlet end and extending to an outlet end, said duct being provided with a central insert at an upstream end, said central insert providing a venturi effect by reducing the cross-sectional flow between the insert and an inner wall of the duct at the upstream end, and said insert and duct providing increased cross-sectional flow areas at downstream locations, said insert ending within said duct upstream of said outlet end;
said air supply system includes a plurality of said ducts, each of said ducts being provided with an insert, and said ducts inlet ends at locations spaced by at least 90° about a cross-sectional center axis of the gas turbine engine;
a diffuser positioned downstream of said compressor, an opening in said housing supplying air to said inlet end, and said diffuser having an outer shroud and an inner shroud with intermediate vanes, and said outer shroud ending at a location upstream of a downstream end of said inner shroud at locations circumferentially aligned with said inlet end, said outer shroud only ends at the upstream location at circumferential locations associated with said inlet end, but extends further downstream at other locations; and
a flow path through said opening, and into said inlet end of said duct has at least a portion formed with a part-circular radius.
20. A gas turbine engine comprising:
a compressor delivering air into a combustion section, and said combustion section and said compressor being housed in a housing, an air supply system communicating through said housing to deliver air from a location between an upstream end of said compressor and an upstream end of said combustor;
said air supply system having a plurality of ducts to move air towards a use; and said ducts having inlet ends at locations spaced by at least 90° about a cross-sectional center axis of the gas turbine engine.
21. A gas turbine engine comprising:
a compressor delivering air into a combustion section, and said combustion section and said compressor being housed in a housing, an air supply system communicating through said housing to deliver air from a location between an upstream end of said compressor, and an upstream end of said combustor;
said air supply system having a duct with an inlet end; and
a diffuser positioned downstream of said compressor, and said diffuser having an outer shroud and an inner shroud with intermediate vanes, and said outer shroud ending at a location upstream of the downstream end of said inner shroud at locations circumferentially aligned with said inlet end.
22. The gas turbine engine as set forth in claim 21 , wherein said outer shroud only ends at the upstream location at circumferential locations associated with said inlet ends, but extends further downstream at other locations.
PCT/US2013/024697 2012-02-06 2013-02-05 Customer bleed air pressure loss reduction WO2013119520A1 (en)

Priority Applications (2)

Application Number Priority Date Filing Date Title
EP13747274.2A EP2812548B1 (en) 2012-02-06 2013-02-05 Gas turbine engine
EP20162505.0A EP3690209A1 (en) 2012-02-06 2013-02-05 Customer bleed air pressure loss reduction

Applications Claiming Priority (2)

Application Number Priority Date Filing Date Title
US13/366,597 US10119468B2 (en) 2012-02-06 2012-02-06 Customer bleed air pressure loss reduction
US13/366,597 2012-02-06

Publications (1)

Publication Number Publication Date
WO2013119520A1 true WO2013119520A1 (en) 2013-08-15

Family

ID=48901708

Family Applications (1)

Application Number Title Priority Date Filing Date
PCT/US2013/024697 WO2013119520A1 (en) 2012-02-06 2013-02-05 Customer bleed air pressure loss reduction

Country Status (3)

Country Link
US (1) US10119468B2 (en)
EP (2) EP2812548B1 (en)
WO (1) WO2013119520A1 (en)

Families Citing this family (2)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US10100675B2 (en) * 2014-12-09 2018-10-16 United Technologies Corporation Outer diffuser case for a gas turbine engine
US10024538B2 (en) * 2015-08-26 2018-07-17 United Technologies Corporation Apparatus and method for air extraction at a gas turbine engine combustor

Citations (9)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US4711084A (en) * 1981-11-05 1987-12-08 Avco Corporation Ejector assisted compressor bleed
US4870826A (en) 1987-06-18 1989-10-03 Societe Nationale D'etude Et De Construction De Moteurs D'aviation (Snecma) Casing for a turbojet engine combustion chamber
US4979587A (en) 1989-08-01 1990-12-25 The Boeing Company Jet engine noise suppressor
US5287694A (en) 1992-10-05 1994-02-22 General Electric Company Fluid channeling system
US5575144A (en) 1994-11-28 1996-11-19 General Electric Company System and method for actively controlling pressure pulses in a gas turbine engine combustor
US20090000307A1 (en) * 2007-06-27 2009-01-01 Toyota Jidosha Kabushiki Kaisha Air-bleed gas turbine
US20100013242A1 (en) * 2007-05-18 2010-01-21 United Technologies Corporation Aircraft combination engines plural airflow conveyances system
US20110138819A1 (en) 2009-12-15 2011-06-16 Kawasaki Jukogyo Kabushiki Kaisha Gas turbine engine with cooling arrangement
US20110192170A1 (en) * 2010-02-11 2011-08-11 Kevin Allan Dooley Air contamination detection in an aircraft air system

Family Cites Families (14)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US3318097A (en) * 1964-10-12 1967-05-09 Garrett Corp Air mass flow multiplier for thrust augmentation
US3441045A (en) 1966-12-02 1969-04-29 Boeing Co Variable orifice nozzle mixing ejector
US4001356A (en) 1975-08-22 1977-01-04 Clinton Graybill Variable venturi downdraft carburetor
US4765131A (en) 1985-02-08 1988-08-23 Allied Signal Inc. Aircraft engine bleed air flow balancing technique
US5155991A (en) 1991-04-01 1992-10-20 United Technologies Corporation Bleed air flow regulators with flow balance
US5211003A (en) * 1992-02-05 1993-05-18 General Electric Company Diffuser clean air bleed assembly
US5351478A (en) * 1992-05-29 1994-10-04 General Electric Company Compressor casing assembly
US5581996A (en) * 1995-08-16 1996-12-10 General Electric Company Method and apparatus for turbine cooling
US6327844B1 (en) * 2000-03-03 2001-12-11 General Electric Company Methods and apparatus for retaining flow restrictors within turbine engines
US6325595B1 (en) 2000-03-24 2001-12-04 General Electric Company High recovery multi-use bleed
US7197881B2 (en) 2004-03-25 2007-04-03 Honeywell International, Inc. Low loss flow limited feed duct
WO2006091138A1 (en) * 2005-02-25 2006-08-31 Volvo Aero Corporation A bleed structure for a bleed passage in a gas turbine engine
US20100031627A1 (en) * 2008-08-07 2010-02-11 United Technologies Corp. Heater Assemblies, Gas Turbine Engine Systems Involving Such Heater Assemblies and Methods for Manufacturing Such Heater Assemblies
GB0820598D0 (en) * 2008-11-11 2008-12-17 Rolls Royce Plc A noise reduction device

Patent Citations (9)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US4711084A (en) * 1981-11-05 1987-12-08 Avco Corporation Ejector assisted compressor bleed
US4870826A (en) 1987-06-18 1989-10-03 Societe Nationale D'etude Et De Construction De Moteurs D'aviation (Snecma) Casing for a turbojet engine combustion chamber
US4979587A (en) 1989-08-01 1990-12-25 The Boeing Company Jet engine noise suppressor
US5287694A (en) 1992-10-05 1994-02-22 General Electric Company Fluid channeling system
US5575144A (en) 1994-11-28 1996-11-19 General Electric Company System and method for actively controlling pressure pulses in a gas turbine engine combustor
US20100013242A1 (en) * 2007-05-18 2010-01-21 United Technologies Corporation Aircraft combination engines plural airflow conveyances system
US20090000307A1 (en) * 2007-06-27 2009-01-01 Toyota Jidosha Kabushiki Kaisha Air-bleed gas turbine
US20110138819A1 (en) 2009-12-15 2011-06-16 Kawasaki Jukogyo Kabushiki Kaisha Gas turbine engine with cooling arrangement
US20110192170A1 (en) * 2010-02-11 2011-08-11 Kevin Allan Dooley Air contamination detection in an aircraft air system

Non-Patent Citations (1)

* Cited by examiner, † Cited by third party
Title
See also references of EP2812548A4 *

Also Published As

Publication number Publication date
EP2812548A4 (en) 2015-10-21
US20130199205A1 (en) 2013-08-08
EP2812548B1 (en) 2020-04-15
EP2812548A1 (en) 2014-12-17
EP3690209A1 (en) 2020-08-05
US10119468B2 (en) 2018-11-06

Similar Documents

Publication Publication Date Title
US10113486B2 (en) Method and system for modulated turbine cooling
CN107246411B (en) High pressure exhaust muffler device with multiple sources
US8082726B2 (en) Tangential anti-swirl air supply
US7712317B2 (en) Flow control systems
EP2943667B1 (en) Two spool gas generator with improved air porting
US10907503B2 (en) Compressors with integrated secondary air flow systems
US8657567B2 (en) Nacelle compartment plenum for bleed air flow delivery system
EP1728992B1 (en) Bleed valve for a gas turbine engine and a corresponding gas turbine.
EP1923575A2 (en) Servo-controlled variable geometry ejector pump
EP1923553A2 (en) Multi-port bleed system with variable geometry ejector pump
GB2342693A (en) Pressure boosted compressor cooling system.
EP3165451B1 (en) Nose cone heat exchanger with radially outer discharge
US10119468B2 (en) Customer bleed air pressure loss reduction
EP1795762A2 (en) Flow redirector for compressor inlet
US7197881B2 (en) Low loss flow limited feed duct
US20200102886A1 (en) Customer bleed air pressure loss reduction
CN115013093A (en) Diffuser discharge assembly
EP2806139B1 (en) Gas turbine engines having plasma flow-controlled intake systems
US11781504B2 (en) Bleed plenum for compressor section
EP4411229A1 (en) Combustor for gas turbine engine with central fuel injection ports
US20180238234A1 (en) Fuel injection device for a gas turbine

Legal Events

Date Code Title Description
121 Ep: the epo has been informed by wipo that ep was designated in this application

Ref document number: 13747274

Country of ref document: EP

Kind code of ref document: A1

NENP Non-entry into the national phase

Ref country code: DE

WWE Wipo information: entry into national phase

Ref document number: 2013747274

Country of ref document: EP