US3365124A - Compressor structure - Google Patents
Compressor structure Download PDFInfo
- Publication number
- US3365124A US3365124A US528860A US52886066A US3365124A US 3365124 A US3365124 A US 3365124A US 528860 A US528860 A US 528860A US 52886066 A US52886066 A US 52886066A US 3365124 A US3365124 A US 3365124A
- Authority
- US
- United States
- Prior art keywords
- air
- slot
- airfoil
- vanes
- vane
- 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
Links
Images
Classifications
-
- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F04—POSITIVE - DISPLACEMENT MACHINES FOR LIQUIDS; PUMPS FOR LIQUIDS OR ELASTIC FLUIDS
- F04D—NON-POSITIVE-DISPLACEMENT PUMPS
- F04D29/00—Details, component parts, or accessories
- F04D29/40—Casings; Connections of working fluid
- F04D29/52—Casings; Connections of working fluid for axial pumps
- F04D29/54—Fluid-guiding means, e.g. diffusers
- F04D29/541—Specially adapted for elastic fluid pumps
- F04D29/542—Bladed diffusers
-
- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F04—POSITIVE - DISPLACEMENT MACHINES FOR LIQUIDS; PUMPS FOR LIQUIDS OR ELASTIC FLUIDS
- F04D—NON-POSITIVE-DISPLACEMENT PUMPS
- F04D27/00—Control, e.g. regulation, of pumps, pumping installations or pumping systems specially adapted for elastic fluids
- F04D27/02—Surge control
- F04D27/0207—Surge control by bleeding, bypassing or recycling fluids
- F04D27/023—Details or means for fluid extraction
-
- 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
- F05D2240/00—Components
- F05D2240/10—Stators
- F05D2240/12—Fluid guiding means, e.g. vanes
- F05D2240/125—Fluid guiding means, e.g. vanes related to the tip of a stator vane
-
- 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
Definitions
- An axial flow compressor includes a row of stator vanes having openings for the extraction of bleed air at their inner tips such that clean air in the inner portion of the flow passageway is extracted.
- the present invention relates to compressor structure and, more particularly, to interstage bleed means as applied to a multi-stage axial flow compressor.
- the main object of the present invention is to provide an interstage bleed structure which operates at the rotor LD. with minimum aerodynamic disturbance.
- Another object is to provide a system which removes clean air with minimum unequal circumferential, axial, or radial pressure distribution in a compressor.
- a further object is to provide a system wherein the losses due to the presence of the removal structure are minimized while obtaining the maximum area for air removal.
- a further object is to provide such a system which employs no stator vane tip shrouding and all of the structure for the removal of the air is contained in the vane itself.
- the invention is for use in a multi-stage axial ow compressor which includes a rotor hub having airfoil blades or vanes extending radially therefrom and a casing with plural stages of stator airfoil vanes between the rotor blades whereby the casing and hub define a iiow passage in the conventional manner.
- At least one of the stator vane stages and the vane platforms are hollow.
- an interstage bleed means is provided comprising an elongated slot in at least some of the vanes of the hollow stage.
- the slot is provided in all of the vanes of the stage.
- the slot is cut back from the tip of the pressure surface and it extends parallel to the hub surface to scoop or scrape the air at the hub into the vane through its concave or high pressure side.
- the air iiows through the supported hollow vane and through the hollow shank or platform into a collecting manifold that is provided on the outside of the casing bridging the hollow portion of the vanes.
- the collecting manifold receives the air and directs it to a point of use.
- the slots are rectangular and disposed between the airfoil leading and trailing edges and set back from each edge.
- the slot is preferably maintained at a minimum height-about equal to the set-back-and the c-ross sectional area of the slot is at least equal to the minimum airfoil cross sectional area for maximum effectiveness.
- FIGURE l is a partial cross sectional View of a typical interstage bleed means of the present invention.
- FIGURE 2 is a partial perspective showing the slot adjacent the hub.
- FIGURE 3 is an enlarged cross section on the line 3 3 of FIGURE l.
- FIGURE l there is shown the central portion of a multi-stage axial iiow compressor that includes a rotor drum or hub 10 having rotating airfoil blades 12 thereon in the conventional fashion.
- a stator casing 14 is provided with the usual inwardly directed radial stator airfoil vanes 16 intermeshed between the rotating blades for raising the pressure of the air passing through fiow passage 18 defined by the casing 14 and hub surface 20 of rotor It).
- the removal means must be such as to minimize amy distortion or pressure unbalance of the airflow through the compressor. Additionally, it is desired to do this with a minimum amount of structure and to avoid vane tip shrouding if possible.
- stator vanes 16 and preferably a complete stage of vanes 16 are hollow to provide an internal passage 22 extending through the vane airfoil and shank portion 24 outside of the casing.
- the structure for doing this consists of an opening, preferably in the form of a rectangular slot 26, that is formed in the high pressure or concave surface 28 as seen in FIGURES 2 and 3. Aerodynamically, this is the non-critical surface of the airfoil as opposed to the low pressure or suction surface 30 on the opposite convex side.
- the slot is preferably cut back from the tip of the pressure surface as shown and is disposed to extend parallel to the hub surface 20, to, in effect, scrape or scoop off a supply of pressurized air.
- the scooped air under pressure then flows radially outward through the vane and shank portion 24 to the outer side of the casing 14.
- a manifold 32 on the casing bridging the hollow platforms or Shanks of the hollow vanes as shown in FIGURE l.
- the manifold may direct the air in any suitable manner to a distant point of use.
- the cross sectional area of the slot must be at least equal to the minimum airfoil cross sectional area if it is not to be a limiting opening.
- the slot area is greater than the airfoil cross section it will be obvious that it will scoop up more air than can be used because the cross sectional area of the airfol is a limiting dimension. If the slot area is less than the airfoil can handle, then the maximum bleed elfe-ct is not obtained.
- the cross sectional area of slot 26 may assume any rectangular shape, It could be tall and narrow and it can be wide and low in height as shown for the same area. It is desired that the slot be low and wide and extend chordwise on the pressure surface as shown to avoid adverse aerodynamic effects that would be present in a siot extending Well into the ilow stream and to obtain the cleanest possible air.
- the strength of the individual vane must be maintained since it is gen erally constructed of thin sheet metal as shown in FIG URE 3 and so it is supported internally by suitable supporting structure 34 which may also form the passages for the outward removal of air.
- the strength is maintained by providing the slot in a set-back from the leading and trailing edges as shown at 36 and 38 respectively in FIGURE 1.
- the slot may be varied depending on the air quantity desired and may take other forms than the rectangular one shown.
- the manifold 32 may extend over more than the hollow stage in order to obtain the necessary area for handling the removed air and still maintain minimum diameter.
- a multi-stage axial flow compressor including a rotor hub having airfoil blades thereon and a casing having plural stages of and including some hollow stator airfoil vanes between the rotor blades,
- said casing and hub defining a flow passage, bleed means comprising,
- said opening means extending in a generally axial direction parallel to the hub surface with its axial extent being substantially greater than its radial extent
- a collecting manifold on said casing bridging the hollow vane base for receiving said air and directing it to a point of use.
- a multi-stage axial ow compressor including a rotor hub having airfoil blades thereon and a casing having plural stages of stator airfoil vanes between the rotor blades, at least one of said vane stages and platforms therefor being hollow, said casing and hub dening a ilow passage, interstage bleed means comprising,
- a collecting manifold around said casing bridging the hollow platforms of said hollow stage for receiving said air and directing it to a point of use.
Landscapes
- Engineering & Computer Science (AREA)
- Mechanical Engineering (AREA)
- General Engineering & Computer Science (AREA)
- Structures Of Non-Positive Displacement Pumps (AREA)
Priority Applications (8)
Application Number | Priority Date | Filing Date | Title |
---|---|---|---|
US528860A US3365124A (en) | 1966-02-21 | 1966-02-21 | Compressor structure |
BE689754D BE689754A (ja) | 1966-02-21 | 1966-11-16 | |
GB51289/66A GB1119479A (en) | 1966-02-21 | 1966-11-16 | Improvements in multi-stage axial flow compressors |
CH1659166A CH452098A (de) | 1966-02-21 | 1966-11-17 | Mehrstufiger Axialverdichter |
NL6616250A NL6616250A (ja) | 1966-02-21 | 1966-11-18 | |
DE19661628261 DE1628261B2 (de) | 1966-02-21 | 1966-11-18 | Luftentnahmevorrichtung an axialverdichtern |
SE15906/66A SE308359B (ja) | 1966-02-21 | 1966-11-21 | |
FR84320A FR1501382A (fr) | 1966-02-21 | 1966-11-21 | Structure de compresseur |
Applications Claiming Priority (1)
Application Number | Priority Date | Filing Date | Title |
---|---|---|---|
US528860A US3365124A (en) | 1966-02-21 | 1966-02-21 | Compressor structure |
Publications (1)
Publication Number | Publication Date |
---|---|
US3365124A true US3365124A (en) | 1968-01-23 |
Family
ID=24107490
Family Applications (1)
Application Number | Title | Priority Date | Filing Date |
---|---|---|---|
US528860A Expired - Lifetime US3365124A (en) | 1966-02-21 | 1966-02-21 | Compressor structure |
Country Status (8)
Country | Link |
---|---|
US (1) | US3365124A (ja) |
BE (1) | BE689754A (ja) |
CH (1) | CH452098A (ja) |
DE (1) | DE1628261B2 (ja) |
FR (1) | FR1501382A (ja) |
GB (1) | GB1119479A (ja) |
NL (1) | NL6616250A (ja) |
SE (1) | SE308359B (ja) |
Cited By (10)
Publication number | Priority date | Publication date | Assignee | Title |
---|---|---|---|---|
DE3504377A1 (de) * | 1984-02-13 | 1985-08-14 | General Electric Co., Schenectady, N.Y. | Profilkoerperkonstruktion und verfahren zur herstellung derselben |
FR2559422A1 (fr) * | 1984-02-13 | 1985-08-16 | Gen Electric | Element a profil de pale creux composite muni d'une structure support interne ondulee et son procede de fabrication |
US4639189A (en) * | 1984-02-27 | 1987-01-27 | Rockwell International Corporation | Hollow, thermally-conditioned, turbine stator nozzle |
US5586859A (en) * | 1995-05-31 | 1996-12-24 | United Technologies Corporation | Flow aligned plenum endwall treatment for compressor blades |
JP2006105134A (ja) * | 2004-09-30 | 2006-04-20 | Snecma | ターボ機械の圧縮機における空気循環方法、この方法を使用する圧縮機の装置、このような装置を組み込む圧縮段および圧縮機、並びにこのような圧縮機を装備した航空機エンジン |
US20060193719A1 (en) * | 2005-02-28 | 2006-08-31 | General Electric Company | Bolt-on radial bleed manifold |
WO2013165281A1 (en) * | 2012-05-02 | 2013-11-07 | Gkn Aerospace Sweden Ab | Supporting structure for a gas turbine engine |
US20180080476A1 (en) * | 2016-09-19 | 2018-03-22 | United Technologies Corporation | Geared turbofan front center body thermal management |
US20180195408A1 (en) * | 2017-01-12 | 2018-07-12 | General Electric Company | Method and system for ice tolerant bleed takeoff |
US20190301301A1 (en) * | 2018-04-02 | 2019-10-03 | General Electric Company | Cooling structure for a turbomachinery component |
Families Citing this family (3)
Publication number | Priority date | Publication date | Assignee | Title |
---|---|---|---|---|
US4182117A (en) * | 1978-01-09 | 1980-01-08 | Avco Corporation | Diffuser vane cusp bleed aperture with automatic ejector control device |
DE19814627C2 (de) * | 1998-04-01 | 2001-02-22 | Man Turbomasch Ag Ghh Borsig | Kühlluftentnahme aus dem Diffusorteil eines Kompressors einer Gasturbine |
DE10355240A1 (de) * | 2003-11-26 | 2005-07-07 | Rolls-Royce Deutschland Ltd & Co Kg | Strömungsarbeitsmaschine mit Fluidentnahme |
Citations (7)
Publication number | Priority date | Publication date | Assignee | Title |
---|---|---|---|---|
US2344835A (en) * | 1943-08-07 | 1944-03-21 | Edward A Stalker | Pump |
US2656146A (en) * | 1948-04-08 | 1953-10-20 | Curtiss Wright Corp | Turbine blade construction |
US2682363A (en) * | 1950-12-08 | 1954-06-29 | Rolls Royce | Gas turbine engine |
US2720356A (en) * | 1952-06-12 | 1955-10-11 | John R Erwin | Continuous boundary layer control in compressors |
US2848155A (en) * | 1950-11-22 | 1958-08-19 | United Aircraft Corp | Boundary layer control apparatus for compressors |
US2859934A (en) * | 1953-07-29 | 1958-11-11 | Havilland Engine Co Ltd | Gas turbines |
US2924425A (en) * | 1953-02-02 | 1960-02-09 | Bristol Aero Engines Ltd | Aerofoil-section bladed structures |
-
1966
- 1966-02-21 US US528860A patent/US3365124A/en not_active Expired - Lifetime
- 1966-11-16 BE BE689754D patent/BE689754A/xx unknown
- 1966-11-16 GB GB51289/66A patent/GB1119479A/en not_active Expired
- 1966-11-17 CH CH1659166A patent/CH452098A/de unknown
- 1966-11-18 DE DE19661628261 patent/DE1628261B2/de active Pending
- 1966-11-18 NL NL6616250A patent/NL6616250A/xx unknown
- 1966-11-21 SE SE15906/66A patent/SE308359B/xx unknown
- 1966-11-21 FR FR84320A patent/FR1501382A/fr not_active Expired
Patent Citations (7)
Publication number | Priority date | Publication date | Assignee | Title |
---|---|---|---|---|
US2344835A (en) * | 1943-08-07 | 1944-03-21 | Edward A Stalker | Pump |
US2656146A (en) * | 1948-04-08 | 1953-10-20 | Curtiss Wright Corp | Turbine blade construction |
US2848155A (en) * | 1950-11-22 | 1958-08-19 | United Aircraft Corp | Boundary layer control apparatus for compressors |
US2682363A (en) * | 1950-12-08 | 1954-06-29 | Rolls Royce | Gas turbine engine |
US2720356A (en) * | 1952-06-12 | 1955-10-11 | John R Erwin | Continuous boundary layer control in compressors |
US2924425A (en) * | 1953-02-02 | 1960-02-09 | Bristol Aero Engines Ltd | Aerofoil-section bladed structures |
US2859934A (en) * | 1953-07-29 | 1958-11-11 | Havilland Engine Co Ltd | Gas turbines |
Cited By (24)
Publication number | Priority date | Publication date | Assignee | Title |
---|---|---|---|---|
FR2559422A1 (fr) * | 1984-02-13 | 1985-08-16 | Gen Electric | Element a profil de pale creux composite muni d'une structure support interne ondulee et son procede de fabrication |
FR2559423A1 (fr) * | 1984-02-13 | 1985-08-16 | Gen Electric | Elements a profil de pale creux composites et leur procede de fabrication |
US4594761A (en) * | 1984-02-13 | 1986-06-17 | General Electric Company | Method of fabricating hollow composite airfoils |
DE3504377A1 (de) * | 1984-02-13 | 1985-08-14 | General Electric Co., Schenectady, N.Y. | Profilkoerperkonstruktion und verfahren zur herstellung derselben |
US4639189A (en) * | 1984-02-27 | 1987-01-27 | Rockwell International Corporation | Hollow, thermally-conditioned, turbine stator nozzle |
US5586859A (en) * | 1995-05-31 | 1996-12-24 | United Technologies Corporation | Flow aligned plenum endwall treatment for compressor blades |
US20060222485A1 (en) * | 2004-09-30 | 2006-10-05 | Snecma | Method for air circulation in a turbomachine compressor, compressor arrangement using this method, compression stage and compressor incorporating such a arrangement, and aircraft engine equipped with such a compressor |
JP2006105134A (ja) * | 2004-09-30 | 2006-04-20 | Snecma | ターボ機械の圧縮機における空気循環方法、この方法を使用する圧縮機の装置、このような装置を組み込む圧縮段および圧縮機、並びにこのような圧縮機を装備した航空機エンジン |
US7581920B2 (en) * | 2004-09-30 | 2009-09-01 | Snecma | Method for air circulation in a turbomachine compressor, compressor arrangement using this method, compression stage and compressor incorporating such a arrangement, and aircraft engine equipped with such a compressor |
CN101082345B (zh) * | 2005-02-28 | 2010-12-08 | 通用电气公司 | 螺栓固定的径向排气管 |
US7374396B2 (en) * | 2005-02-28 | 2008-05-20 | General Electric Company | Bolt-on radial bleed manifold |
JP2006242184A (ja) * | 2005-02-28 | 2006-09-14 | General Electric Co <Ge> | 抽気マニホルドおよびコンプレッサー・ケースアセンブリ |
US20060193719A1 (en) * | 2005-02-28 | 2006-08-31 | General Electric Company | Bolt-on radial bleed manifold |
EP2844880A4 (en) * | 2012-05-02 | 2016-04-20 | Gkn Aerospace Sweden Ab | SUPPORT STRUCTURE FOR A GAS TURBINE ENGINE |
JP2015516537A (ja) * | 2012-05-02 | 2015-06-11 | ゲーコーエヌ エアロスペース スウェーデン アーベー | ガスタービンエンジンの支持構造 |
US20150176494A1 (en) * | 2012-05-02 | 2015-06-25 | Gkn Aerospace Sweden Ab | Supporting structure for a gas turbine engine |
WO2013165281A1 (en) * | 2012-05-02 | 2013-11-07 | Gkn Aerospace Sweden Ab | Supporting structure for a gas turbine engine |
US9797312B2 (en) * | 2012-05-02 | 2017-10-24 | Gkn Aerospace Sweden Ab | Supporting structure for a gas turbine engine |
US20180080476A1 (en) * | 2016-09-19 | 2018-03-22 | United Technologies Corporation | Geared turbofan front center body thermal management |
US20180195408A1 (en) * | 2017-01-12 | 2018-07-12 | General Electric Company | Method and system for ice tolerant bleed takeoff |
EP3568575A4 (en) * | 2017-01-12 | 2020-08-12 | General Electric Company | FROST TOLERANT PURGE TAKE-OFF PROCESS AND SYSTEM |
US10968771B2 (en) * | 2017-01-12 | 2021-04-06 | General Electric Company | Method and system for ice tolerant bleed takeoff |
US20190301301A1 (en) * | 2018-04-02 | 2019-10-03 | General Electric Company | Cooling structure for a turbomachinery component |
US10808572B2 (en) * | 2018-04-02 | 2020-10-20 | General Electric Company | Cooling structure for a turbomachinery component |
Also Published As
Publication number | Publication date |
---|---|
NL6616250A (ja) | 1967-08-22 |
FR1501382A (fr) | 1967-11-10 |
BE689754A (ja) | 1967-05-16 |
GB1119479A (en) | 1968-07-10 |
SE308359B (ja) | 1969-02-10 |
DE1628261A1 (de) | 1970-12-10 |
CH452098A (de) | 1968-05-31 |
DE1628261B2 (de) | 1971-04-01 |
Similar Documents
Publication | Publication Date | Title |
---|---|---|
US2720356A (en) | Continuous boundary layer control in compressors | |
US8262340B2 (en) | Turbomachine exerting dynamic influence on the flow | |
US7200999B2 (en) | Arrangement for bleeding the boundary layer from an aircraft engine | |
US5503529A (en) | Turbine blade having angled ejection slot | |
US7364404B2 (en) | Turbomachine with fluid removal | |
CA2507972C (en) | Method and apparatus for assembling gas turbine engines | |
US6976826B2 (en) | Turbine blade dimple | |
US3365124A (en) | Compressor structure | |
US9726084B2 (en) | Compressor bleed self-recirculating system | |
US6358003B2 (en) | Rotor blade an axial-flow engine | |
US2801790A (en) | Compressor blading | |
US2660401A (en) | Turbine bucket | |
EP1022432A2 (en) | Cooled aerofoil for a gas turbine engine | |
JP2007536459A (ja) | 遷音速ガスタービンの衝撃波誘起境界層の抽気 | |
US2915279A (en) | Cooling of turbine blades | |
US3535873A (en) | Gas turbine cooling devices | |
EP0201770A2 (en) | Turbine engine with induced pre-swirl at the compressor inlet | |
US2846137A (en) | Construction for axial-flow turbomachinery | |
CN112983885B (zh) | 用于燃气涡轮发动机的风扇的分流器和转子翼型件的围带 | |
US3142438A (en) | Multi-stage axial compressor | |
GB742477A (en) | Improvements in or relating to bladed stator or rotor constructions for fluid machines such as axial-flow turbines or compressors | |
US20160319835A1 (en) | A gas turbine engine integrally bladed rotor with asymmetrical trench fillets | |
US11933193B2 (en) | Turbine engine with an airfoil having a set of dimples | |
US2848155A (en) | Boundary layer control apparatus for compressors | |
US20210372288A1 (en) | Compressor stator with leading edge fillet |