Search Images Maps Play YouTube News Gmail Drive More »
Sign in
Screen reader users: click this link for accessible mode. Accessible mode has the same essential features but works better with your reader.

Patents

  1. Advanced Patent Search
Publication numberUS2735612 A
Publication typeGrant
Publication dateFeb 21, 1956
Filing dateApr 20, 1950
Publication numberUS 2735612 A, US 2735612A, US-A-2735612, US2735612 A, US2735612A
InventorsGeorge F. Hausmann
Export CitationBiBTeX, EndNote, RefMan
External Links: USPTO, USPTO Assignment, Espacenet
hausmann
US 2735612 A
Abstract  available in
Images(2)
Previous page
Next page
Claims  available in
Description  (OCR text may contain errors)

Feb. 21, 1956 G ANN 2,735,612

BLADE PASSAGE CONSTRUCTION FOR I COMPRESSORS AND DIFFUSERS Filed April 20, 1950 2 Sheets-Sheet l FICBI FIC5.2

INVENTOR GEORGE F. HALJSMANN AGENT Feb. 21, 1956 G, F. HAUSMANN 2,735,612

BLADE PASSAGE CONSTRUCTION FOR COMPRESSORS AND DIFFUSERS Filed April 20, 1950 2 Sheets-Sheet 2 INVENTOR GEORGE F. l-IALJSMANN BY WKZJM AGENT BLADE PASSAGE CONSTRUCTION FOR COMPRESSORS AND D'IFFUSERS George F. Hausmann, Hartford, Comm, assignor to United Aircraft Corporation, East Hartford, Conn, a corporation of Delaware Application April 20, 1950, Serial No. 157,133

2 Claims. (Cl. 230122) This invention relates to improvements in confined fluid flow and more specifically to improved passage configurations for interblade passages in compressors and difiusers as for example in the air inlet diffusers for high performance aircraft.

It is an object of this invention to provide improved means for redistributing or energizing boundary layer flow along the confining wall of a fluid passage particularly in the immediate vicinity of rapid expansion regions such as is experienced between adjacent stator vanes of compressors, rotor blades or diffuser type turning vanes in air inlet passages.

Another object of this invention is to provide an improved boundary layer energizing mechanism of the type described herein comprising a substantially streamlined protrusion extending from the confining surface and located adjacent the blade extremities for varying the local pressure gradient by means of varying the flow passage along the axis of flow in order to obtain the particular aerodynamic advantages described hereinafter.

A further object of this invention is to provide protrusions of the type described to delay separation over the blades thereby obtaining maximum efficiency of diffusion and energy conversion within a minimum distance along the axis of flow. Therefore, a feature of this invention resides in improving the flow efficiencies over higher ranges of Mach numbers and lift coefficients of adjacent diffuser blades having a cascade arrangement.

These and other objects of this invention will become readily apparent from the following detail description of the drawings in which:

Figs. 1 and 1A illustrate flow separation over an airfoil shaped blade as caused by adverse effects of the boundary layer along an adjacent confining surface.

Figs. 2 and 3 illustrate the flow improving protrusions according to this invention.

Fig. 4 is a partial view of an aircraft fuselage illustrating a flush air inlet having diffuser type turning vanes for providing a high diffusion rate within a minimum of axial length.

Fig. 5 is a side view taken along the line 55 of Fig. 4.

Fig. 6 is an enlarged detail view of a portion of Fig. 4.

Fig. 7 is a detail view taken along the line 77 of Fig. 6.

Figs. 8 through 11 illustrate various modifications of this invention as applied to axial flow compressors.

in confined fluid flow or in fluid passages wherein a cascade of airfoil shaped vanes are arranged, for example so as to form diffuser passages therebetween, it is desirable to obtain a high rate of diffusion within a short distance along the axis of flow while also insuring maximum'efficiency. In confined fluid flow where the blades ends are in substantially juxtaposed relationship with the confining surface, it has been found that the boundary layer along the confining surface sets up secatent O 2,735,612 Patented Feb. 21, 1956 ondary flows particularly in the diffuser passages between the blades so that fluid separation obtains and the aerodynamic efficiency of the blades is not at an optimum value.

By way of example, and referring to Fig. l, a blade 20 is shown extending from a wall 22 of a flow confining surface. The blade 20 may be one of a group of blades spaced transversely of the axis of flow so as to provide diffuser passages therebetween, as for example the statorblades of a compressor. As seen herein, the boundary layer along the confining surface having a lower velocity than that of the main fluid stream sets up secondary flows particularly in the vicinity of the blade 26 where rapid expansion of theair may be taking place so that an adverse pressure grading and subsequent fluid separation over the cambered surface 26 may result. Hence, although the blade itself will normally have a boundary layer over its major surfaces which may under certain conditions cause fluid separation, the interaction of the boundary layer of the confining surface therewith accelerates these poor flow conditions and effects the flow a substantial distance away from the confining surface over the span of a particular blade or blades as illustrated by the arrows in Figs. 1 and 1A. Therefore, it is apparent that the range of blade lift coefficients at which high efiiciency can be maintained is substantially reduced over the ideal flow condition.

To this end then and in order to increase the ranges of efficient operation, a protrusion 30 is provided which extends into the fluid stream from the confining surface adjacent the blade 24. These protrusions may normally have a span transversely of the axis of How such as to extend from one blade to the other, as for example illustrated in Fig. 6. The protrusion 30 is of substantially streamline shape and has its leading edge 34 located within the first third of the chordwise dimension of the blade 24 measured from the leading edge of the blade. The maximum point of protrusion 36 is preferably located within the last fourth of the chordwise dimension of the blade 24 measured from the leading edge thereof. A downstream or trailing portion 38 of the protrusion 36 may terminate adjacent the trailing edge of the blade 24 as illustrated in Fig. 2, or it may assume the shape as illustrated in the pr0- trusion 40 in Fig. 3.

It has been found that extending the trailing edge of tie protrusion in the manner illustrated in Fig. 3 does not provide a great improvement over a trailing edge of the type illustrated in Fig. 2. The reason for the slight difference is attributed to the fact that immediately aft of the blade 24 where diffusion between the blades has caused an increase in pressure and a decrease in velocity, an abrupt contour like the trailing edge 38 of Fig. 2 is of little consequence.

It is then apparent that the protrusion 3th in efiect produces a gradual convergence of the confining fluid surface in the vicinity of the blade 24 so as to acceleratethe boundary layer air in the critical region Where it might otherwise cause secondary flows which are adverse to that of the main stream so as to cause fluid separation over the adjacent blade. This increase in boundary layer velocity reaches its peak near the trailing ed e of the blade 24 thereby delaying separation from the blade in this vicinity to minimize the dilatorious effect which would obtain over a large portion of the span of the blade.

The utilization of a protuberance on the hub wall and/or outer casing wall adjacent the outer end of the vanes has particular effect on the flow conditions where boundary layer flow along the duct wall and the flow conditions over the vanes interact to cause unsatisfactory conditions leading to separation and high drag. The maximum point of protuberance of the wall contour is located as shown downstream of the maximum thickness location of the vane. Location of the protuberance in this manner provides a local acceleration of the flow over the airfoil shaped vane in the vicinity of the vane-wall intersection. Considering the flow over the vane then, a normal decrease in velocity is experienced aft of the maxi mum thickness of the vanes due to the diifusion taking place. This condition along with the rise in pressure tends to prematurely retard flow of the boundary flow along the duct wall causing local separation along the wall and also over a large section of the adjacent vane.

From the foregoing it will be evident that the leading edge of the protrusion as shown is located downstream of the leading edge of the vanes since local acceleration of the boundary layer along the duct wall is desirable at approximately the point where diffusion commences between the major vane surfaces.

The advantages of the use of protrusions 30 makes them particularly adaptable for air inlets of high performance aircraft for it is desirable to reduce the velocity and increase the pressure of the air within a very short distance.

Referring to Fig. 4, a fuselage 50 is shown having imbedded therein one or more turbo-jet power plants 52 which induct air via a passage 54. A substantially flush air inlet 56 is provided in the fuselage 50 and includes a plurality of vanes 58 of airfoil shape which are arranged in a cascade so as to form a plurality of diffuser type diverging passages 60 therebetween (Fig. 6). The blades 58 extend completely across the air inlet passage and are of such configuration with regard to camber and angle of incidence so as to provide a high rate of expansion therebetween within a relatively short distance along the axis of flow. The expansion, of course, causes a transfer of velocity energy into pressure energy. In order to obtain this high rate energy conversion the blades must be highly aerodynamically loaded and operate efficiency over a wide range of Mach numbers.

Hence, in order to improve the efiiciency of the diffusion and to improve pressure recovery at the engine, a plurality of protrusions 30 are provided on the confining wall of the air inlet passage 56 adjacent the extremities of the blades 58 with the protrusions extending transversely of the axis of fiuid flow so as to span the spaces between the blades 58 as seen in Fig. 7. It will be noted that the protrusions 30 will then progressively vary the area of the fluid passages 60 and in effect further provide diffusion at their trailing edges in a plane parallel to the span of the blades 58. The passage 54 may gradually diverge along the axis of flow so as to constitute a diffuser in itself to further increase the pressure of the fluid after it has moved past the vanes 53.

The use of protrusions in the manner described is also readily adaptable to axial flow compressors, as shown for example by the modifications illustrated in Figs. 8 to 11. The protrusions in these modifications assume a shape similar to the protrusions 30 described above. Referring to Fig. 8, for example, the first stage compressor blades 80 and the second stage compressor blades 82 may have their rotor rims 84 formed with protrusions 86 while the stator blades 83 may have their adjacent confining walls 90 and 91 provided with protrusions 92. Fig. 9 illustrates a similar arrangement with the additional feature that a protrusion 96 is provided in the outer wall of the annular compressor passage 98 while the compressor blade 100 has its tip extremity 102 indented so as to complement the configuration of the projection 96.

Figs. and 11 illustrate further modifications of the configurations shown in Figs. 8 and 9. Thus in Fig. 10 the stator vane 110 has protrusions 112, 114 adjacent the extremities there while the confining Walls in the vicinity 4 of the rotor blades 116, 118 are devoid of any protrusions. However, the rotor rim 120 and the adjacent confining wall portion 122 are of larger diameter than the upstream confining surface so that the annular passage 124 is of lesser cross-sectional area than the upstream portion of the passage as is conventional in multi-stage compressors.

In Fig. 11, on the other hand, the extremities of both the rotor blade and the stator blade 132 have inwardly directed protrusions adjacent thereto While the confining walls 134, 136 gradually converge to diminish the crosssectional area of the passage 140 in a downstream direction.

The amount that the protrusion extends into the main stream is preferably determined by test so that it will produce the maximum flow improvement consistent with 'the flow parameters such as the size of the boundary layer along the main confining surface, and the boundary layer over the blade surfaces as effected by blade camber and blade chordwise and spanwise dimensions.

As a result of this invention it is apparent that a simple but effective means has been provided for increasing flow efficiencies through blades having a cascade arrangement as for example in compressors, diffusers and the like while maintaining high aerodynamic loading on the blades.

Although certain embodiments of this invention have been illustrated and described herein, it will be apparent that various changes and modifications may be made in the arrangement and construction of the various parts without departing from the scope of this novel concept.

What it is desired to obtain by Letters Patent is:

1. In a compressor having inner and outer walls of predetermined diameters respectively forming an annular passage, said passage having a longitudinal axis, a row of vanes fixed to said walls of airfoil shape and circumferentially spaced transversely of the axis of said passage, each of said vanes substantailly spanning said annular passage in a radial direction and extending from one of said walls to the other, said vanes having a chordwise length extending along the longitudinal axis of said passage, and a protrusion extending into said passage from at least one of said walls and spanning the entire space between circumferentially adjacent vanes, said protrusion having a smoothly curved shape diverging at a predetermined rate from the diameter of said one wall and subsequently converging at a greater rate to the diameterof said one wall in a direction downstream along said longitudinal axis, said protrusion having its point of maximum divergence located approximately within the last downstream quarter of the chordwise length of the vanes but upstream of the trailing edge thereof. 7

2. In a compressor according to claim 1 wherein said protrusion has its leading edge located within the first third of the chordwise dimension of said vanes.

References Cited in the file of this patent UNITED STATES PATENTS 1,333,142 Ulmer Mar. 9, 1920 2,017,043 Galliot Oct. 15, 1935 2,216,046 Peck Sept. 24, 1940 2,340,195 Maag Ian. 5, 1944 2,474,258 Kroon June 28, 1949 2,503,973 Smith Apr. 11, 1950 2,527,971 Stalker Oct. 31, 1950 2,575,682 Price Nov. 20, 1951 2,648,492 Stalker Aug. 11, 1953 2,648,493 Stalker Aug. 11, 1953 2,650,752 Hoadley Sept. 1, 1953 FOREIGN PATENTS 564,336 Great Britain Sept. 22, 1944 596,784 Great Britain Jan. 12, 1948 988,736 France Oct. 30, 1951

Patent Citations
Cited PatentFiling datePublication dateApplicantTitle
US1333142 *Apr 9, 1919Mar 9, 1920Ulmer TheodoreIntake-manifold
US2017043 *Sep 12, 1931Oct 15, 1935Galliot NorbertDevice for conveying gaseous streams
US2216046 *Apr 12, 1937Sep 24, 1940Robert E PeckAir conditioning conduit fitting
US2340195 *Apr 23, 1941Jan 25, 1944Maag George AAirplane construction
US2474258 *Jan 3, 1946Jun 28, 1949Westinghouse Electric CorpTurbine apparatus
US2503973 *Jan 29, 1946Apr 11, 1950Power Jets Res & Dev LtdAir intake arrangement for supersonic aircraft
US2527971 *May 15, 1946Oct 31, 1950Edward A StalkerAxial-flow compressor
US2575682 *Feb 14, 1944Nov 20, 1951Lockheed Aircraft CorpReaction propulsion aircraft power plant having independently rotating compressor and turbine blading stages
US2648492 *May 14, 1945Aug 11, 1953Edward A StalkerGas turbine incorporating compressor
US2648493 *Oct 23, 1945Aug 11, 1953Edward A StalkerCompressor
US2650752 *Aug 27, 1949Sep 1, 1953United Aircraft CorpBoundary layer control in blowers
FR988736A * Title not available
GB564336A * Title not available
GB596784A * Title not available
Referenced by
Citing PatentFiling datePublication dateApplicantTitle
US2788172 *Dec 6, 1951Apr 9, 1957Stalker Dev CompanyBladed structures for axial flow compressors
US2819837 *Jun 19, 1952Jan 14, 1958Laval Steam Turbine CoCompressor
US2918254 *May 10, 1955Dec 22, 1959Hausammann WernerTurborunner
US2944731 *May 17, 1956Jul 12, 1960Lockheed Aircraft CorpDebris traps for engine-air inlets
US3059834 *Feb 20, 1958Oct 23, 1962Werner HausammannTurbo rotor
US3069848 *Jan 11, 1960Dec 25, 1962Rolls RoyceJet lift gas turbine engines having thrust augmenting and silencing means
US3471080 *Jun 13, 1968Oct 7, 1969United Aircraft CorpLow noise generation fan
US3529631 *Jun 4, 1969Sep 22, 1970Gilbert RiolletCurved channels through which a gas or vapour flows
US3968935 *Jul 24, 1974Jul 13, 1976Sohre John SContoured supersonic nozzle
US4167376 *May 20, 1977Sep 11, 1979Papst-Motoren KgAxial fan
US4199296 *Jan 9, 1978Apr 22, 1980Chair Rory S DeGas turbine engines
US4208167 *Sep 22, 1978Jun 17, 1980Hitachi, Ltd.Blade lattice structure for axial fluid machine
US4305248 *Oct 5, 1979Dec 15, 1981The United States Of America As Represented By The Secretary Of The Air ForceHot spike mixer
US4315715 *Jun 8, 1979Feb 16, 1982Nissan Motor Company, LimitedDiffuser for fluid impelling device
US4465433 *Jan 17, 1983Aug 14, 1984Mtu Motoren- Und Turbinen-Union Muenchen GmbhFlow duct structure for reducing secondary flow losses in a bladed flow duct
US4512158 *Jun 16, 1983Apr 23, 1985United Technologies CorporationHigh blockage diffuser with means for minimizing wakes
US4844692 *Aug 12, 1988Jul 4, 1989Avco CorporationContoured step entry rotor casing
US5215439 *Aug 25, 1992Jun 1, 1993Northern Research & Engineering Corp.Arbitrary hub for centrifugal impellers
US5365731 *Apr 8, 1993Nov 22, 1994United Technologies CorporationEfficient anti-ice exhaust method
US5397215 *Jun 14, 1993Mar 14, 1995United Technologies CorporationFlow directing assembly for the compression section of a rotary machine
US6283713 *Oct 22, 1999Sep 4, 2001Rolls-Royce PlcBladed ducting for turbomachinery
US6471474Oct 20, 2000Oct 29, 2002General Electric CompanyMethod and apparatus for reducing rotor assembly circumferential rim stress
US6511294Sep 23, 1999Jan 28, 2003General Electric CompanyReduced-stress compressor blisk flowpath
US6524070Aug 21, 2000Feb 25, 2003General Electric CompanyMethod and apparatus for reducing rotor assembly circumferential rim stress
US6669445 *Mar 7, 2002Dec 30, 2003United Technologies CorporationEndwall shape for use in turbomachinery
US6969232Oct 23, 2002Nov 29, 2005United Technologies CorporationFlow directing device
US7189055May 31, 2005Mar 13, 2007Pratt & Whitney Canada Corp.Coverplate deflectors for redirecting a fluid flow
US7189056May 31, 2005Mar 13, 2007Pratt & Whitney Canada Corp.Blade and disk radial pre-swirlers
US7220100 *Apr 14, 2005May 22, 2007General Electric CompanyCrescentic ramp turbine stage
US7244104May 31, 2005Jul 17, 2007Pratt & Whitney Canada Corp.Deflectors for controlling entry of fluid leakage into the working fluid flowpath of a gas turbine engine
US7484935Jun 2, 2005Feb 3, 2009Honeywell International Inc.Turbine rotor hub contour
US7690890Sep 22, 2005Apr 6, 2010Ishikawajima-Harima Heavy Industries Co. Ltd.Wall configuration of axial-flow machine, and gas turbine engine
US7874794Mar 21, 2006Jan 25, 2011General Electric CompanyBlade row for a rotary machine and method of fabricating same
US7887297 *May 2, 2006Feb 15, 2011United Technologies CorporationAirfoil array with an endwall protrusion and components of the array
US8061980Aug 18, 2008Nov 22, 2011United Technologies CorporationSeparation-resistant inlet duct for mid-turbine frames
US8100643Apr 30, 2009Jan 24, 2012Pratt & Whitney Canada Corp.Centrifugal compressor vane diffuser wall contouring
US8141588May 21, 2008Mar 27, 2012Fuel Tech, Inc.Flow control method and apparatus
US8366399May 2, 2006Feb 5, 2013United Technologies CorporationBlade or vane with a laterally enlarged base
US8402769 *Jan 25, 2008Mar 26, 2013Siemens AktiengesellschaftCasing of a gas turbine engine having a radial spoke with a flow guiding element
US8425188Jun 7, 2012Apr 23, 2013Pratt & Whitney Canada Corp.Diffuser pipe and assembly for gas turbine engine
US8500399Apr 25, 2007Aug 6, 2013Rolls-Royce CorporationMethod and apparatus for enhancing compressor performance
US8511978 *May 2, 2006Aug 20, 2013United Technologies CorporationAirfoil array with an endwall depression and components of the array
US8517686Nov 20, 2009Aug 27, 2013United Technologies CorporationFlow passage for gas turbine engine
US8591176 *Nov 4, 2009Nov 26, 2013Rolls-Royce Deutschland Ltd & Co KgFluid flow machine with sidewall boundary layer barrier
US8926267Feb 14, 2013Jan 6, 2015Siemens Energy, Inc.Ambient air cooling arrangement having a pre-swirler for gas turbine engine blade cooling
US20040081548 *Oct 23, 2002Apr 29, 2004Zess Gary A.Flow directing device
US20100031673 *Jan 25, 2008Feb 11, 2010John David MaltsonCasing of a gas turbine engine
USRE30720 *Jul 12, 1978Aug 25, 1981 Contoured supersonic nozzle
USRE43710 *Jun 5, 2001Oct 2, 2012United Technologies Corp.Swept turbomachinery blade
DE1209807B *Feb 22, 1960Jan 27, 1966Rolls RoyceGasturbinenstrahltriebwerk mit einem die Umgebungsluft in einem aeusseren Ringkanal beschleunigendem Strahlapparat
EP0997612A2 *Oct 20, 1999May 3, 2000ROLLS-ROYCE plcBladed ducting for turbomachinery
EP1074697A2 *Aug 4, 2000Feb 7, 2001United Technologies CorporationApparatus and method for stabilizing the core gas flow in a gas turbine engine
EP1126133A2Feb 16, 2001Aug 22, 2001General Electric CompanyConvex compressor casing
EP1632648A2 *Aug 26, 2005Mar 8, 2006MTU Aero Engines GmbHGas turbine flow path
EP1799989A1 *Oct 6, 2005Jun 27, 2007Volvo Aero CorporationGas turbine intermediate structure and a gas turbine engine comprising the intermediate structure
EP2159398A2 *Aug 18, 2009Mar 3, 2010United Technologies CorporationSeparation-resistant inlet duct for mid-turbine frames
EP2518269A2Apr 27, 2012Oct 31, 2012Hitachi Ltd.Gas turbine stator vane
EP2541069A1 *Jun 26, 2012Jan 2, 2013Pratt & Whitney Canada Corp.Radial compressor diffuser pipe with bump to reduce boundary layer accumulation
WO1982002418A1 *Dec 30, 1981Jul 22, 1982Bessay RaymondTurbine stage
WO1992013197A1 *Jan 8, 1992Aug 6, 1992Northern Res & EngArbitrary hub for centrifugal impellers
WO1993022548A1 *Apr 13, 1993Nov 11, 1993United Technologies CorpExhaust vent for de-icing system of aircraft nacelle
WO1996000841A1 *Jun 28, 1994Jan 11, 1996United Technologies CorpFlow directing assembly for the compression section of a rotary machine
WO2004038180A1 *Oct 23, 2003May 6, 2004United Technologies CorpApparatus and method for reducing the heat load of an airfoil
WO2009082665A1Dec 18, 2008Jul 2, 2009Fuel Tech IncA flow control method and apparatus
Classifications
U.S. Classification415/208.1, 415/183, 415/914, 138/39, 415/199.5, 138/111
International ClassificationF02C7/04, F01D5/14
Cooperative ClassificationY10S415/914, F01D5/143, F02C7/04, Y02T50/673
European ClassificationF01D5/14B2B, F02C7/04