US2628768A

US2628768A - Axial-flow compressor

Info

Publication number: US2628768A
Application number: US657529A
Authority: US
Inventors: Kantrowitz Arthur
Original assignee: Individual
Current assignee: Individual
Priority date: 1946-03-27
Filing date: 1946-03-27
Publication date: 1953-02-17
Anticipated expiration: 1970-02-17

Description

Feb. 17, 1953 A. KANTRowlTz 2,628,758

v AxIAL-FLow COMPRESSOR Filed March 27, 1946 2 SHEETS-SHEET: 2

v I V EN TOR.

N QH.

Patented Feb. 17, `1953 UNITED STATES PATENT OFFICE AXIAL-FLOW COMPRESSOR ArthurfKantrowitz, Hampton, Vai ApplicationMa-rclr 2&7, 1946, Serial. No.. 657,529

, Thisinvention relates to the-shapes of. blades for use'in an axial-flow, compressor, andlis.I intended to produce aahigherpressure ratio per stage than Yhash-ithertobeen; effected. in this type of machine. Heretofore, in orderv tol-achievehigh efficiency in machines of this type, the local gas velocities entering the rotors and stators have been restricted tov less.` than the velocity of sound inthe gas. This has; meant thatthe. compression ratio producedfby'an axial-flow compressor stage has1 been. restrictecilv to about 1,:;3.

It is, aIrobj-ect of this` invention. to provide an improved method" for compressing lgas so that a largevr compression; ratio;v can: be obtained sin a singleystage and. Witlrhigh eiliciency. n

Anotherfobjectistoprovide an improved blade shape; and correlation. of; bla-ding` for effecting larger compression ratios; nzsingle stage, axial.- flow' compressors.

With this* inventionsupersonic: gas; velocity is used and the: resulting sho-ck; Waves; are controlled'so.v that high. compressor efliciency is ob` tainedrgin spitefoff, the, s-hockl.s It is this control `of the shockf5waves'. thatmakes it. possible to obtain suchA higher: pressure-risesv in a single stage.

Features-.I of.v the invention ,relate to: the bla-ding ofthe -;o'mpre'ssorand..` the preferred embodiment ot:the'invention'has bladesofnovelv shape andina particular: relation with one. another' for ef.` fecting decelerationV of. gas! iow` aga-insti the back'. pressurev with. control.; ofA the, shock Vwaves producing the. velocityy change.

Other olof|ects-v featuresand advantages of. the inventionfwill appear or hepointed-out as the-.descri-lotion prcceedsr In; the. drawing; formi-ng. a part-,thereof .in-.whichlike reference; characters indicate corresponding I parts inA all. the views,

Fig, L is a` perspective View of. a compresso-r embodying thisV invention, the casing being shown insectioni-norderv to: i-llustratei-the blading..

Fig. 2.is a fragmentary enlarged sectionalzview through-afpcrtion-ofthe rotor and stator of the compressor shown in Eig. 1I.. v

Fig. 3 is a development of the blading-a'lfeng thefcylinder A-Ao-f;1ig., 2. 1 rLheaxial.compressor shown inFig. 1 includes a casing `SI-l-iavfing an inlet opening. 9- at one end. Within,i thezcasing there is a fixed. stator hab IIL The gasor air enters the. compressor through the annular space Il between the inside wall of.v the casing Il` andtheperipheral surface ofthe stator hub.` The stator has. fixed blades I2 at equiangular positions around' its entire circumferenceand1t`heouter ends oi these blades are rigidly: 'co,nnectecl'` to theA casing 8..

2 Other stator blades t3' are similarlyv connected. to thecasing, and the bladesl3 arespacedfrom' the blades IZto. leave room for a rotor I4 having blades. II5- constructed. and arranged-Ito. draw air through. theA passages between the stator. blades I2, andto discharge. the. air through the. passages between. the stator blades I3'.

Fig. 2 shows the relation of one, oi the rotor. blades l5' to the stator 4blades I2V and?. 3`. The outer edge ofA the. blade [5, extends., close `to the wall of the compressor casing 8;..andinthe con hstruction illustrated. the radial'` length oiV the blade, I5 isv slightly greater toward'. its intake side. The internal. diameter. of the casing. 8dev creases proportionateli;f to leave, ample running clearance for the blade. it'.` `Constrnotion' avoids too rapid' an increase in the, cross section of.' the passage betweensuccessive blades. I5.. Too rapid an increase in the passage area. wouldjresult in dead air. regions in the, passages which would' lead `to reduced efficiency.

The. gas passages?" of` the. rotorare. the passages bounded on the sides by successive rotorl blades, on the bottom by the surface I'I' of" the rotor between blades, and on the` top by the i. side surface area of the casing across which the outer ends ofl they bladesswe'ep as-tlie rotor r`e volves. The rotor can be constructed"with` a cylindrical shell connectedwiththe. tips of the rotor blades. Such shrouds are used on turbines andA axi'alfflow compressors for the purpose, of preventing; tip clearance losses and.. for reducing` vibratiomand' are Well understood in the art.

The rotor I`4` is connected with adrivngi shaft I8' that tunis in bearings within the compressor..

and power is appliedt'o thedrive'shat I`8'fron`1'a motor or any external: source-of power;`

The priricilole` of the invention, the shape; ofv

the blades, and the correlation ofthe blades `can best beV unclersto'odv by considering. the` operation in connection with two successive. blades: wand As the. rotational speen"v of' anesnall flow' conf;-

pressor increased, a yspeed' is4 reached at'which the relative gas velocity' vexceeds theloeal speed of' sound attsome; point on the blad'ing. Ifftnei ahead of the. leading edge off the adjacent blade I5. A shockwave of' this. type Whichkexists up.- streaml fromy the leading edge ofi a. blade will be' referred to, herein. as a detachedll bow Waver These detached bow Waves existing. ahead' oi' each of the blades eventually extend far out from the cascade and form an extended wave system. The losses due to this extended wave system are so large that designers have avoided the occurrence of supersonic velocities in order to avoid these losses.

In order to maintain the compressor eiciency it is necessary to have the bow wave move back far enough to attach tothe next successive blade l. Such attachment prevents the setting up of an extended wave system.

The amount that the bow wave moves back along the blade l5 depends u'pon several different factors. One is the blade speed. The bow wave tends to move back as the velocity of the blade through the air increases. Another factor is the wedge angle, that is, the sharpness of the leading edge of the succeeding blade I5. The bow wave moves back further, at a given speed, ifV the leading edge of the succeeding blade is sharp. If the angle of the blade surfaces that meet to form the leading edge is greater than 90, it is not, possible in air to get attachment no matter how high the blade speed may become. Another factor affecting the ultimate position of the bow wave is the width of the narrowest section between the blades, that is, the area of cross section of the space between the blades I5 and I5 in a plane normal to the direction of the gas flow between the blades. If there is any considerable reduction in section along the passage between the blades the bow waves cannot become attached to the blades. For reasons that will become apparent, it is desirable to have a slightly restricted throat between the blades, and to have the shock wave between the blades downstream of the minimum section of the passage between the blades.

The position of the bow wave is affected also by the back pressure ratio, that is, the ratio of back pressure to inlet pressure. A higher back pressure ratio opposes downstream movement of the bow wave, but with sonic or supersonic velocity of the air at the minimum throat section 34, the back pressure ratio is not directly effective on theposition of the detached bow wave.

If,`however,the throat section is large enough, 'the back pressure low enough, the wedge angle small enough and the rotational speed high enough the bow waves become attached. The portion of the waves upstream from the cascade then nearly disappears and the portion between the blades moves downstream and is confined entirely within the blading. rIhis confined shock decelerates the air thru the speed of sound and in this process compresses the air efciently.

Although it is possible to accelerate gas to supersonic velocity without shock, it is not possible to decelerate through the sonic velocity without having a shock wave in the gas stream.

V`In the operation of the compressor, the rotor is driven at sufficiently high speed to make the relative velocity of the air entering the rotor blading higher than the sonic velocity. The air velocities relative to the stator blades, both inlet and discharge, is subsonic and the design of the stator blading is conventional.

` The blade surfaces at the leading edge meet in an angle less than 30 degrees and preferably less than 20 degrees. In the illustrated embodiment of the invention the blade faces form an angle of approximately degrees at the leading edge. The leading edge thus forms a sharp wedge so that bow waves can become attached to 4 the blades. In this specification and in the claims, the term sharp as applied to the leading edge designates an edge in which the effective angle between the forward and rearward blade surfaces is less than 30 degrees, and the leading edge radius is less than 2 percent of the circumferential distance between the leading edges of successive blades; or the blades are made of such thin material that the thickness of the leading edge is less than 2 percent of the circumferential distance between blades.

The cross-sectional area of the inlet of the gas lpassage between the blades I5 and I5 is indicated by the plane 33. All gas passage crosssections referred to in this specification and in the claims are areas of planes normal to the direction of gas ow through the passage at the location where the cross-section is taken.

As the result of the increase in thickness of the blades l5 and I5 downstream from the passage inlet 33, the cross-section of the rotor gas passage between these blades decreases progressively to a throat section 34 where the crosssectional area is a minimum, but not substantially less than the passage entrance or inlet 33. The area of the throat section 34 is approximately 93 percent of the cross-section of the inlet 33 in the compressor shown in the drawing. This value is given merely as an illustration.

The back pressure at the discharge-region 35 is so correlated with the inlet pressure and the rotor speed that the air flow decreases through the sonic velocity and undergoes normal shock at the region 36 just downstream from the throat 34. The gas passage between the blades l5 and I5 increases in cross-section beyond the throat 34, and this insures stability of the shock wave.

Although the shock wave tends to be substantially normal to the direction of gas flow, it is, in practice distorted by interaction with the boundary layers on the blades, the rotor and the casing. Angular waves given off from the boundary layers interact with the normal shock and sometimes result in multiple reflections. In the compressor of this invention, this entire normal shock complex is conned within the blade passage in the region 36; and the term region of shock, as used herein, refers to the space occupied by this normal shock complex.

the walls of the rotor gas passage. The blading,

therefore, must have sufficient chord length from the leading edges 3| to the trailing edges 31, as compared with the spacing between the blades, so that planes normal to the gas flow throughout the region of shock is bounded on both ends by the blade surfaces; and it is important that this be true at the top as well as at the lower ends of the blades where they connect to the body of the rotor, and are closer together than at their tips.

I In order to reduce the intensity, and hence the energy losses, in the normal shock complex at the region just beyond the throat 34, the rearward surface of each of the blades I5 is made with a concave portion 39 for a short distance back from the edge 3|. This concave region Ioriginates compression waves which reduce the velocity of the air smoothly before the air reaches the region of normal shock. These compression waves will thus reduce the intensty of the shock and the losses that go with it;

In order that supersonic*flbvvi intc the-fcascade mayl be startedyi-t isneces'sary that the narrowestsectioriof the passaaes-formed by the blades downstreamfrom the inlet section 33 be nearly as large as thearea available for gas ow at the inlet section 33. In starting the machine, an extended-'wave system is set up in the air and this entails large losses. In order tot haveV the desired fiowconditions; setup, it isnecessary that the air after undergoing these large losses, be stillt capable. of passing through the blading. If atlrroat with considerably lessair flow area than theinlet` section 3'3 were included in the -blade desigmthen it wouldnot be. possible for the air to get;through thisthroat in` the starting condition, and the vdesired"flow conditions'- could not be set up.

It will be understood that the blades l5 of the rotor may be used in the stator on the *discharge side of a rotor that delivers the gas to the downstream stator blades at supersonic velocity. The relative velocity ofthe blades and gas is the important consideration whether or not the blades are moving in space. The term cascade is used herein to designate a circle of blades whether secured to a rotating or stationary part of the apparatus. In any event, the blading offers the advantage that the gas having supersonic velocity with respect to the gas passages can decelerate to a subsonio value with the shock waves contained entirely within the blade passages.

In the case of stationary blading the forward and rearward surfaces of the blades are the surfaces corresponding to the forward and rearward surfaces of the blades when used on a roy tor. The direction of the `air entering the passages between the blades has both an axial and la tangential component, and the direction of the tangential component is toward the f-orward surfaces of the blades and away from their rearward surface.

The preferred embodiment and method of this invention has been illustrated and described, but `changes and modifications can be made and some featuresof the invention can be used in diierent combinations without departing from the invention as defined in the claims.

I claim as my invention:

1. A supersonic compressor including a cascade of blades, angul-arly spa-ced from one another around a surface of revolution, said blades having sharp leading edges which'are disposed at acute angles to a plane normal to the axis of the body of revolution and being of suicient chord length so that successive blades intersect common parallel planes that are spaced from one 'another by a substantial distance and that are normal'to the direction of gas now between the blades, the confronting faces of said blades diverging from one another in the direction of gas now for at least a portion of their length between the parallel planes.

2. A supersonic axial-flow compressor comprising a casing, a bladed rotor supported within the casing for rotation with running clearance between the blade tips and the inside surface of the casing, angularly spaced blades on the rotor with sharp leading edges disposed at acute angles to both the axis of rotation and the plane of rotation of the rotor, said blades being of suiiicient chord length so that successive blades intersect common parallel planes that are spaced `from one another by a substantial distance 4and that are normal to the direction of gas 'new' between the: biaces=,{ the l weborama faces of said blades diverging from one another in the direction of gas flow' for atleast apos#- tion of their length betweengthe parallel planes.

3. A supersonic axial-how` compressor` coni- DI'SIL,4 a Cascade haV'ng angulgfly; Spacedprofor rearward face just beyond the leading edge of the blade, each of said blades increasing in thickness along the portion of its length that has said concave surface.

5. An axial-flow compressor for gas having supersonic velocity, said compress-or comprising a casing and a bladed rotor that rotates within the casing with running clearance between *thel blade tips of the rotor and the inside surface of the casing, said rotor having blades of sulcient chord length to enclose a gas passage in which a plane perpendicular to the gas ilow through the passage intersects the confronting faces of the blades that form the gas passage along a region of the passage where the blades areshaped to provide rst a throat and then a region of expanding cross-section in the direction of gas flow, the blades having concave surfaces on their rearward faces at regions spaced va short distance beyond the leading edges of the blades for controlling the gas velocity, each of said blades increasing in thickness along the portion of its length that has said concave surface.

6. A supersonic axial-flow compressor comprising a stator, a rotor, and blades on the rotor having leading edges formed by forward and rearward blade faces that meet in an effective angle of less than 30 degrees, and with the bisector of such angle at an acute angle to the plane of rotation of the rotor, the rearward faces of the blades just behind the leading edges being concave for controlling the velocity of iow, each of said blades increasing in thickness along the portion of its length that has said concave surface and successive blades of the rotor being of sufiicient chord length to intersect common parallel planes that are spaced from one another by a substantial distance and that are normal to the direction of gas ow between said blades, the confronting faces of said blades diverging from one another in the direction of gas flow for at least a portion of their length between the parallel planes.

7. An axial-now compressor including a stator, a rotor having blades for moving gas from the stator at supersonic veloci-ties with respect to the leading edges of the rotor blades, successive blades being shaped to provide gas passages with entrances that have a cross section slightly larger than the cross-section of the passages at an intermediate thro-at region, and said blades `being of suicient length so that successive blades intersect common parallel planes that are spaced from one another by a substantial distance and that are normal to the direction of gas flow in the passage just beyond the throat in the downstreamY direction, the confronting faces of said blades diverging from one another in the direction ci gas flow for at least a portion of their length that lies beyond the throat but between the parallel planes.

- 8. A supersonic velocity compressor including a cascade of blades angularly spaced around a surface of revolution and havingr leading edges at their upstream ends and portions of greater thickness intermediate their ends providing surfaces that cooperate Wit-h confronting surfaces of preceding and succeeding blades to form gas passages that decrease slightly in cross section to a throat area, some distance back from the leading edges of the blades, and then increase in cross section toward the outlet ends of said passages, and concave surfaces on the rearward faces of the blades upstream from said throat areas.

ARTHUR KANTROWITZ.

REFERENCES CITED The following references are of record in the nie of this patent:

UNITED STATES PATENTS Number Name Date 408,864 Vogelgesang Aug. 13, 1889 853,363 Holzwarth May 14, 1907 997,678 Jalonick July 11, 1911 15 2,378,372 Whittle June 12, 1945 2,435,236 Redding Feb. 3, 1948