US2805818A

US2805818A - Stator for axial flow compressor with supersonic velocity at entrance

Info

Publication number: US2805818A
Application number: US261571A
Authority: US
Inventors: Ferri Antonio
Original assignee: Individual
Current assignee: Individual
Priority date: 1951-12-13
Filing date: 1951-12-13
Publication date: 1957-09-10
Anticipated expiration: 1974-09-10

Description

p 10, 1957 A. FERRI 2,805,818

STATOR FOR AXIAL FLOW COMPRESSOR WITH SUPERSONIC VELOCITY AT ENTRANCE Filed Dec. 13, 1951 4 Shee ts-Sheet 1 FIG. 2

FIG. I

- INVENTOR. Q ANTON/0 F ERR/ ATTORNEYS sept- 0, 1957 A. FERRI 2,805,818

STATOR FOR AXIAL FLOW COMPRESSOR WITH SUFERSONIC VELOCITY AT ENTRANCE Filed Dec. 13, 1951 4 Sheets-Sheet 2.

FIG. 3

INVEN TOR.

ANT/W0 FER/5! Maxim A T TQR/VEYS p 10, 1957 A. FERRl 2,805,818

STATOR FOR AXIAL. FLOW COMPRESSOR WITH SUPERSONIC VELOCITY AT ENTRANCE Filed Dec. 13, 1951 4 Sheets-Sheet 3 inn? INVENTOR. AN TON/O FER AT TORNE Y8 p 10, 1957 A. FERRI 2,805,818

STATOR FOR AXIAL. 110w COMPRESSOR WITH SUPERSONIC VELOCITY AT ENTRANCE Filed Dec. 13 1951 4 Sheets-Sheet 4 FIG. 6

sussomc DIFFUSER zomz SUPERSONIC INVENTOR. ANTON/0 FER/i! ATTORNEYS FIG. 8

United States Patent STATOR FOR AXIAL FLOW' CQMPRESSOR WITH SUPERSQNIC VELGCETY AT ENTRANiIE Antonio Fen-i, Rockville Centre, N. Y. Application December 13, 1951, Serial No. 261,571

1 Claim. (Cl. 230-114) (Granted under Title 35, U. S. Code (1952), see. 266) This invention relates to improvements in compressors, and more particularly, deals with improvements in stators for compressors which are provided with upstream rotors constructed and arranged to deliver fluid to the stators at Mach numbers greater than unity.

Previous stators were capable of handling only subsonic fluid flow throughout and could not, except possibly with prohibitively low efficiency, narrow range operation and flow instability, accomplish the main object of this invention, which is to provide a means for efriciently diifusing air or other fluid from supersonic to subsonic velocity in the flow through a row of stator blades.

The major, although not exclusive, fleld of use of the invention is in present day high speed aircraft, wherein efliciency of operation is mandatory. The instant invention makes possible an eflicient compressor in which the velocity of the gas leaving the rotor is supersonic, realizing the high pressure ratio and high efliciency which a compressor of this type is capable of producing. The inability of previous stators to decelerate the flow from supersonic velocity with reasonable efliciency has limited the design of compressors to the type in which the velocity of the gas leaving the rotor is subsonic. In this type, the deceleration from supersonic to subsonic velocity occurs in the rotor, and the angle through which the gas is turned by the rotor is restricted to a small value, with the result that the maximum obtainable pressure ratio per stage is inherently lower than for a machine with supersonic velocity at the rotor exit.

Accordingly a more specific object of the invention is to provide an improved compressor stator which has a plurality of blades dividing the compressor passage into a group of channels which function as supersonic difiusers, the contraction ratio of each being larger than the maximum contraction ratio of a converging-diverging supersonic diffuser calculated from one-dimensional flow theory, the construction being such that if the velocity of fluid has a component parallel to the compressor rotation axis, which is supersonic, the lower surface of each channel makes an angle with the entering velocity at the leading edge such that a shock wave is produced at the design condition of operation, the shock wave meeting the upper surface of the channel slightly behind the leading edge and being cancelled or only partially reflected. But if the axial component of entering velocity is subsonic, the lower surface of each channel is parallel to the entering velocity at the leading edge, and the region behind the leading edge is so shaped that it produces compression waves which are confined within the channel, the upper surface of the channel making an angle with the direction of entering velocity at the leading edge.

Another object is 'to provide difiusers constituted by the stator blades which have two zones therein, the upstream zone being for supersonic fluid entry, while the downstream zone has the fluid turning in the axial direction therein, this being the subsonic zone.

Other objects and features will become apparent in 'ice following the description of the illustrated embodiments of the invention.

In the drawings:

Fig. l is a fragmentary sectional and largely diagrammatic view of a typical compressor having one of the stators therein;

Fig. 2 is a fragmentary perspective view of another stator constructed in accordance with the principles of the invention;

Fig. 3 is a fragmentary perspective view of the stator in Fig. 1;

Fig. 4 is a fragmentary perspective view of a further form of the invention;

Fig. 5 is a fragmentary perspective view of another form of the invention;

Fig. 6 is a diagrammatic view of several stator blades which are shaped to dilfuse fluid efliciently when the axial component of entry velocity is subsonic although the entry velocity is supersonic;

Fig. 7 is a diagrammatic view of several stator blades which are shaped to diffuse fluid efficiently when the axial component of entry velocity is supersonic; and

Fig. 8 is a schematic view applicable to any of the illustrated stators, showing compressor operation with the use of the invention.

In Fig. 1, there is a fragmentary part of a compressor 10 of the type adapted for use in an aircraft, the compressor including a housing or casing 12 which has an annular fluid, as air or other gas, passage 14 extending therethrough. A rotor 16 is mounted in the casing 12 in advance of a stator 18, the rotor 16 being capable of delivering fluid to the stator 18 at supersonic axial speeds. Such arrangement is typical of each embodiment disclosed herein.

In Fig. 2, the casing 12 is provided with a boundary layer removal slot 20, such an expedient being optional. The rotor 16 includes among other standard elements, a hub 22 from which extend the rotor vanes 24, buckets or the like. The requirement of the rotor 16 is that it deliver the working fluid at Mach numbers greater than unity. A large number of compressors today rely on the rotor to difiuse the fluid, the present invention (see Fig. 8) having rotor delivered fluid moving at supersonic axial velocities, with diffusion taking place within the stator in order to realize the benefits of a higher compression ratio in the system. Stator 26 includes a plurality of blades or

vanes

28, 29, 30 and 31 which are mounted on

spindles

32, 33, 34 and 35, the latter being rotatively carried by the casing 12 and a stator ring 38 which is located in the casing 12. Since the spindles are radially arranged in the casing and the stator blades are mounted axially thereof, the

blades

28, 29, 30 and 31 are also radial.

This stator has its

blades

28, 29, 30 and 31 and all other identical blades so constructed that the angle setting of the blades may be varied during, prior to, or after operation of the compressor. The arrangement is such that the contraction ratio may be changed along with the setting of the blades. A general discussion of contraction ratios and the matter of figuring them is set forth in the publication entitled Elements of Aerodynamic Supersonic Flow, published in 1949 by the McMillan Company at pages 181 to 195. In order to adjust the angle setting of the blades, mechanical means may be used, as the gear 40 mounted in bearings in the compressor casing 12, which is engaged by pinions 42, one being fixed to each spindle. Upon rotation of the gear 40 by standard means, as the arm 44 which is ultimately actuated by mechanical, electrical or hydraulic equipment, the pinions, spindles and stator blades are rotated to obtain the desired setting. It is to be observed that the stator blades have straight leading and trailing edges but are-contoured to allow for diiferences in velocity of the fluid entering the root and tip portions of the stator 26. Moreover the blades ,of the stator 26, as in each embodimennhave sharp leading edges, this being important to have efiiciency in'accepting supersonic flow from the rotor.

The blades of the stator 26-are staggered and spaced from each other to provide channels '44, 46, 48 and others, each of which serves as a diffuser. The. upstream zone of each diifuse-r accepts and'diffuses supersonic flow, while the downstream zone of each channel constitutes a subsonic zone.

In Fig. 3 the casing is identical to the casing 12, and the rotor 52 is the same as its counterpart in Fig. 2. Stator 54 is located behind the rotor 52 and behind the boundary layer removal slot 156 which is formed in the casing 50. The compressor of a jetengine must operate efliciently over a wide range, In this connection I have made each blade 56 of a frontsection 58 and a rearsection 60. The front sections 58 of the blades form the walls of the supersonic zone of the channels in the stator 54 While the rear sections 60 form the walls of the subsonic zone of each channel. The sections 60 are fixed to the stator ring 62, but the sections 58 are secured to the spindles 64, whereby upon rotation of the gear 66, the sections 58 of the blades 56 are adjusted thereby changing the capacity and character of the supersonic zones of the diffuser channels. Such adjustment should be made in accordance with the output of the rotor 52 in order to maintain the highest available efiiciency of the stator.

Another embodiment is illustrated in Fig. 4, wherein the casing 70 is provided with a stator 72. In this instance part of the turning of the flow occurs in the subsonic region and, is produced independently of the supersonic zone. Structurally, the stator 72 consists of a ring 74 which is provided with a series of staggered fixed vanes 76, spaced to constitute channels forming the subsonic zone in the stator. In advance of the vanes 76 there is a plurality of blades 78, each beingspaced from the other to form channels constituting diffusers. The means of actuating the blades 78 are the same as the means of operating the blades 80 of the embodiment of Fig. 5 and also the previously described blade adjusting means.

The distinction between the forms of Figs. 2 and 5 is as follows: The blades 28 have straight leading and trailing edges, While the blades 80 are twisted in order to allow difierent velocity diagrams at the root and tip of each blade. Hence the

blades

28 and 80 present alternate solutions'for the same problem of varying the velocity diagrams at the root and tip of the blades.

For more detailed operation; attention is invited to Figs. 6-8. In Fig. 6 the blades 84 are shaped slightly different from the blades 85 of Fig. 7. It has been found that by adjusting the stator blades through rotation of their spindles, the contraction ratio may be varied and in addition the efliciency may be maintained high through a rather large range of Mach numbers. However, in Figs. 6 and 7 the stator blades are fixed, taking into consideration two operational conditions. Even though the entry flow may be supersonic in each stator (Figs. 6 and 7), in Fig. 6 the design of the channels 86 is such as to accept flow when the axial component of velocity is supersonic, and the channels 87 are designed to accept flow when the axial component is subsonic.

For supersonic flow where the component of the fluid flow parallel to the axis of rotation of the compressor is supersonic, the blades 84 are so shaped that the lower surface 88 of each channel 86 makes an angle with the entering velocity at the leading edge such that a shock wave 90 is produced at design condition of operation. The shock wave meets the upper surface 91 of the channel 86 slightly behind the leading edge and is cancelled or only partially reflected.

It" the axial component of the entering velocity is sonic even though total flow is supersonic, the lower surface 92 of each channel is made parallel to the entering a velocity at the leading edge, and the surface behind the leading edge is contoured to produce compression waves 93 which are confined within the channel. The upper surface of the channel 87 makes an angle with the direction of entering velocity at the leading edge, and the shock wave 94 is produced as illustrated. By use of stators shown in Figs. 6 and 7 two design conditions are considered. In the embodiments which have movable sections or blades the contraction ratio (h1=height of entering air stream divided by h2=height of diffuser I channel at smallest point) may be varied, and the-setting may be made such as to efliciently handle supersonic flow whether the axial component is supersonic or subsonic. Accordingly the movable blade embodiments are capable of adjustment to various positions which make it possible to realize the operation as described in connection with both of the fixed vane embodiments.

It is apparent that modifications may be made as come within the scope and purview of the claim.

The invention described'herein may be manufactured and used by or for the Government of the United States of America for governmental purposes without the payment of any royalties thereon or therefor.

What is claimed is:

A stator for an axial flow gas compressor having a flow passage in which the axial component of gas velocity is supersonic, said stator comprising a stator ring, a plurality of radially extending blades aboutthe periphery of said ring, said blades having sharp leading edges and dividing said passage into a plurality of supersonic diffusers, each of said blades-being fixed on a radial spindle, a pinon on each spindle inwardly of said blades, a ring gear engaging said spindle, and means to rotate said ring gear, whereby said blades may be rotated to vary the angle of attack of each blade so that the leading edge of each blade may be varied in accordance with the velocity of the gas stream.

References Cited in the file of this patent UNITED STATES PATENTS 1,544,288 Van Ormer June 30, 1925 2,224,519 McIntyre Dec. 10, 1940 2,258,793 New Oct. 14, 1941 2,316,452 Pfenninger Apr. 13, 1943 2,341,974 Browne Feb. 15, 1944 2,384,000 Wattendorf Sept. 4, 1945 2,435,236 Redding Feb. 3, 1948 2,455,251 Hersey Nov. 30, 1948 2,555,576 Criqui June 5, 1951 2,575,682 Price Nov. 20, 1951 2,613,029 Wilde Oct. 7, 1952 2,623,688 Davidson Dec. 30, 1952 2,659,528 Price Nov. 17, 1953 2,689,680 Lovesey Sept. 21, 1954 FOREIGN PATENTS 7 580,841 Great Britain Sept. 23, 1946 615,219 Great Britain Ian. 4, 1949 724,553 Germany Aug. 29, 1942