US2870958A

US2870958A - Mixed blade compressor

Info

Publication number: US2870958A
Application number: US558874A
Authority: US
Inventors: Edward A Pinsley
Original assignee: United Aircraft Corp
Current assignee: Raytheon Technologies Corp
Priority date: 1956-01-13
Filing date: 1956-01-13
Publication date: 1959-01-27
Anticipated expiration: 1976-01-27

Description

Jan. 27, 1959 PlNSLEY 2,870,958

MIXED BLADE COMPRESSOR Filed Jan. 15, 1956 INVEN TOR EDWARD A. PINSLEY A TTORNEV a United States Patent OfiFice.

Patented Jan. 27,1959

MIXED BLADE COMPRESSOR Edward A. Pinsley, East Hartford, Conn assignor to United Aircraft Corporation, East Hartford, Conn, a corporation of Delaware Application January 13, 1956, Serial No. 558,874 3 Claims. (Cl. 230-122 This invention relates to turbomachinery and more specifically to compressors and the arrangement of the blading therefor.

, Itis an object of this invention to provide a blading arrangement for turbomachinery which prevents the formation of rotating stall in any given row of blading in turbomachinery.

This and other objects of this invention will become readily apparent from the following detailed description ofth'e drawing: a

Fig. 1 is a cross-sectional illustration of an axial-flow 1 row having transonic or supersonic type of blading, and

illustrates a mixture of transonic typeblading according to this invention. H

Referring to Fig. l, a turbojet power plant is generally indicated at as having an axial flow compressor section 12, a burner section 14, a turbine 16 for driving the compressor 12 and an exhaust nozzle 18.

The compressor 12 may be subject to rotating stall during certain operating conditions so as to adversely affect the life of the compressor. It is the purpose of this invention to avoid this undesirable stall condition. At certain ofl-design operating conditions ofaxial-flow compressors, for example, it is known that axially unsymmetrical regions of stall exist in one or more blade rows. These stall cells rotate about the axis of rotation of the compressor at some fraction of the rotor speed and produce fluctuating air loads of unpredictable frequencies on the individual blades. These fluctuations, in turn, cause mechanically destructive vibratory stresses. In

any given row of blading the phenomenon is manifested as a circumferentially propagating blade stall where the stall progresses from one blade to the next blade adjacent to its suction surface. The tendency of a compressor, for example, to operate in a rotating stall regime may be reduced by the installation of alternate blades of different geometry in one or more stator or rotor rows.

Referring to Fig. 2, a cross section through several blades in a conventional blade row is illustrated. The direction of flow relative to the blade row is indicated by the arrow 19. As the total flow through the machine decreases, the angle-of-attack of the relative flow increases, thereby increasing the tendency of the blades to stall. One blade, as for example blade 20, in any given row will stall first, the stall region being illustrated by separation of the flow from the blade suction surface and the a formation of the dead-flow region 21. This stalling of the blade 20 in turn causes a local increase in the relative angle-of-attack of the blade adjacent to its suction surface, as for example blade 22. This local increase of angle-of-attack is illustrated for example by the arrow 23.

The increased angle-of-attack causes blade 22 in turn to stall and this stall condition progresses such that next the blade 24 begins to stall, etc. When any one blade stalls, asfor example the blade 20, the blade passage 26 between blade 20 and blade 22 is referred to as a stalled blade passage. As the number of stalled blade passages increases the flow through the remaining passages of that particular row of blades must in turn increase, thus reducing the relative angle-of-attack of the remaining blades and hence. reducing the'tendency of these blades to stall Eventually, the relative angle-of-attack of, for example, blade 20 becomes small enough to permit unstalled flow which in turn reduces the tendency of the next adjacent blade (blade 22) to stall. The unstalling effect lags the stalling effect by the width of one ormore' blade passages ,giving rise to a propagating stall cell. It should be noted that at any one time there may be more than one of these stall cells propagating in any given blade row but the essential flow mechanics is unchanged.

Fig. 3 shows a cross section through a blade row incorporating the present invention. As shown, blades 30, .33 and 35 are chosen for good high Machnumber .flow performance at the design point while blades 32 and 34 are chosen for proper unstalled performance at higher angles-of-attack than the

blades

30, 33 and 35 With proper choice of blading it is possible to obtain uniform exit flow angles for the blade row through most ofthepperating range of the machine. Thus as the flowthrough the blade row in. Fig. 3 is decreased, a bladeof the type such as 30, 33 or 35 will stall. Although the flow through the passage 38 between blades 30 and 32 will decrease, the local increase of angle-of attack, as shown by the arrow 40 adjacent the leading edge of the blade; 32, will not be largeenough to cause blade 32 to stall provided the unstalled operating range of blade 32 is sutficiently large. Thus for some flow range any or all of the blades such as 30, 33 and 35 may stall without initiating the stall propagation obtained by the conventional blade row having blades of identical shape.

Under conditions of machine operation during which the total flow continues to decrease beyond the point where all the blades such as 30, 33 and 35 are stalled,

.the condition will eventually be reached Where a blade a the flow about blade 33 since it will be already operating in the stalled condition. Thus with all the blades tending to operate at a stalled condition, the tendency for stall propagation and hence rotating stall to exist at this operating point will be minimized.

Thus in constructing a row of blading which is in accordance with the present invention and which will provide the desired air turning angle at the design relative inlet air angle the following procedure is employed. First, one selects a blade profile for a given radial station which will satisfy the above requirements at the proper blade spacing and which has a small radius of curvature at the leading edge, as for example, an NACA (National Advisory Committee for Aeronautics) 65-series blower blade section. Second, one would select a blade profile having a larger radius of curvature at the leading edge, for example, an NACA (National Advisory Committee for Aeronautics) four-digit profile, which would also satisfy the design turning requirements at the proper blade spacing. These two types of blades would be alternated circumferentially in any given blade row, either rotor or stator. In any event the alternate blades are chosen such that one has a relatively large leading edge radius and the other has a relatively small leading edge radius. In other words, the camber of the suction surface at the leading edge of alternate blades is relatively high (large leading edge radius) as compared to the camber of the leading edge of the suction surface of the remaining blades (low leading edge radius).

The present invention may be employed in blade rows where the inlet air approaches the blade leading edge at supersonic relative velocities. For efficient performance at design speeds, the leading edges of these blades must be sharp. Fig. 4 shows the blading of a conventional transonic or supersonic rotor. As the flow through such a blade row decreases during off-design operation, the relative approach velocities eventually become subsonic and rotating stall may form in the same manner as in the case of blades with rounded leading edges.

Fig.6 shows a blade row having sharp leading edges which also incorporate the present invention. The leading-edge camber of the suction surfaces of alternate blades has been increased to decrease the susceptibility to stall at high angles-of-attack and subsonic velocities. The remaining blades will stall at lower angles-of-attack and hence rotating stall will be prevented in the same manner as discussed in connection with Fig. 3.

As a result of this invention, an increase in surge margin may be obtained in compressors as a result of the improvement in the rotating-stall characteristics. In addition, the new result is obtained aerodynamically, rather than by some mechanically moving parts. In addition, while there being no weight penalty involved, the improvement in rotating-stall characteristics will decrease vibratory stress levels, allowing in at least some cases, a reduction in blade size and weight.

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

I claim:

1. A rotor for an axial flow compressor having a plurality of stages, at least one of said stages comprising a plurality of circumferentially spaced blades, and adjacent blades being of different blade section whereby one of said blades has a relatively sharp leading edge and the next adjacent blade has a relatively blunt leading edge, with the blades with blunt leading edges alternating with blades with sharp leading edges, all of said blades having leading and trailing edges terminating in the same positions respectively along the axis of flow and all of the blades having relatively sharp trailing edges.

2. A rotor for an axial flow compressor having a plurality of stages, at least one of said stages comprising a plurality of circumferentially spaced blades each having relatively sharp trailing edges, and adjacent blades being of dilierent blade section whereby one of said blades has a suction surface terminating at its upstream end in a relatively high cambered leading edge and the next adjacent blade has a suction surface terminating at its upstream end in a relatively low cambered leading edge and all of the suction surfaces having the same relative position with respect to each other, alternate blades being of similar contour and all of the blades being of' substantially the same chordwise dimension.

3. A rotor for an axial flow compressor having a plurality of stages, at least one of said stages comprising a plurality of circumferentially spaced blades, and adjacent blades being of different blade airfoil section, one of said blades being of the NACA -series and the next adjacent blade being of the NACA four-digit series,said blades of different series being alternated circumferentially, all of said blades having the same chordwise dimension and having relatively sharp trailing edges.

References Cited in the tile of thispatent UNITED STATES PATENTS 2,351,516 Jandasek June 13, 1944 2,406,499 Iandasek Aug. 27 1946 FOREIGN PATENTS 1566/26 Australia Apr. 26, 1926 276,239 Switzerland June 30, 1951 685,979 Germany Dec. 29, 1939