US2934259A

US2934259A - Compressor blading

Info

Publication number: US2934259A
Application number: US592118A
Authority: US
Inventors: George F Hausmann
Original assignee: United Aircraft Corp
Current assignee: Raytheon Technologies Corp
Priority date: 1956-06-18
Filing date: 1956-06-18
Publication date: 1960-04-26
Anticipated expiration: 1977-04-26

Description

April 26, 1960 G. F. HAusMANN COMPRESSOR BLADING 2 Sheets-Sheet i Filed June 18, 1956 Tf/e r y t nx.

/ ,ear/swam cam/E/vr/o/VAL ma` mp/4f@ Fo@ Max/MUM N VE N TOR GEORGE E HUSMNN Br vf/ A T TORNEY April 26, 1960 G. F. HAUSMANN 2,934,259

COMPRESSOR BLADING Filed June 18, 1956 2 Sheets-Sheet 2 M (HX/HL) Max/Mam rH/cA/A/ss /N VEN TOR GEORGE E HAUSMANN ATTORNEY United 'States Patent O F COMPRESSOR BLADING George F. Hausmann, Glastonbury, Conn., assignor to United Aircraft Corporation, East Hartford, Conn., a corporation of Delaware Application June 18, 1956, Serial No. '592,118

4 Claims. (Cl. 5230-122) This invention relates to compressors and more specically to axial flow transonic or supersonic compressors.

It is an object of this invention to provide a compressor having a blade coniiguration which provides maximum ecieney for a given pressure rise across the blading while conforming to structural limitations imposed by stress considerations.

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

Fig. 1 is a partial schematic and partial cross-sectional View of a compressor rotor or stator having conventional blading therein.

Fig. 2 is similar to Fig. 1 but illustrates the proposed blading of this invention.

Fig. 3 is a schematic illustration in partial cross section showing a modified form of blading according to this invention, and

Fig. 4 is a plan view of one of the blades of Fig. 3.

1t has been determined from two-dimensional transonic and supersonic cascade tests that the total pressure losses are minimized in axial ow compressors for most conventional blade shapes when the blades are arranged at a gap-chord ratio (r/c) of approximately .5. However, the use of this gap-chord ratio (f/c) at the tip sections (when the relative velocities are transonic or supersonic) requires a gap-chord ratio at the root which is lower in proportion to the hub-tip diameter ratio. Since the use of a gap-chord ratio lower than .3 or .4 at the root of the blades of an axial flow compressor would severely compromise the performance of these sections (choking or low critical Mach numbers), it would be necessary to utilize inverse taper on the planform of the blade to obtain the desirable aerodynamic characteristics at the root and tip. Furthermore, it would also complicate the mechanical structure of the blade attachment fixtures because of the space limitations on the compressor rotor disc. However, the use of inverse taper imposes a low stress limit on the blade.

The blade passage described herein would provide the desirable aerodynamic characteristics of a low gap-chord ratio at the compressor tip without resorting to inverse planform taper.

Fig. l depicts an array of blades at the tip section of a conventional transonic or supersonic axial flow compressor. '-It has been determined from experiment that blade passages of this type have a minimum total pressure loss for a given pressure rise and turning, which describe the pressure rise in the corresponding compressor element. This particular cascade has a gap-chord ratio (T/c) of approximately .5, which is undesirable from compressor stress considerations. The maximum thickness of the blades also occurs at approximately the 50% chord position. The favorable aerodynamic performance of this cascade is due to the characteristics of the flow at the location of the intercept of the shock wave from the leading edge of the succeeding blade on the upper surface of the preceding blade. The Mach Y P2,934,259 Patented Apr. 26, 1960 ICC number at this intercept is essentially the same as the relative inlet Mach number and the boundary layer thickness is lower than that at any other -point within the passage. If the gap-chord ratio (r/c) were increased, the Mach number and boundary layer thickness at the shock intercept would both be increased. Downstream of the shock intercept in Fig. 1, the ow diffuses and turns subsonically. However, the pressure rise and compressor work associated with this process are small relative to the turning and pressure rise which is obtained across the shock wave system.

In order to obtain these desirable aerodynamic characteristics without severely compromising the structural considerations, it is proposed to terminate the chord of the blades shown in Fig. l at a position slightly downstream of the shock intercept, as shown in Fig. 2. This modication places the maximum thickness at approximately the -percent chord station 'and results in a blunt, or relatively blunt trailing edge. In this respect, the subsonic diffusion and turning would be compromised. However, it has been demonstratedby test that this cornpromise is not severe. The test results have indicated that the favorable shock-interaction characteristics are maintained, whereas the gap-chord ratio is essentially doubled.

It is believed that the nomenclature shown in Figs. 1 and 2 is sufficiently clear so that extended application of reference characters is not necessary. lt might be pointed out, however, that in the Fig. 2 configuration according to this invention the maximum thickness of the blading is approximately at the location indicated by the numeral 10. This location may range from 75% of the chord, as shown, or in any position aft thereof to the chord location. This makes the gap-chord ratio (r/c) ranging between .75 to 1.0. It should be added that the low pressure side of each blade will have a surface forward of the maximum thickness point which is substanitally aligned or`parallel to the direction of the relative inlet flow. The point of maximum thickness then forms the passage throat between the blades in cooperation with the bottom surface (high pressure side) of the next adjacent blade. The bottom surface of each blade will be at an angle to the relative inlet ow an amount approximately equal to the Wedge angle of the blade.

The blade section of this invention would be used only on the outboard portion of the blades where the relative approach Mach numbers are greater than one. Thus the type of blade shown in Figs. 3 and 4 is one embodiment of this invention. In other words, the flow becomes more critical toward the outboard end of the blades 20, 22 so that the maximum thickness may be located as, for example, as shown respectively at 24 and 26 in the 50% chordwise location at the root portion. However, adjacent the tip portion of these blades the maximum thickness point 30 is located in the last quarter of the chordwise dimension of the blade section at that point. The blade 20 is better illustrated in Fig. 4 illustrating that the maximum thickness point of the blade 24 is located approximately at the 50% chordwise position at the root portion of the blade and the maximum thickness point changes until it reaches approximately a location within the last 25% of the chordwise dimension adjacent the tip of the blade 20.

it should be added that the trailing edge of the blades of this invention aft of the maximum thickness point is relatively blunt compared to the leading edge of the blades. In other words, aft of the maximum thickness point the blade may be made relatively blunt or rounded to any suitable shape commensurate with good design practice and fabrication limitations.

As a result of this invention, it is apparent that a sim-` ple yet highly efficient blade configuration has been prothe construction and arrangement of the various parts n without departing from the scope of this novel concept.

I claim:

1. A compressor comprising a casing structure, a compressor rotor journaled axially in said casing structure and defining therewith an annular ow passage, blading carried by said rotor and extending across said annular flow passage in a cascade arrangement, said rotor blades being substantially arfoil in section from root to tip and having a substantially sharp leading edge and a relatively blunt trailing edge, said blades having their point of maximum thickness adjacent said trailing edges, the low pressure or top of the leading blade having at least two surfaces at an angle relative to each other and the high pressure or bottom of the next adjacent blade having a single surface at a small angle relative to the oncoming stream, said blades ybeing arranged so that the inlet edge of each blade together with the maximum thickness portion of the adjacent leading blade defines a restricted throat in the passage between the blades, said throat being substantially aligned with the direction of how of the gaseous medium flowing through the blade passage for effecting a normal compression shock to gaseous medium entering said throat at superacoustic velocity relative to the moving blades and thereby provide an immediate reduction in the velocity of said medium and an immediate increase in the pressure thereof, the iiow downstream of said shock being subsonic and confined in a passage defined by only one of two of said adjacent blades of said cascade.

2. A compressor comprising a casing structure, a compressor rotor journaled axially in said casing structure and defining therewith an annular fiow passage, blading carried by said rotor and extending across said annular flow passage in a cascade arrangement, stator vanes carried by said casing structure and extending across said flow passage downstream of said rotor blades and arranged to direct the gaseous medium to be compressed at a proper angle to the next adjacent rotor, said rotor blades being substantially airfoil in section from root to tip and havingr a substantially sharp leading edge and relatively blunt trailing edge, said blades having their point of maximum thickness in the 75% to 100% chordwise position, said blades being arranged so that the inlet edge of each blade together with the maximum thickness portion of the adjacent leading blade delines in cooperation with the leading edge of the next following blade a restricted throat in the passage between the blades, said throat being substantially aligned with the direction of flow of the gaseous medium iiowing through the blade passage for effecting a normal compression shock to gaseous medium entering said throat at superacoustic velocity relative to the moving blades and thereby provide an immediate reduction in the velocity of said medium and an immediate increase in the pressure thereof.

3. A compressor comprising a casing structure, a compresser rotor journaled axially in said casing structure and defining therewith an annular flow passage, blading carried by said rotor and Vvextending across said annular flow passage in a cascade arrangement, said rotor blades being substantially airfoil in section from root to tip and having a substantially sharp leading edge and a relatively blunt trailing edge, said blades having top and' bottom surfaces which define the low and high pressure sides of each of said blades, respectively, said blades having their point of maximum thickness adjacent said trailing edges and said point being formed on the top or low pressure cambered surface of each blade, said point being defined by a maximum departure of said top or low pressure surface from the chord line of the blade, said blades being arranged so that the inlet edge of each blade together with the maximum thickness portion of the adjacent leading blade defines a restricted throat in the passage between the blades, said top surface upstream of said throat being substantially aligned with the relative inlet flow and said throat being substantially aligned with the direction of iiow of the gaseous medium owing through the blade passage for effecting a normal compression shock to gaseous medium entering said throat at superacoustic velocity relative to the moving blades and thereby provide an immediate reduction in the velocity of said medium and an immediate increase in the pressure thereof, the flow downstream of said shock being subsonic and confined in a passage defined by only one of two of said adjacent blades of said cascade.

4. In an axial flow supersonic compressor having at least one rotor, said rotor including a plurality of circumferentially spaced blades having a predetermined angle of attack relative to the axis of rotation of said rotor, said blades in cross section having relatively lsharp leading edge portions extending a substantial portion of the chord- Wise dimension of the blades, said blades having top or low pressure and bottom or high pressure surfaces with the top surface leading in motion during rotation of the blades about said axis, the low pressure or top surface of said blades having upstream and downstream surface por tions with the upstream portion extending over a majority of the chordwise length ofthe blade and being substantially parallel to the relative direction of the inlet flow, the downstream surface portion being at an angle relative to said upstream portion, and the point of maximum thickness of the blades adjacent the outer tip thereof being located adjacent the trailing edge of said blades thereby forming a throat with the leading edge of the following blade whereby the leading edge shock extending at substantially right angle from the high pressure surface of one blade intercepts the trailing edge region of the low pressure surface of the next adjacent blade approximately inthe to 100% chordwise range.

References Cited in the file of this patent UNITED STATES PATENTS 2,721,693 Fabri et al. Oct.r25, 1955