US3014640A

US3014640A - Axial flow compressor

Info

Publication number: US3014640A
Application number: US740875A
Authority: US
Inventors: Joseph N Barney; Leonard B Lisher
Original assignee: Motors Liquidation Co
Current assignee: Motors Liquidation Co
Priority date: 1958-06-09
Filing date: 1958-06-09
Publication date: 1961-12-26
Anticipated expiration: 1978-12-26

Description

Dec. 26, 1961 Filed June 9, 1958 4 Sheets-Sheet 1 INVENTORS Dec. 26, 1961 J. N. BARNEY ETAL 3,014,640

AXIAL FLOW COMPRESSOR Filed June 9, 1958 4 Sheets-Sheet 2 INVENTORS 5 ATTORNEY Dec. 26, 196 1 J. N. BARNEY ETAL 3,014,640

AXIAL FLOW COMPRESSOR Filed June 9, 1958 4 Sheets-Sheet 3 co nsereo me Fzaw EFFIC7NCY ALTITUDE L033 COEFFICIENT REYNOLDS NUMBER ATTORNEY Dec. 26, 1961 J. N. BARNEY ETAL 3,014,640 AXIAL FLOW COMPRESSOR Filed June 9, 1958 4 Sheets-Sheet 4 M4! STREAM FLOW a 31 STREAM 10W 7? w H x %g 1 5 X;

v A INVENTORS 222$2 Our invention relates to turbocompressors, and is directed to improving the performance of such machines, particularly axial-flow compressors.

The invention relates particularly to improving the characteristics of compressors by controlling boundary layer effects to improve flow characteristics and efiiciency.

The nature of the invention and its advantages will be clear to those skilled in the art from the succeeding detailed description of the preferred embodiment of the invention and the accompanying drawings thereof.

FIGURE 1 is a partial sectional view of an axial-flow compressor taken on a. plane containing the axis of the compressor.

FIGURE 2 is an axonometric view of a rotor blade.

FIGURE 3 is an elevation View of the base portion of a rotor blade.

v FIGURE 4 is a cross-sectional view of a rotor blade taken on the plane indicated by the line 4-4 in FIG- URE 1.

FIGURE 5 is a fragmentary sectional view taken on the plane indicated by the line 5-5 in FIGURE 3.

FIGURE 6 is a cross-section of a stator blade taken on the plane indicated by the line 6--@ in FIGURE 1.

FIGURE 7 is a plot illustrating variation with altitude of corrected air flow and compressor efficiency in a typical axial-flow compressor.

FIGURE 8 is a plot illustrating typical variation of compressor loss coeflicient with Reynolds number.

FIGURE 9 is a plot illustrating the effect of the invention on compressor efficiency.

FIGURE 10 is a cross-section of a prior art blade, either rotor or stator.

FIGURE llis a similar cross-section of a blade embodying one feature of the invention. 7

FIGURE 12 is an elevation view of the low pressure face of a prior art rotor blade.

FIGURE 13 is a partial sectional view of the same taken on the plane indicated by the line 1313 of FIG- URE 12.

FIGURE 14 is a view similar to FIGURE 12 illustrating the effect of a second feature of the invention.

FIGURE 15 is a partial sectional View of the same taken on the plane indicated by the line l5--15 in FIG- URE l4.

FIGURE 1 illustrates an axial-flow compressor of known type incorporating this invention. The compressor stator comprises an inlet structure 11, a case 12, inlet guide vanes 13, and rows of stator blades 14.

The inlet structure mounts a bearing 17 supporting the forward end of a rotor comprising disks 18, 19, and a tie bolt 21. Rotor blades 22 are mounted on the disks. Only the first two stages are illustrated; as manymay be provided'as are desired.

FIGURES 2 to 5 illustrate a typical rotor blade, such as a first stage blade 22. It comprises an airfoil blade portion 26' and a dovetail root 27 for mounting in disk.

13. The blade portion has a curved mean. camber line, a convex low pressure face 28, and a concave or high pressure face 29. As will be seen from FIGURE 6,

. the stator blades 14also are airfoils with a curved mean camber, line, a convex low pressureiface 28' and a concave high pressure face 29. v "j v p The physical structure so far described is well known,

and is described more fully, in US. Patents 2,640,679 and ted States Patent ice 2,675,174. Such compressor structures may be improved by the present invention, which comprises two features, one applicable to both rotor and stator blades, the other applicable to rotor blades only.

Before proceeding to the description of the compressor structure as modified by the invention, reference will be made to FIGURES 10 to 15, which are generalized views illustrating the nature of the build-up of boundary layer air on the compressor blades and the effect of modification of the blades according to the invention to prevent such build-up.

FIGURES l0 and 11 are cross-sections of blades, which may be either stator or rotor blades. The blade B has a leading edge E, a trailing edge T, a low pressure face L and a high pressure face H. In the prior art blade of FIGURE 10, the spiral arrows over the rear part of the low pressure face indicate the build-up of a thick layer of low energy viscous flow which is characteristic of operation at lower Reynolds numbers. FIGURE 11 illustrates a blade upon which very small ridges 34 extending spanwise of the blade have been distributed chordwise of the blade on the low pressure face. The preferred structure of such an addition to the blade will be described in detail. FIGURE 11 illustrates by the curledarrows how the spanwise ribs deflect the boundary layer air away from the surface of the blade so that the high energy main stream air flow will scour it from the blade and carry it away. Since the boundary layer air is scoured away at several locations chordwise of the blade, it vdoes not build up into the thick layer. illustrated in FIGURE 10.

So far as the rotor blades alone are concerned, there is a second phenomenon which tends to increase the buildup of the stagnant boundary layer. air toward the trailing edge of the low pressure face of the blade, illustrated. by FIGURES 12 and 13 showing a prior art rotor blade. The blade is identified by B and the faces and edges by the same legends as in FIGURES l0 and 11. Theblade is shown as mounted in a rotor disk 18 in FIGURE 13. There will be a certain amount of stagnant boundary layer air adjacent the surface of the rotor 18. This will be drawn into the low pressure area principally toward the trailing edge of the blade where it adds to the boundary layer air present on the face of the blade. Because of the rotation of the rotor there is a centrifugal force which tends to pump this air outwardly along the blade towards the tip. However, this action does not diminishthe thicknessof the boundary layer air, but merely moves it spanwise. The separation of the main air flow from the blade is increased by the addition of rotor boundary layer air to the blade boundary layer.

FIGURES 14 and 15 illustrate the second feature of the invention, the provision of a rib 31 extending from the low pressure face of the blade and preferably from the leading edge to the trailing edge as illustrated by the. arrows in these figures. The rotor boundary layer air after a short initial travel along the surface of the blade strikes the rib 31, and is deflected outwardly into the fast: moving main air stream which carries it away.

The general nature of the invention being clear from the preceeding general description with reference to-FIG- URES 10 to 15, we may now proceed to the more specific description of an actual installation with reference to FIGURES 1 to 7. i

The feature of the invention applicable only to the rotor blades is provided by a rib 31 extending substantially through the compressor. The preferred configuration of 3 this chrodwise-extending rib is most clearly apparent from FIGURES 2 and 4.

The purpose of the rib is to detach boundary layer air from the low pressure surface of the blade, particularly near the trailing edge. As is known, a more or less stagnant boundary layer develops at the margin of the air flow path along the surface of the rotor defined by the disks 18, 19, and the upper surfaces of the blade roots 27. This stagnant air rotates with the rotor. A low pressure area develops on the convex side of the rotor blades. There also a boundary layer develops, and grows toward the trailing edge of the blade. The stagnant air at the rotor surface tends to move outward along the face of the blade into this low pressure area, influenced by centrifugal force, and augment the boundary layer already present. Because of the thick boundary layer, it is not swept off by the air flowing axially through the compressor. The addition of the stagnant air from the inner boundary of the path, augmenting the blade boundary layer air, enhances flow separation from the blade and reduces efficiency.

The rib 31 intercepts this outwardly-flowing stagnant air and deflects it into the fast-moving main air stream between adjacent blades, which sweeps it away. Elimination of this extraneous stagnant air from the greater part of the blade span improves the stage efficiency.

The second feature of the invention lies in means to increase the turbulence of the boundary layer air on the low pressure faces of the blades of both rotor and stator. This aspect of the invention is particularly applicable to compressors of moderate size, rather than large compressors, since the Reynolds number is less in small compressors.

The usual practice in compressors is to polish or otherswie finish the blades so that they have a smooth surface. This has its advantages, but it has the disadvantage that below a certain value of Reynolds number, which may be approximately 90,000, the boundary layer on the low pressure face of the blade builds up to a considerable thickness, and there is thus an increase in separation of flow from the low pressure face of the blade. In a typical compressor used in an aircraft engine, for example, the etficiency of the compressor decreases at higher altitudes because of the decrease in Reynolds number. The Reynolds number may vary from about 200,000 at sea level to about 50,000 at 40,000 feet altitude, because of decreasing air density. The purpose of the present invention is to impart turbulence to the otherwise laminar boundary layer, which will minimize the growth of the boundary layer and separation of the flow from the blade.

This result is obtained by providing a rough surface on the blade which will impart turbulence to the boundary layer without affecting too adversely the air flow over the blade.

The preferred structure for this purpose comprises a number of very small ridges 34 extending spanwise of the blade. As will be apparent from FIGURES 4 and 6, these ridges are quite small compared to the span of the blade and are of a triangular cross section. It is believed that the ridges may preferably be located approximately at the 7 and 3 chord values measured from the leading edge. As an indication of the size of the ridges, it is contemplated that they may be approximately $1 of an inch high.

The effect of these ridges on compressor performance may be explained by reference to FIGURES 7 to 9..

FIGURE 7 illustrates the effect of altitude on the performance of a compressor with smooth blades. The variation in corrected air flow is a function of corrected rotor speed. However, because a turbo-prop engine generally operates at a nearly constant rotor speed and the airplane Mach number is fairly low, a generalized curve such as FIGURE 7 will aid in the explanation. Considering the air at sea level as standard, the corrected air flow increases with altitude. The corrected air flow is the weight of air flowing through the engine in unit time multiplied by the square root of the ratio of ambient temperature to standard sea level temperature and divided by the ratio of ambient air pressure to standard pressure at sea level.

The efliciency of the compressor decreases very slowly with altitude from its sea level value up to a point at which the curve drops more steeply with increase in altitude. If the compressor is used in a gas turbine aircraft engine, the normal flight altitude may lie beyond the droop in the efficiency curve of FIGURE 7. One important reason for this loss of efliciency is the decrease in Reynolds number of the air flowing through the compressor with altitude. This is illustrated by FIGURE 8, in which the loss coeflicient which represents loss of power in the compressor is plotted as a function of Reynolds number. At higher values of Reynolds number, the loss coeflicient is nearly constant, but as the Reynolds number decreases the curve begins to rise rather steeply. This point, which is indicated as the break point, will occur approximately at a Reynolds number of 90,000. The low Reynolds numbers are those occurring at high altitude and the point on the curve indicated as altitude may represent the highest normal flight altitude of an airplane incorporating the compressor. Gas turbine powered aircraft fly more efliciently at high altitudes than low and, therefore, are ordinarily flown at high altitudes such as 25,000 feet or more. This loss of compressor efficiency, which reduces the efiiciency of the engine and, therefore, increases its fuel consumption for a given power output, is quite undesirable.

The present invention is directed to improving the compressor efliciency. or reducing the compressor loss coeflicient, and thereby increasing engine efliciency at high altitudes, by setting up turbulence in the boundary layer on the low pressure face of the blades. This has an effect equivalent to an increase of Reynolds number. While the resulting disturbance of flow will decrease engine efiiciency at low altitude where the Reynolds number is high, it will tend to shift the break point to lower values of Reynolds number and thereby increase high altitude eflicieny. This is illustrated by FIGURE 9, in which the broken line represents the variation of efficiency of a compressor with smooth blades as a function of Reynolds number. The solid line represents the variation of efficiency with the ridged blades of the invention. It will be noted that a compressor incorporating the invention is somewhat less efficient at the higher Reynolds numbers encountered at low altitude. However, it is more efficient over a considerable range of high altitude flight in which a gas turbine aircraft normally operates. The result is that the overall fuel consumption of the aircraft is considerably improved.

It will be apparent that, so far as the rotor blades are concerned, the rib 31 and the ridges 34 cooperate in eliminating or preventing the development of thick, stagnant boundary layers on the low pressure face of the rotor blades and thus cooperate to minimize separation of flow from the blade and promote efliciency of the compressor.

We claim:

A dynamic compressor comprising a stator and a rotor, at least one compressor stage constituted by a row of rotor blades mounted on the rotor and a row of stator blades mounted on the stator, the said blades being of generally airfoil section and having a low pressure face, and ridges on said blades extending spanwise on the low pressure face thereof adapted to impart turbulence to the boundary layer air thereon, the low pressure faces of the blades being smooth except as broken by the ridges, the ridges being spaced chordwise of the blades, and at least some of the said ridges being adjacent the mid-chord of each blade, each rotor blade bearing a rib extending chordwise thereof adjacent to the rotor on the low pressure face thereof adapted to intercept boundary layer air flowing from the rotor spanwise of the blade and thereby prevent blanketing of the ridges by the said spanwise flowing boundary layer air.

Clark June 28, 1932 Telfer Jan. 16, 1934 FOREIGN PATENTS Great Britain Aug. 16, 1938 Great Britain Sept. 20, 1946 Great Britain June 13, 1956 Great Britain Apr. 9, 1958 France July 18, 1938 France July 8, 1957