EP0774050B1

EP0774050B1 - Interrupted circumferential groove stator structure

Info

Publication number: EP0774050B1
Application number: EP95922886A
Authority: EP
Inventors: Robert S. Mazzawy
Original assignee: United Technologies Corp
Current assignee: Raytheon Technologies Corp
Priority date: 1994-06-14
Filing date: 1995-05-24
Publication date: 1999-03-10
Anticipated expiration: 2015-05-24
Also published as: DE69508256T2; JPH10501318A; WO1995034745A1; TW268070B; JP3640396B2; EP0774050A1; DE69508256D1

Description

This invention relates to axial flow rotary machines, and more particularly to a casing extending circumferentially about a compression stage for a gas turbine engine.

An axial flow, rotary machine, such as a gas turbine engine for powering aircraft, has a compression section having a plurality of compression stages. As the working medium gases are drawn into the engine, the gases are compressed to raise the temperature and pressure of the gases. The gases are burned with fuel and expanded through a turbine to develop useful power and work. The work is transferred from the turbine section to the compression section via a rotor assembly. The rotor assembly in the compression section does work on the incoming gases to compress the gases.

One example of a compression section is the fan section of a high bypass ratio turbofan engine. The fan section may have a diameter of 150 to 200 mm (six to eight feet). A plurality of fan blades extend radially outward from the rotor assembly across the flowpath for working medium gases. The fan blades may be rotated at speeds in excess of 2,500 revolutions per minute about an axis of rotation. The rotating fan blades drive the gases rearwardly, compressing the working medium gases as the gases enter the fan section of the engine.

Each fan blade has a tip. The tip is shrouded by the non-rotating adjacent casing; the fan blades do not have a shroud at the tip that rotates with the blades. A clearance is provided between the tips and the adjacent casing to accommodate transient movement of the blade outward with respect to the casing. This region is commonly referred to as the end wall region. The transient movement may occur in response to maneuver loads and the normal growth of the blade in response to rotational forces acting on the blade.

The end wall region experiences different aerodynamic conditions than does the remainder of the compression stage. In the end wall region, the aerodynamic interaction between the tip of the rotor blade and the adjacent wall causes boundary layer effects and drag effects. These effects make it more difficult to force the flow rearwardly. As a result, the pressure rises by a smaller amount in the endwall region than in the mid-span region of the fan blade. Leakage also occurs at the end wall region between the opposite axially extending sides of the blade in response to pressure gradients across the tip of the blade. As a result, the surge margin and aerodynamic efficiency of the compression stage are adversely effected.

Many approaches have been proposed to vary the contour of the outer wall outwardly of the fan blades to maintain an adequate surge margin. An important issue is the effect that such contouring has on the aerodynamic efficiency of the compression stage.

One example of such an approach is shown in U.S. Patent No.: 4,239,452 issued to Roberts, Jr. entitled Blade Tip Shroud for a Compression Stage of a Gas Turbine Engine and which discloses an axial flow machine having the features of the preamble of claim 1. In Roberts, the casing radially outwardly of the blade has a recess. Discontinuities are disposed in the recess. The discontinuities effect the aerodynamic interaction between the tip of the blade and the wall. As shown in Figure 6 of Roberts, the discontinuities increase surge margin with some penalty in aerodynamic efficiency which results from the aerodynamic interaction with the discontinuities.

In one embodiment in Roberts, for example, the recess is provided with circumferentially extending grooves (Figure 3). The grooved construction provides an increase in surge margin over a smooth wall (Figure 6, curve C) with a decrease in aerodynamic efficiency. Alternatively, the wall might be provided with a plurality of axially extending skewed chambers (Figure 2) or a combination of both a circumferentially extending grooves and axially extending skewed chambers (Figure 4). The Figure 4 construction shows the greatest increase in surge margin with the smallest decrease in aerodynamic efficiency over a smooth wall.

The axially skewed chambers decrease efficiency because the rotor blades use some of the work transferred to them to pump and re-pump the working medium gases in the axially skewed chamber. Pumping increases the level of pressure attainable in the end wall region in response to a sudden back pressure. Thus, pumping is the mechanism by which the axially extending chambers increase surge margin.

As can be seen in Figure 6, the circumferentially grooved construction has a moderate gain in surge margin and a moderate loss in the aerodynamic efficiency in comparison to the smooth wall. Thus, if less surge margin were required, it could be accomplished with circumferentially extending grooves with less of a loss in aerodynamic efficiency than some of the other constructions illustrated in Roberts.

The above art notwithstanding, the Applicant's scientists and engineers are seeking to develop other constructions of casings which increase surge margin in comparison to smooth wall constructions while decreasing the adverse effect such constructions have on aerodynamic efficiency.

This invention is predicated upon the recognition that interrupting circumferential grooves with a critical number of dams related to the number of adjacent fan blades results in an improvement in surge margin with a negligible effect on aerodynamic efficiency.

Thus the invention is characterised over the aforementioned prior art by the features of the characterising portion of claim 1.

Thus, according to the present invention, a stator structure for the compression section of an axial flow rotary machine includes a casing which extends circumferentially about the array of rotor blades, the casing having an inner wall which includes a plurality of circumferentially extending grooves radially outwardly of and axially aligned with the blades, each of the grooves having a plurality of circumferentially spaced dams which interrupt the circumferential continuity of the groove and direct high velocity flow from the grooves into the adjacent boundary layer.

In accordance with one embodiment of the present invention, the number of circumferentiaily spaced dams lies in a range which is a function of the number of blades of the array of rotor blades; and, the circumferentially spaced grooves are disposed radially outwardly only of the mid-chord region of the rotor blade to increase the level of the effect of the grooves and dams on the boundary layer and decrease the impact on the aerodynamic efficiency.

A primary feature of the present invention is a casing radially outward of an array of rotor blades in a compression section. The casing has a plurality of circumferentially extending grooves which are each circumferentially interrupted by a plurality of dams. In one embodiment, the dams are circumferentially spaced and extend in grooves having a depth which is twice the axial width of the groove. The lands extending between adjacent grooves are half the width of the axial width of the grooves. The number of dams lies in a range between three-fourths the number of rotor blades and one and one-half times the number of rotor blades. The circumferential length of the dams is approximately equal to the width.

In one of the particular embodiments, the number of dams in each groove is equal to the number of rotor blades and the dams are equally spaced about the circumference of the groove. In another embodiment, the grooves are disposed in a groove region which extends over the mid-chord region of the blade such that the groove region and the adjacent surface of the case outwardly of the blades are at the same radial height. The adjacent surface of the case may be either recessed with respect to the inner surface of the casing or at the same radial height as the inner surface of the casing.

A primary advantage of the present invention is the increase in surge margin which results from disposing a plurality of circumferentially interrupted grooves radially outward of an array of rotor blades in the compression section of the engine. Still another advantage of the present invention is the incremental increase in surge margin for a given decrease in aerodynamic efficiency which results from use of the circumferentially interrupted grooves whether the tip of the fan blade is at the same radial height as the remainder of the casing or recessed with respect to the surface

A preferred embodiment of the present invention will now be described, by way of example only, with reference to the accompanying drawings in which:

Figure 1 is a perspective view of a turbofan gas turbine engine showing a casing which circumscribes the tips of the fan blades.

Figure 2 is a developed view (looking outwardly) of a portion of the fan casing shown in Figure 1, illustrating the relationship of circumferentially grooves which are interrupted by circumferentially spaced dams.

Figure 3 is a developed view corresponding to the view shown in Figure 2.

Figure 4 is an alternate embodiment of the construction shown in Figure 2 in which the grooves are staggered circumferentially.

Figure 5 is a side-elevation view taken along the lines 3-3 of Figure 2.

Figure 6 is a graphical representation of the change in percentage points of surge margin and aerodynamic efficiency for the construction shown in Figure 3 to the number of circumferentially spaced dams in the circumferential grooves expressed as a function of the number N of blades in the array of rotor blades.

Figure 1 is a perspective view of a turbofan gas turbine engine 10 of the type used to power commercial airliners. The engine includes a core section 12 and a fan section 14. A plurality of unshrouded fan blades 16 extend radially outward from a rotor 18. The rotor is disposed about an axis of rotation A.

Each fan blade has an airfoil 22 having a tip 24. The tip does not have a rotating shroud, but may have a part span shroud. The fan blade has a platform 26 which adapts the fan blade to engage an attachment slot in the rotor assembly. The length of the fan blade may exceed 0.61 m (two feet) and in the particular embodiment shown, is approximately 0.69 m (twenty-seven (27) inches long) as measured from the platform. The outer diameter of the tip of the fan blade is about 2.39 m (ninety-four (94) inches) as measured from the axis of rotation A.

The fan (compression) section 14 includes a primary flow path 28 for working medium gases which extends through the core section. A secondary flow path 32 for working medium gases extends through the fan section outwardly of the primary flowpath. The secondary or bypass flow path is annular in shape. A fan casing 34 extends circumferentially about the flow path to bound the flow path at its outermost portion. The casing has an inner surface 36, formed of an abradable material which extends circumferentially about the rotor blades. The inner surface is spaced radially from the rotor blades leaving a clearance gap G therebetween. The clearance gap in the present embodiment is approximately 2.29 mm (ninety-thousandths (0.090) of an inch).

Figure 2 is an undeveloped view and Figure 3 is a developed view of the casing 34 looking outwardly toward the inner surface 36 of the casing A plurality of grooves 38 extends circumferentially about the interior of the casing. A plurality of dams 42 are disposed in each groove. Each dam extends axially across the groove and radially in the groove to circumferentially interrupt the groove. Each dam is spaced from the adjacent dam by a pitch distance P. The pitch distance is the distance from a point on one dam to the corresponding point on an adjacent dam. The pitch distance P is equal to the circumference of the inner surface divided by the number of dams. The number of dams is equal, for example, to the number of blades N of the array of rotor blades. As will be seen in Figure 6, there is an optimum design range for the number of dams. (The design range is a function of the number of rotor blades.)

Two rotor blade tips 24, each associated with an adjacent fan blade 16, are shown in Figure 2 in full and in Figure 3 in phantom. The rotor blades sweep upwardly (Figure 2) and from right to left (Figure 3) under operative conditions as shown by the arrow V which indicates the direction of motion of the fan blades. Each fan blade has a leading edge 44 and a trailing edge 46. A suction sidewall 48 on one side of the blade extends from the leading edge to the trailing edge. A pressure sidewall 52 on the other side of the blade extends from the leading edge to the trailing edge. The fan blade has an axial length L as measured parallel to the axis of rotation. The fan blade has a leading edge region 54 which extends for the first twenty percent (20%) of the axial length L of the blade and a trailing edge region 56 which extends for the last twenty percent (20%) of the axial length L of the fan blade. The fan blade has a mid-chord region 58 which extends between the leading edge region and the trailing edge region. The casing has a groove region 62 which extends circumferentially at a location which is radially outwardly of the mid-chord region of the blade and contains the grooves.

Figure 4 is a view corresponding to the view shown in Figure 3 of an alternate embodiment of the casing shown in Figure 3. In the alternate embodiment, the circumferential grooves 38 are interrupted by dams 42 which are staggered circumferentially with respect to the adjacent dam.

Figure 5 is an enlarged, simplified cross-sectional view of the outer casing 36 and fan blade tip 24 shown in Figure 1, Figure 2, and Figure 3. The inner surface 36 of the casing outwardly of the fan blade tip is at the same radial location as the adjacent structure. As will be realized, in response to operative forces, the fan blade 16 may rub against the inner (abradable seal) surface 36, cutting a recess 64 which would result in a relationship to the inner surface as shown by the phantom lines. As shown with the phantom lines, after the tip has cut into the inner (abradable seal) surface, the surface 36 outwardly of the fan blade tip is at a different radial location than the inner surface of the adjacent structure. In alternate constructions, the recess may be deliberately formed in the outer case as shown, for example, in U.S. Patent No.: 4,239,452 discussed above.

Figure 5 shows the plurality of grooves 38. As shown in Figure 2, Figure 3, and Figure 5, each groove 38 has an axial width W , and a radial depth D as measured from the inner surface of the casing outwardly of the rotor blade. The ratio of the radial depth D to the axial width W is greater than or equal to one. In the embodiment shown, the ratio of the radial depth D to the axial width W is equal to two ( D / W = 2). Each groove is spaced axially from the adjacent groove by a distance equal to one half of the axial width W , leaving a land 66 therebetween. As discussed above, a plurality of dams 42 at each groove extend axially from the land to circumferentially interrupt the full depth of the groove. The circumferential length L_C of the dam is equal to the axial width W of the groove 38. Experience indicates that no adverse effects occur for dams which extend up to three times the width W of the groove as shown in Figure 4. It is possible to have the dam not extend for the full radial depth of the groove, but this will compromise the performance of the embodiment to a serious extent.

Figure 6 shows the relationship of the change in aerodynamic efficiency (curve E ) and change in surge margin (curve S ) for the interrupted groove 38 construction as a function of the number of dams (interruptions) 42 and pitch P . The number of dams is expressed in terms of the number of fan blades N in the array of fan blades. The changes are measured in comparison to a base line construction which does not have dams that interrupt the circumferential continuity of each groove. The relationships shown in Figure 6 are based on an analysis of empirical data.

As shown in Figure 6, the addition of dams 42 increases the surge margin with little or no decrease in the aerodynamic efficiency. For example, in embodiments that have a quantity of interruptions or dams that is equal to three-fourths the number of fan rotor blades, the surge margin increases by approximately four points with no apparent decrease in aerodynamic efficiency. As the quantity of dams increases, there is a slight decrease in aerodynamic efficiency until the quantity of dams is about equal to one and one-half times the number of rotor blades. Beyond that point, the surge margin does not measurably increase and aerodynamic efficiency continues to decrease. Accordingly, the optimum design range for the quantity of dams is believed to be equal to three-fourths to one and one-half times the number of rotor blades. The pitch P (spacing) between the dams for this range increases from about two-thirds of the circumference of the inner wall divided by the number N of rotor blades to one and one-third times the circumference of the inner wall divided by the number N of rotor blades.

During operation of the gas turbine engine, shown in Figure 1, working medium gases are drawn into the secondary flowpath 32 and compressed by the fan blades 16. As the fan blades rotate at high velocity about the axis of rotation A , leakage effects occur at the endwall region 68 between the tip 24 of each rotor blade and the inner case 36. These leakage effects are, in part, due to flow in the direction P-S from the pressure sidewall to the suction sidewall over the tip. The circumferential grooves channel the leakage in the tangential or circumferential direction (as opposed to the direction P-S perpendicular to a line connecting the leading edge to the trailing edge). This reduces the component of velocity of the leakage flow acting opposite to the direction in which the flow is driven by the fan in the downstream direction. This increases the surge margin of the embodiment and as shown in Figure 6, has a small impact on the aerodynamic efficiency of the compression stage.

Although the aerodynamic interactions in the end wall region are not well understood, the following is a working hypothesis which explains the empirical results. During operation of the engine, the rotating blades transfer energy to the adjacent gasses. In the end wall region 68, a boundary layer develops. The boundary layer has a much lower velocity than the wheel speed of the adjacent fan blade because of drag effects. In the grooves 36, however, the velocity of the working medium gases is much higher. Accordingly, high velocity working medium gases are circumferentially traversing the array of fan blades 14 outwardly of the boundary layer of relatively low velocity gases.

The low velocity gases in the boundary layer can cause significant losses because the boundary layer may separate from the wall 36. High velocity gases in the circumferentially extending grooves 38 are driven inwardly by the dams into the low energy boundary layer. These high-energy gases act to energize the boundary layer and avoid separation of the boundary layer. This avoids concomitantly the losses associated with separation of the boundary layer and results in an aerodynamic efficiency which remains constant even as the circumferential grooves provide an improvement in surge margin. Accordingly, adding the circumferentially spaced dams to the circumferential grooves allows the surge margin to increase without an unacceptable decrease in aerodynamic efficiency of the array provided the quantity of dams equals a number of fan blades within the optimum design range.

For example, the high velocities in the grooves are made possible by the pitch P of the dams which provides sufficient length of uninterrupted groove for the velocities to develop. In embodiments having too many dams, the high velocities never develop. In embodiments having too few dams, the high velocities develop but are not injected with desirable frequency into the boundary layer.

Claims

An axial flow rotary machine having an axis of rotation (A), a compression section (14) which includes an annular flowpath (28) for working medium gases, a rotor assembly which includes an array of rotor blades (16) which extend radially outwardly across the flowpath, each rotor blade having a tip (24), and a stator structure which comprises: a casing (34) having an inner surface (36), the casing extending circumferentially about the flowpath (28) for working medium gases to bound the flowpath, extending circumferentially outwardly of the rotor blades (16) and being spaced radially from the rotor blades leaving a clearance gap (G) therebetween, the casing (34) having a groove region which is radially outwardly of each blade, the groove region having a plurality of grooves (38),each groove extending circumferentially about the interior of the casing, and each groove being spaced axially from the adjacent groove leaving a land (66) therebetween, at least one of said grooves having an axial width (W) a radial depth (D) as measured from the inner surface,
characterised in that said at least one groove has a ratio of radial depth (D) to axial width (W) which is greater than or equal to one (D/W≥1), and has a plurality of dams (42) disposed in said groove, the number of dams being at least equal in number to three fourths of the number of blades (N) (0.75N) and extending axially from the land (66) to circumferentially interrupt the groove, and in that said inner surface (36) of said casing (34) is formed of an abradable material.
The machine as claimed in claim 1 wherein the number of dams (42) is equal to a number which lies in a range of three fourths of the number of blades ( N ) (0.75 N ) to one and one half times the number of blades ( N ) (1.5 N ).
The machine as claimed in claim 1 or 2 wherein each dam (42) is spaced circumferentially from the adjacent dam by a pitch distance ( P ) equal to the circumference of the inner surface divided by ( N ).
The machine as claimed in any preceding claim wherein the number of dams (42) is equal to the number of blades ( N ).
The machine as claimed in any preceding claim wherein said groove (38) has a ratio of radial depth ( D ) to axial width ( W ) which is equal to two.
The machine as claimed in claim 5 wherein the axial width of the land (66) is equal to one-half of the axial width ( W ) of the groove (38).
The machine as claimed in any preceding claim wherein at least five of the grooves have a plurality of dams (42).
The machine as claimed in any preceding claim wherein at least two of the grooves (38) have a plurality of dams (42) and the dams of one groove are staggered circumferentially With respect to the other groove.
The machine as claimed in any preceding claim wherein said casing (34) has a recess (64) in the casing which faces the rotor blades and wherein the groove region is disposed in said casing.
The machine as claimed in any preceding claim wherein the tip (24) of the rotor blade (16) has a mid-chord region which extends for sixty percent (60%) of the lenght ( L ) of the tip in the axial direction, and which is spaced by a distance equal to twenty percent (20%) of the length ( L ) from the leading edge (44) and from the trailing edge (46) of the blade and wherein the groove region is radially outwardly of the mid-chord region of each blade