GB2092681A

GB2092681A - Circumferentially Grooved Turbine Shroud

Info

Publication number: GB2092681A
Application number: GB8133245A
Authority: GB
Original assignee: Pratt and Whitney Aircraft of Canada Ltd
Current assignee: Pratt and Whitney Canada Corp
Priority date: 1981-01-27
Filing date: 1981-11-04
Publication date: 1982-08-18
Also published as: CA1158563A; GB2092681B; FR2498679B2; FR2498679A2

Abstract

The stationary shroud 22 includes a plurality of spaced-apart lands 26 defining therebetween circumferential grooves 24 to reduce the flow leakage past the tips of rotor blades 12. The grooves are inclined at a specified angle ??? to the radial plane and the maximum thickness of the blade is greater than the depth d of the grooves. <IMAGE>

Description

SPECIFICATION Circumferentially Grooved Shroud Liner The present invention relates to.gas turbine engines, and is an improvement on the invention described and claimed in British Patent Application Serial No. 2017228.

Rotor blades in a turbine, especially in smaller turbines, are normally unshrouded. The root of the blade is fixed to a hub, but the outer radial end or tip of the turbine rotor blade is free. A liner of cylindrical shape is normally provided in the stationary housing acting as a shroud for the blades.

However, since the shroud liner is stationary, it is necessary to leave a gap or tolerance between the tip of the blade and the shroud liner so as to avoid rubbing. It is necessary to provide a sufficient gap between the tip and the shroud in order to provide for the differences in the expansion of the respective metal components. A minimum practical clearance has been found to be 1% of the blade height.

However, when such gaps are provided, the gases on the high pressure side of the blade tend to leak over the tip of the blade at a relatively higher velocity than the rotating velocity of the blade to thereby interfere with the low pressure side of the blade and deteriorate the flow pattern of the gases on the low pressure side increasing, for instance, separation of the gas flow from the surface of the blade on the low pressure side thereof. Furthermore, the gap causes some of the gas to bypass the rotor, and thus not contribute to the work, since it is not turning relative to the rotor blades.

Various improvements and developments have been made to reduce the actual gap between the tip of the blade and the shroud liner as, for example, the ablative seal described in U.S. Patent 3,836,156, issued September 17, 1974, Hector B. Dunthorne, and assigned to the applicant. This helps to reduce the size of the gap, but does not eliminate the tip leakage. One common method for reducing the amount of tip leakage flow, as well as its disruptive influence on the flow on blade low pressure surface, is to equip the rotor tip with a shroud.

Each blade has at its tip a segment of a ring, such that when assembled in a rotor disc, these shroud segments form a continuous ring which prevents flow from within rotor blade passages from leaking around the blade tips. Due to the necessary tip clearance, some flow will still leak past the rotor blades, but at least it does not disrupt the mainstream flow on blade low pressure surfaces.

However, any such shroud causes a relatively large amount of metal to be added to rotor blade tips. This is most undesirable in the case of first stage turbine blades, because of the high gas temperature at the exit of the combustion chamber immediately preceding the turbine. The additional mass of the blade created at the tip causes the centrifugal stresses in the rotor blades to be substantially increased, with the result of a much reduced rotor blade life; while there are ways to alleviate this problem, such as a reduced gas temperature, or blade roots with very large metal areas, hence thick discs, considerations of overall engine efficiency and weight usually dictate the elimination of such shrouds on first stage blades.

It is an aim of the present invention to provide a mechanical means of aerodynamically controlling and reducing the leakage of the gases from the high pressure side of the blade to the low pressure side of the blade over the tip thereof without using shrouds at the blade tips.

British Patent Application Serial No. 2017228 defines a gas turbine engine having a rotor provided with a plurality of radially extending blades and a stationary shroud surrounding the rotor, each blade having a high pressure side and a low pressure side relative to the fluid flow and a generally smooth tip end surface, wherein the stationary shroud comprises in its radially inward facing surface a plurality of lands providing spaced-apart grooves in the surface formed so as to reduce leakage of working fluid from the high pressure side to the lower pressure side of the rotor, the thickness of each land, the width of the grooves and the depth of the grooves being selected as a function of the flow characteristics of the turbine so as to effectively thicken the boundary gas layer between the blade tips and the shroud.

It has now been discovered that an improvement in engine efficiency can be effected by controlling the inclination and depth of the grooves.

According to the present invention there is provided a gas turbine engine having a rotor provided with a plurality of radially extending blades and a stationary shroud surrounding the rotor, each blade having a high pressure side and a low pressure side relative to the fluid flow and a generally smooth tip end surface, wherein the stationary shroud comprises in its radially inward facing surface a plurality of spaced-apart lands providing parallel circumferentially extending grooves in the surface formed so as to reduce leakage of working fluid from the high pressure side to the lower pressure side of the rotor, the thickness of each land, the width of the grooves and the depth of the groove being selected as a function of the flow characteristics of the turbine, wherein the parallel grooves are inclined relative to a radial plane towards the fluid flow at an angle from 100 to 400, and wherein t/d > 1 where t is the width dimension of the widest part of the blade at the tip and d is the depth of the grooves.

An embodiment of the invention is hereafter described with reference to the accompanying drawings, in which: Figure 1 is a fragmentary perspective view of a turbine wheel with a stationary shroud associated therewith; Figure 2 is a fragmentary radial cross-section through the embodiment of Figure 1; Figure 3 is a cross-section of a turbine blade, as in Figure 4; Figure 4 is a graph of engine efficiency y against t/d; and Figure 5 is a graph of engine efficiency y against the inclination angle of the grooves.

Referrin'g now to the drawings, there is shown a turbine wheel 10 with a plurality of blades 12 each having a high pressure or concave side 14 and a low pressure or convex side 16. Each of the blades is fixed to a hub 18. Each blade has a free end referred to as tip 20. A cylindrical housing surrounding the rotor includes a cylindrical stationary shroud 22. The shroud 22 is spaced from the tips 20 of the blades 1 6 by the gap (g). This distance (g) represents the necessary clearance to allow for differences in expansion of the respective m4.tals.

The shroud, in the present embodiment, is provided with a plurality of parallel grooves 24 having a bight 28 and separated by thin lands 26. In a typical example of a turbine which was tested, the parameters were as follows: Blade chord length .540" Thickness of blade trailing edge .018" Gap between blade and shroud (g) .014" to .020" Depth of groove 24 (d) .1 50" Width of groove (b) .050" Thickness of land 26 (a) .020" Mean angle of blade to axis of rotor 45.3 The groove parameters would, as a function, be determined by the size and shape of the blades as well as the gap (g).

It has been found that, in addition to the advantages mentioned above, there is the added advantage that there is less rub area for the blade tips and the geometry of the shroud is better suited for blade containment in the case of accidental dislodgement of the blades.

In the embodiment shown in Figures 1, 2 and 3, the bight 28 of the groove 24 is semi-circular in cross-section, and the grooves 24 run circumferentially and obliquely to the radial plane.

As discussed in Application Serial No. 2017228, next to the stationary shroud 22, the primary stream of hot gas forms a boundary layer. The moving blades 12 continuously cut into this slower moving boundary layer, which has the effect of opposing the tip leakage flow, in effect forming a partial aerodynamic seal. This phenomenon occurs on all unshrouded blades.

The grooves formed by lands 26 have the effect of thickening this boundary layer, improving the effectiveness of this "seal" (by virtue of the larger.- surface area scrubbed by the gas in the tip region), and by directing this boundary layer more exactly in the direction of the prevailing tip leakage.

This direction, however, will not be the same for all possible turbine rotor blades, but will depend on the inlet and exit angles of the blade.

It has been discovered through tests that there appears to be a relationship between the thickness of the blade tip t at the widest part of the blade and the depth of the grooves. The efficiency of the engine will be increased if the ratio of the thickness of the blade to the depth of the groove in the shroud is greater than 1. In other words, t/d > l. Figure 4 shows graphically this phenomenon. In the case of very thin blade tips, small caps or mini-shrouds could be provided at the tip thereof in order to effectively widen the tip of the blade. The following results were taken from actual tests, and the efficiency is measured against an engine having a shroud without grooves but with the remaining characteristics the same.

Blade Tip Groove Groove Thickness (t) Depth (d) Width (b) (inches) (inches) (inches) t/w tld y lSy y Example .171 .150 .050 3.42 1.14 90.9 +.1 Example II .073 .150 .050 1.46 .49 90.2 -.4 Example Ill .073 .060 .050 1.46 1.22 90.7 +.1 y=efficiency hy=difference in efficiency The tests shown above, it is noted, were taken with a groove width being constant, that is, at .050 inches in width.

In the embodiments illustrated the grooves are distributed over an axial extent equal to the chord length of the turbine blade.

It has been found that with turbine rotor blades, separation of flow from the low pressure surface is generally present, near the tip of the blade. This separation is caused by the flow leakage over the blade tip, as discussed. Tip sections of different turbine blade design have different surface pressure distribution, and thus different positions of flow separation when tip leakage flow is present. Since the grooves redirect and retard the tip leakage flow, the optimum location (and direction, for that matter) of the grooves is likely to differ from one blade design to the next. For example, in the case of a blade tip having a large negative pressure gradient nears it trailing edge, grooves are distributed over the trailing edge area of the blades, to delay separation.In the case of a blade having a negative pressure gradient near its leading edge, the optimum placement of grooves would be over the area of the leading edges of the blades.

In the embodiment illustrated the grooves 24 are inclined at an angle to the radial plane, and this angle cP can be from 100 to 400 from the radial plane x. The angle of inclined grooves 24 (or lands 26) provides a further control of the flow in the grooves, that is, a changed flow co-efficient.

Figure 5 shows the increase in efficiency of the engine as the angle of inclination from the radial plane x increases in a direction upstream of the flow. In specific tests, a=0.020" b=0.050" d=0.1 50" =20 to 30 As can be seen from Figure 5, the preferred angles of inclination are between 20 and 300 although improved results are obtainable with angles of inclination from 100 to 400.

Schedule of Metrix Equivalents .014 ins .356 .356 mm .018 ins .457 .457 mm .020 ins ------ .508 mm .050 ins 1.270 mm .060 ins 1.524 mm .073 ins 1.854 mm .150 ins 3.810 mm .171 ins ----- 4.343 mm .540 ins ----- 13.716 mm

Claims

1. A gas turbine engine having a rotor provided with a plurality of radially extending blades and a stationary shroud surrounding the rotor, each blade having a high pressure side and a low pressure side relative to the fluid flow and a generally smooth tip end surface, wherein the stationary shroud comprises in its radially inward facing surface a plurality of spaced-apart lands providing parallel circumferentially extending grooves in the surface formed so as to reduce leakage of working fluid from the high pressure side to the lower pressure side of the rotor, the thickness of each land, the width of the grooves and the depth of the groove being selected as a function of the flow characteristics of the turbine, wherein the parallel grooves are inclined relative to a radial plane towards the fluid flow at an angle from 100 to 400, and wherein t/d > 1 where t is the width dimension of the widest part of the blade at the tip and d is the depth of the grooves.

2. A turbine engine as claimed in claim 1 , wherein the angle of inclination is between 200 and 300.

3. A turbine engine as claimed in claim 1 or claim 2, wherein the grooves have a depth from 0.050 inch (1.27 mm) to 0.150 inch (3.81 mm).

4. A turbine engine as claimed in any preceding claim, wherein there are not less than four grooves distributed over the chord length of the tip.

5. A turbine engine as claimed in any preceding claim, wherein the grooves are distributed over an axial extent equal to the chord length of the turbine blade.

6. A turbine engine as claimed in any of claims 1 to 4, wherein the grooves are distributed over an axial extent covering a fraction of the chord length of the turbine blade.

7. A turbine engine as claimed in claim 6, wherein the grooves are distributed over the area of the leading edge of the turbine blades.

8. A turbine engine as claimed in claim 6, wherein the grooves are distributed over the trailing edge of the turbine blades.

9. A turbine engine substantially as described herein with reference to the accompanying drawings. ~~~~~~~~ ~~~~~