Gas-cooled Turbine Blade

Abstract

A cooled turbine blade for gas turbines and the like having passageways for conducting a relatively cool fluid through the blade to its tip and incorporating a walled cavity at the tip of the blade. The wall of the cavity has a cutaway portion which permits the cooling fluid to discharge into a main gas stream through the turbine at the area of lowest pressure on the blade surfaces. This permits the flow of fluid from the discharge orifices of the passageways to be more uniform from the leading to the trailing edge of the blade and reduces the required gas pressure on the entire cooling system.

Description

BACKGROUND OF THE INVENTION

In constructing turbine blades for gas turbine engines, numerous difficulties are encountered in their design because of the high speeds at which such blades operate and because of the high temperatures to which they are subjected. The temperature of a blade may be held materially lower than that of the gases in which it operates by providing cooling passages in the blade extending from its root to its tip and by conducting a stream of cooling fluid, such as air, through these passageways to cool the blade.

Normally, the cooling passages terminate at openings in the extreme end or tip of the blade. This, however, has several disadvantages. First, the pressure on the blade at its leading edge is greater than at its trailing edge. As a result, the cooling fluid in the passages at the leading edge must discharge into an area of greater pressure than the passages at the trailing edge, meaning that the flow of fluid and heat transfer efficiency at the leading edge is impeded. Secondly, the tip clearance between the end of the blade and the stationary turbine wall must be greater than it otherwise would be; and the tolerances become more critical. Finally, the pressure on the cooling fluid must be increased to the point where it can be effectively discharged into a high-pressure gas stream.

SUMMARY OF THE INVENTION

In accordance with the present invention, the foregoing drawbacks associated with prior art cooling systems are overcome by a construction wherein the blade is provided with a walled cavity at its tip. The cooling passages terminate at the bottom of the cavity; while the extremity of the cavity wall is adjacent the turbine casing, and, in effect, constitutes the end of the blade. A portion of the wall is cut away on the low-pressure side of the blade adjacent its trailing edge such that the passages in the blade do not discharge into the main turbine gas stream but rather into a relatively constant pressure area within the cavity. Furthermore, by virtue of the fact that the wall of the cavity is cut away adjacent to the trailing edge of the blade on its low-pressure side, the cavity is exposed to the lowest discharge pressure possible with a minimum disturbance to the main flow through the turbine. Thus, the pressure on the cooling system can be reduced and, at the same time, the flow of fluid through the cooling passages from the leading edge of the blade to its trailing edge is essentially the same throughout.

The above and other objects and features of the invention will become apparent from the following detailed description taken in connection with the accompanying drawings which form a part of this specification.

DESCRIPTION OF THE PREFERRED EMBODIMENT

FIG. 1 is a fragmentary elevational view of a turbine rotor having the blades of the invention thereon;

FIG. 2 is a top view taken substantially along line II--II of FIG. 1, but on a larger scale, showing the walled cavity tip construction of the blades of the invention;

FIG. 3 is a cross-sectional development view through the curved center of a blade taken substantially along line III--III of FIG. 2; and

FIG. 4 is a perspective view of the blade tip of the invention.

With reference now to the drawings, and particularly to FIG. 1, there is shown, by way of example, part of a gas turbine rotor structure including a rotor shaft 10 which is connected to a rotor disc 12 projecting radially from the shaft axis. The rim 14 of the disc 12 is formed with a continuous series of transverse "side entry" recesses 16, each of which is broadly V-shaped but having sidewalls formed with adjacent secondary transverse serrations or grooves 17.

The turbine blades 18, as perhaps best shown in FIGS. 2 and 4, have the usual airfoil shape with rounded leading edges 20, thin trailing edges 22 and curved profiles. The blade root 24 (FIG. 1) is in the form of a triangular elongated block with the blade 18 integral with its one side and the two other sides having adjacent transverse grooves therein, which grooves conform to and are adapted to register with the grooves 17 formed in the transverse recesses 16 of the disc rim 14, thus forming the well-known "fir tree" base connection. Attachment is made by sliding the blade root into the disc rim recess so that the grooves and ridges of the two parts are in cooperative engagement. Suitable means, not shown, are provided for maintaining the roots 24 within their cooperating recesses 16. Furthermore, suitable means, not shown, but well within the skill of the art, are provided for conducting a cooling fluid such as air up through the root portions 24 and into radial passages extending through each of the blades 18.

The radial passageways 26 are shown in FIGS. 2, 3 and 4 and extend all the way from the root 24 to the tip or outer radial extremity of each blade. The lower sides of the blades 18 as viewed in FIG. 2 are referred to as the high-pressure sides 29 and are those sides against which hot gases under pressure are directed in order to cause the rotor to turn in the direction of arrow 30 shown in FIG. 1. The other side 32 of each blade also extends between the leading and trailing edges 20 and 22 and is referred to as the low-pressure side. Hot gases, in passing through the blades, generally follow the path of arrows 34. The point of highest pressure on the blades is at the high-pressure side 29 adjacent the leading edge 20. As the gas travels from the leading to the trailing edge of the blade, it expands and its pressure reduces. Consequently, the point of lowest pressure is at the low-pressure side 32 near the throat area B.

In accordance with the present invention, the tip of each blade 18 is provided with a walled cavity, generally indicated by the reference numeral 36 in FIGS. 2-4, which is formed by a wall 38 extending along the top of the high-pressure side 29 from the trailing edge 22 to the leading edge 20 and thence along approximately one-half the length of the low-pressure side 32 to point A which is just downstream of the throat B between the trailing edge 22 of one blade and the low-pressure side of the following blade. The radial passageways 26 terminate at the bottom 37 of the cavity 36, as perhaps best shown in FIG. 3. In this manner, the gas discharges into an area within the cavity 36 of relatively constant pressure, meaning that the pressure which gases issuing from the passageways 26 at the leading edge 20 experience is approximately the same as that experienced by the gas leaving the passageways 26 adjacent the trailing edge 22. The depth of the cavity 36 is adjusted to provide an area at C which is substantially greater than the combined area of the cross sections of the cooling passages 26 upstream of point A. This permits the discharge pressure for the upstream holes to approach the pressure at point A, approximately the lowest pressure on the blade surface.

The present invention thus provides a means for attaining a uniform flow of cooling fluid through radial passages in a turbine blade from its leading to its trailing edge, independent of operating tip clearances between the blade tip and the stationary turbine walls. At the same time, the cooling air enters the main stream through the turbine in more nearly the correct direction for mixing with the main gas stream, thereby reducing mixing losses. Although the invention has been shown in connection with a certain specific embodiment, it will be readily apparent to those skilled in the art that various changes in form and arrangement of parts may be made to suit requirements without departing from the spirit and scope of the invention.

Claims

I claim as my invention:

1. In a turbine blade of the type having a root portion, a tip portion, leading and trailing edge portions, a low-pressure surface extending between said leading and trailing edges, a high-pressure surface continuous with said low-pressure surface and extending between said leading and trailing edges, and a plurality of passages of uniform cross-sectional area extending from said root portion to said tip portion of the blade for conducting a cooling fluid through the blade; the improvement comprising a walled cavity at the outermost tip portion of said blade having a bottom at which said passages terminate and a cutaway portion in its wall extending between said trailing edge and a point on said low-pressure side which permits the cooling fluid to discharge into the main gas stream through the turbine at the point of substantially lowest pressure on the blade surfaces, and said wall being a continuous extension of the outer high-pressure and low-pressure surfaces.

2. The improvement of claim 1 wherein said point is opposite the trailing edge of a next successive turbine blade when said blades are assembled on a rotor.

3. The improvement of claim 1 wherein the area between the point on said low-pressure side where said wall terminates and the opposite wall on the high-pressure side of the blade is substantially greater than the combined areas of the cross sections of the passages upstream of said point between said point and said leading edge.

4. The improvement of claim 1 wherein said blade is of airfoil construction with a rounded leading edge portion of greater radius than the trailing edge portion.