US2966331A

US2966331A - Hollow, air cooled turbine blades

Info

Publication number: US2966331A
Application number: US720774A
Authority: US
Inventors: Creek John Oliver
Original assignee: Orenda Engines Ltd
Current assignee: Orenda Engines Ltd
Priority date: 1958-03-11
Filing date: 1958-03-11
Publication date: 1960-12-27
Anticipated expiration: 1977-12-27

Description

Dec. 27, 1960 J. o. CREEK y 2,956,331

HOLLOW, AIR cooLED TURBINE BLADES Filed March 1l, 1958 I4 II is l l l l l 1 1 l '6,63 '5 FIG. 2

lSd Il INVENTOR J. O. CREEK United States Patent 2,965,331 HOLLOW, AIR COOLED TURBINE BLADES John Oliver Creek, Brampton, to Orenda Engines Limited, a corporation of Canada Filed Mar. 11, 1958, Ser. No. '720,774 3 Claims. (Cl.l 253--39.15)

Ontario, Canada, assigner Malton, Ontario, Canada,

This invention relates to gas turbine blades and more particularly to gas turbine blades having cooling passages formed integrally with a sheet metal skin.

In constructing turbine blades for gas turbine engines, numerous difiiculties are encountered in the design because of the high speeds at which such blades operate and because of the high temperatures to which they are subjected. VThe temperature of the blades may be held materially lower than that of the gases in which they operate by providing cooling passages in the blade and conducting a stream of cooling fluid, such as air, through these passages to cool the blade. The constructions which have been suggested so far have attempted to fulfill these requirements but none have proved completely satisfactory.

In some of the suggested structures, the working portion of the blade is fabricated by covering a ribbed in- Zi Patented Eec. 2?, 195@ ferred embodiments both of which are illustrated in the accompanying drawings in which like reference characters denote corresponding parts in the several views and in which:

Figure 1 is a partial perspective view of the blade portion of a gas turbine blade made in accordance with one embodiment of the present invention;

Figure 2 is a side elevation of a piece of sheet stock from which the blade of Figure 1 may be made;

Figure 3 is a side elevation view of the sheet stock of Figure 2 after a forming operation;

ner core having generally an aerofoil cross-section, witha thin skin; the skin serving as a closure for the channels dened ,between the ribs to form passages through which a cooling fluid may be passed to cool the blade. The chief drawback of this construction lies in the dificulty of accurately forming a ribbed core having the proper aerofoil cross-section, this being a necessity since the covering skin is usually too thin the permit machining or other types of forming after it has been applied to the core. It is a very exacting and costly process to control the size of the ribs and the channels between the ribs while, at the same time, producing the desired aerofoil shape. Also, it has been found that the skin often lacks dimensional stability after application, warping being common unless the skin is relatively thick, a feature which is undesirable if efcient cooling of the hot blade surface is to be achieved.

The principal object of this invention, accordingly, is to provide a turbine blade having cooling passages near the hot surfaces of the blade.

A further object of the invention is to provide a gas turbine blade which is provided with cooling passages and which is fabricated solely from sheet stock in a manner which is adapted to mass production techniques.

It is a further object of the present invention to provide a gas turbine blade formed from sheet stock wherein the cooling passages are integral with the sheet material and which serve to stiften the sheet material.

According to the invention a gas turbine blade compris a skin of sheet metal bent over upon itself to form a blade portion having a smooth, continuous outer surface of aerofoil cross-section, one pair of edges of the skin being joined together to form the trailing edge of the blade, the inner surface of the blade skin being provided, at points intermediate its leading and trailing edge, with spaced, integrally formed ribs extending inwardly from the inner surface of the skin, each rib defining, with an adjacent rib, a groove, and means spanning the groove between adjacent ribs enclosing the grooves to form closed passages for cooling fluid.

The invention is described with reference to two pre- Figure 4 is a partial perspective view of the blade portion of a gas turbine blade made in accordance with the second embodiment of the present invention, and

Figure 5 is a side elevation view of a portion of the blade of Figure 4.

Before preceeding further with the specific description of the structural features of the present invention it is desired to define certain terms which will be used in the description and in the claims. The term gas turbine blade is intended, throughout the specification and claims, to mean a blade of the type suitable for use in a gas turbine engine and the term is intended to include both compressor blades and turbine blades of both the stationary and rotating type. The term blade portion is intended to mean that portion of the gas turbine blade which extends into the moving stream of working fluid y structure.

The root or other mounting means associated with,

the blade portion forms no part of the present invention and may take several forms, any one of which is well known in the art.

Referring now to Figure l it will be seen that the blade 10 is formed from a single piece of sheet metal 11 which is bent over upon itself along a fold line 12, the fold line 12 lying substantially at the centre of the leading edge of the blade. One pair of edges of the sheet metal skin 11 are joined together at 13 to form the trailing edge of the blade and the inner surface of the blade skin is provided, at points intermediate its leading and trailing edge, with spaced integrally formed ribs 14 which extend inwardly from the inner surface of the skin substantially normal thereto.

It will be seen, by reference to Figures 1, 2 and 3 that each rib 14 defines, with an adjacent rib, a groove 15, both the ribs and the grooves lying spanwise of the blade as may be seen in Figure 1.

Referring now to Figures 2 and 3 it will be seen that. each rib 14 is provided with an extension 16 of a thickness less than the thickness of the rib 14, the thickness of the extension 16 being, preferably, substantially half of the thickness of the rib 14. The length of the extension 16 beyond the end of the rib 14 is slightly longer than the width of the groove 15 so that when the extension 16 is bent over -at right angles to the rib 14 to form cover means for the groove 15 it will completely span the groove 15 (see Fig. 2) between adjacent ribs 14 to enclose the groove to form a closed passage 15a (see Fig. 3) for cooling fluid immediately adjacent to the skin 11 of the gas turbine blade.

At the point of juncture between the rib 14 and the extension 16 there is provided a shoulder 17 against which the end 16a of the extension 16 of an adjacent rib may abut as shown in Figure 3, the end 16a of the extension 16 being brazed to the shoulder 17 of the adjacent rib.

The ribs 14 are provided, in a spanwise direction of the blade, over an area intermediate the leading edge 12 and 3 the trailing edge 13 of the gas turbine blade 10. The rib 14a is separated from the leading edge 12 by a space 18 and rib fab is separated from the trailing edge 13 by a space 19. The ribs 14 do not extend towards the leadingY edge farther than the position shown in Figure 1 for the reason that if they did the curvature ofthe skin 11 at this point in the blade would be such that the ribs would be distorted and the extension 16 would not be able to occupy the position shown in Figure 3 without buckling. The ribs 14 do not extend towards the trailing edge farther than shown in Figure l for the reason that if they did their height would be such that the rib 14b on the concave side Ztl of the blade 10 would interfere with the corresponding rib db on the convex` side 21 of the blade 10. In addition, it is well known that the leading edge and the trailing edge ofthe gas turbineblades are commonly areas of greater temperature than the mid-chord region of the blade and, accordingly, with the present construction, passages 1S and 19 are provided at the leading and trailing edges respectively which provide for the passage of rlarger quantities of cooling iiuid to maintain the ternperature of the blade at these points at a satisfactory level;

The ribs 14 and the extensions 16 are formed integrallyl with the skin 11 by any suitable means such as rolling, forging, stamping, machining, grinding or by wrapping a sheet of metal about a mandrel and turning a helical groove in the surface ofthe sheet by revolving the man drel against a cutting tool. When the sheet is removed from the mandrel and laid fiat,V the grooves will run diagonally across the sheet and an appropriate section may be cut from the sheet tol form the blade shown in, Figure l. `The ribs which will, in the machined state, occupy the

space

18 and 19 may be removed by grinding, planing or any other machining operation.

Referring now te` Figures 4 and 5 an alternative embodiment of theinventionwill be, described andvinthese. views the` blade 110 is., shownl as beingl formed from; ak sheet 111 by bending about a fold linel 11,2v which also, constitutes the leading edge of the blade, one pair of edges of the blade being joined together at 113 to form the trailing edge of the blade. Ribs 114 are provided integrally with the inner surface of the skin 111 andextend inwardly from the skin substantially normal thereto. The rib 114e is separated from the leading edge 112 by a space 118 and the rib 114th is separated from the trailing edge 113 by a space 11S. Each rib 114 defines, with an adjacent rib, a groove 115 which is enclosed by cover means 116 formed by a channel member having legs 122 and 123 which extend into the groove 115 between adf jacent ribs 114, thel external surfaces of legs 122 and 123 lying in close abutment with the side walls` 114C of the groove 115. Each cover means 116 is brazed in one of the grooves 115 in the position shown in Figure 5 to provide completely enclosed grooves 115 lying spanwise of the blade immediately adjacent to the inner surface of, the skin 111.

The passages 118 and 119 are provided adjacent the leading and trailing edges respectively to carry relatively large quantities of cooling fluid to adequately cool these hot areas of the blade.

It will be appreciated that the construction disclosed in this specification provides a blade skin which is more,y

rigid, due to the reinforcing action of the ribs 114, than` would be the case if only an unsupported thin skin were employed. The overall weight of the blade is, of course, materially less than that of a blade manufactured with a solid` ribbed core and the advantages of that construction, in providing cooling passages immediately adjacent the surface of the skin, is obtained without the disadvantage of the excess weight.

The construction of the ribs and skin is such that it is well adapted to mass production techniques as is the case with the channel members 116 which may be rolled, drawn or extruded from thin sheet stock.

The brazing operation may be carried out economically and simply by any of the known brazing techniques either before or after the sheet has been fabricated into ablade section.

While the invention has been described in particular detail in reference to two preferred embodiments it is to be appreciated that minor modifications may be made in the construction thereof without departing from the spirit of the invention or the scope of the appended claims.

What I claim as my invention is:

l. Av gas turbine blade comprising a sheet metal skin having a smooth, continuous one-piece outer surface of aerofoil cross-section, extending from a leading edge to a trailing edge, the blade skin having an inner surface on which are provided, intermediate its leading and trailing edge,- spaced, parallel ribs integrally formed with the skin and projecting inwardly from its inner surface and extending spanwise of the blade, each rib defining, with an adjacent rib, a groove, and cover means for each groove supported by the two ribs defining such groove and separate from the cover means for every other groove, each cover means comprising a U-shaped channel of sheetmetal, the legs of the U extending between adjacent ribs and in contact with their adjacent surfaces and brazed thereto to enclose each groove to form passages for cooling fluid.

2. A gas` turbine blade having a. smooth, continuous, one-piece sheet metal skin presenting an outer surface of aerofoil cross-section extending from a leading edge to a trailing edge' through a concave side and' a convex side, a first set of a plurality of spaced, parallel ribs integrallyformed with the skin and projecting inwardly from the inner surface of the concave side of the skin and extendingspanwise of the blade, each rib defining, with an adjacent rib, a groove, a second set of a plurality of spaced, parallel ribs integrally formed with the skin and projecting inwardly from the inner surface of the convexV side ofl the skin and extending spanwise of the blade, each rib of the second set defining, with an adja centV rib of the second set, a groove, the ribs on the concave and convex sides of the blade projecting towards each other but terminating Short of Contact with one another, and cover means for each groove supported. by the two ribsdefining such grooves and separate from the cover means for every other groove to enclose each groove to form passages for cooling fluid, the inner surface of the skin and the covermeans for the grooves defining a central cavity within the blade extending spanwise of the blade and from leading edge to trailing edge.

3. A gas turbine blade as claimed in claim 2 in which the cover means for each groove between adjacent ribs comprises a U-shaped channel of sheet metal, the legs of; the U extending. between adjacent ribs and in contact with their adjacent surfaces and brazed thereto.

References (-Iited in the file of this patent UNITED STATES PATENTS 2,779,565 Bruckmann Jan. 29, 1957 2,787,441 Bartlett Apr. 2, 1957 2,807,437 Roush Sept. 24, 1957 2,828,106 Schramm Mar. 25, 1958