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Publication numberUS3834833 A
Publication typeGrant
Publication dateSep 10, 1974
Filing dateFeb 16, 1973
Priority dateFeb 18, 1972
Also published asCA1009125A1, DE2211830A1
Publication numberUS 3834833 A, US 3834833A, US-A-3834833, US3834833 A, US3834833A
InventorsFaber G, Maggi C
Original AssigneeBbc Brown Boveri & Cie
Export CitationBiBTeX, EndNote, RefMan
External Links: USPTO, USPTO Assignment, Espacenet
Blade construction for axial-flow turbo-machines and method of protecting turbo-machine blades against stress corrosion cracking
US 3834833 A
Abstract
A steel blade for an axial-flow turbo-machine is disclosed which has a hardened leading edge portion. In accordance with the invention, the hardened edge portion is imparted with a permanent compressive prestress in the direction of the longitudinal axis of the blade. This protects the blade against stress corrosion cracking.
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Description  (OCR text may contain errors)

United States Patent Faber et a1.

[ BLADE CONSTRUCTION FOR AXIAL-FLOW TURBO-MACHINES AND METHOD OF PROTECTING TURBO-MACHINE BLADES AGAINST STRESS CORROSION CRACKING Inventors: Guy Faber, Oberrohrdorf; Carlo Maggi, Baden, both of Switzerland Assignee: Brown, Boveri & Company Ltd.,

Baden, Switzerland Filed: Feb. 16, 1973 Appl. No.: 333,156

Foreign Application Priority Data Feb. 18, 1972 Switzerland 2373/72 US. Cl. 416/241, 416/224 Int. Cl. F0ld 5/28 Field of Search 416/224, 241

References Cited UNITED STATES PATENTS 5/1959 Pekarek 416/241 [45.] Sept. 10, 1974 3,148,954 9/1964 Haas 416/241 X FOREIGN PATENTS OR APPLICATIONS 195,050 7/1923 Great Britain 416/224 Primary Examiner-Everett A. Powell, Jr. Attorney, Agent, or Firm-Toren, McGeady and Stanger [57] ABSTRACT A steel blade for an axial-flow turbo-machine is disclosed which has a hardened leading edge portion. In accordance with the invention, the hardened edge portion is imparted with a permanent compressive prestress in the direction of the longitudinal axis of the blade. This protects the blade against stress corrosion cracking.

2 Claims, 1 Drawing Figure PATENTED SE? 1 M974 3. 884.833

BLADE CONSTRUCTION FOR AXIAL-FLOW TURBO-MACHINES AND METHOD OF PROTECTING TURBO-MACHINE BLADES AGAINST STRESS CORROSION CRACKING FIELD OF INVENTION The invention is directed to a procedure for protecting the hardened leading edge of blades or vanes (hereinafter blades) of axial-flow turbo-machines against stress corrosion cracking.

BACKGROUND INFORMATION It is general practice to harden the leading edges of blades of axial-flow turbo-machines for the purpose of minimizing erosion of the blades. The desired hardness may be obtained by various heat treatments, for example by flame hardening or by means of induction heat by high frequency.

It is well known, however, that hardened steel is more susceptible to stress corrosion cracking sometimes referred to in the art as stress crack corrosion than is unhardened steel. The phenomenon of stress corrosion is discussed in The Making, Shaping and Treating of Steel, United States Steel Corporation, 1964, page 933. Hardened alloy steel, for example stainless steel containing 12 percent of chromium, exhibits a particular great tendency towards stress corrosion cracking. This susceptibility is dependent on various factors, among others the composition and the heat treatment condition of the steel. Further, it is dependent on the contaminants, for example chlorides, which are present in the flowing working medium of the turbo-machine and also on the tensile stresses which act on the respectively endangered portion of the blade. Since a blade of the indicated kind, during operation of the machine, is subjected to tensile stresses due to the centrifugal forces, and sometimes also due to forces caused by the working medium, also a hardened edge of such blades is attacked by stress corrosion cracking.

It has previously been proposed to reduce the mechanical stresses on the blades by construction measures. This, however, results indisadvantages such as,

SUMMARY OF THE INVENTION It is the primary object of the present invention to provide a procedure by means of which the hardened leading edge of a blade of an axial-flow turbo-machine is largely protected against stress corrosion cracking.

Another object of the invention is to achieve such protection by technological, as distinguished from constructive, measures.

Generally it is an object of the present invention to improve the characteristics of the blades of axial-flow turbo-machines.

Briefly, and in accordance with the invention, the leading edge of such a blade, and while it is being hardened, is imparted with a permanent, compressive prestress in the direction of the longitudinal axis of the blade.

If a blade which has been imparted with such an initial compressive stress is, during operation, subjected to tensile stresses caused by centrifugal or other forces, that portion of the blade which exhibits the compressive prestress is then completely, or at least to a very large extent, relieved of the pressure. Any tensile forces which may still arise in this portion of the blade are so insignificant that they do not cause any substantial. in.- crease in the stress corrosion cracking.

The blades may be imparted with the permanent compressive prestress in exceedingly simple manner. Generally, hardening of steel is effected by heating to a temperature above 850C. and subsequent cooling or quenching. Due to the martensite formation in the steel, a volume increase takes place. If only a portion of a steel work piece is hardened in the indicated manner, as is the case in the hardening of the leading edge of a blade, then a compressive stress or prestress occurs in the hardened portion, provided the remainder of the work piece, which remains cold or unheated during'the hardening, has a sufficiently large'cross-sectional area in order to enable it to absorb the resulting tensile stresses without being plastically deformed, i.e. beyond the yield point.

In a blade having an oblong profile of which merely one end portion, such as the leading edge, is to be hardened, the conditions are particularly unfavorable. Extensive experiments have indicated that a permanent compressive prestress in the direction of the longitudinal axis of the blade, which prestress thus also remains in cold condition, is imparted with certainty only to the hardened portion if the cross-sectional area of this hardened blade portion occupies at the most 20 percent of the total blade cross section. The remaining percent of the cross-sectional area of the unhardened blade portion can then effectively absorb the corresponding tensile stresses without. deformation.

Accordingly, and pursuant to the invention, the leading edge portion of the blade to be hardened is not permitted to have a cross-sectional area in excess of 20 percent of the total cross section of the blade and the thus remaining 80 percent are maintained in cold, unheated condition during the hardening, whereby the hardened portion is imparted with the desired permanent compressive prestress.

The various features of novelty which characterize the invention are pointed out with particularity in the claims annexed to and forming a part of this specification. For a better understanding of the invention, its operating advantages and specific objects attained by its use, reference should be had to the accompanying drawing and descriptive matter in which there is illustrated and described a preferred embodiment of the invention.

The single FIGURE of the drawing is a crosssectional view of a blade of an axial-flow turbomachine.

It will be noted that the cross-sectional area of the blade is formed by the portions 1 and 2, portion 1 being the cross-sectional area of the hardened zone, to wit, the area of the leading edge of the blade. This hardened zone 1 extends not only adjacent the leading edge but also adjacent a portion of the suction side of the blade. The area 2 occupies the remainder of the crosssectional area of the blade. In accordance with the invention and in order to create during the hardening the desired permanent compressive prestress, the crosssectional area 1 amounts to at the most, and preferably less than, 20 percent of the total cross-sectional area, to wit, the sum of the cross-sectional area 1 and 2. The hardening of the area 1 may be: accomplished by any conventional tempering treatment as previously re ferred to.

blade is protected against stress corrosion cracking.

2. The improvement of claim 1, wherein the crosssectional area of said hardened edge portion amounts to at the most twenty percent of the total crosssectional area of the blade.

Patent Citations
Cited PatentFiling datePublication dateApplicantTitle
US2888244 *May 24, 1956May 26, 1959Thompson Ramo Wooldridge IncFluid directing member
US3148954 *Jun 13, 1960Sep 15, 1964Haas IreneTurbine blade construction
US3371908 *Nov 1, 1966Mar 5, 1968Tokyo Shibaura Electric CoTurbine blading components and process of producing the same
US3564689 *May 27, 1968Feb 23, 1971Boehler & Co Ag GebMethod of fabricating a turbine blade having a leading edge formed of weld metal
US3729345 *Sep 20, 1971Apr 24, 1973Mitsubishi Heavy Ind LtdMethod for making propellers of high-strength and high-toughness cast steel
GB195050A * Title not available
Referenced by
Citing PatentFiling datePublication dateApplicantTitle
US4385866 *Mar 17, 1980May 31, 1983Tokyo Shibaura Denki Kabushiki KaishaCurved blade rotor for a turbo supercharger
US4778345 *Mar 14, 1986Oct 18, 1988Ngk Spark Plug Co., Ltd.Ceramic blade heat connected to hardened steel shaft
US5120197 *Jul 16, 1990Jun 9, 1992General Electric CompanyTip-shrouded blades and method of manufacture
US5348446 *Apr 28, 1993Sep 20, 1994General Electric CompanyNickel-aluminum alloy, cooling
US5620307 *Mar 6, 1995Apr 15, 1997General Electric CompanyLaser shock peened gas turbine engine blade tip
US5742028 *Jul 24, 1996Apr 21, 1998General Electric CompanyPreloaded laser shock peening
US6004102 *Aug 11, 1997Dec 21, 1999Abb Patent GmbhTurbine blade for use in the wet steam region of penultimate and ultimate stages of turbines
US6155789 *Apr 6, 1999Dec 5, 2000General Electric CompanyGas turbine engine airfoil damper and method for production
US6551064Jul 24, 1996Apr 22, 2003General Electric CompanyLaser shock peened gas turbine engine intermetallic parts
Classifications
U.S. Classification416/241.00R, 416/224
International ClassificationF01D5/28, C21D9/00
Cooperative ClassificationF01D5/286, C21D9/0068
European ClassificationF01D5/28D, C21D9/00P