Abstract: A lightweight high temperature multiproperty rotor disk and method of manufacture for use in a gas turbine engine. The rotor disk of a nickel super alloy and having a plurality of lightweight high temperature single crystal blade attachment lugs bonded thereto. The single crystal high temperature rotor blade attachment lugs being cast in a substantially rectangular shape and bonded to the circumference of the disk. A method of manufacturing the rotor includes positioning the plurality of single crystal attachment lugs on a positioning ring that has locating nests formed thereon. Subsequently, the rotor disk is placed within the bonding ring and a restraining ring is placed around the entire subassembly. Upon subjecting the assembly to high temperature the rotor disk, having a coefficient of thermal expansion greater than the coefficient of thermal expansion of the restraining member, forces the components into a mating relationship.

Abstract: A gas turbine blade includes an airfoil, platform, and dovetail. The blade is cast with a chamfer along one edge of the platform thereof. The platform edge is then machined to truncate the chamfer in a simple and precise method.

Abstract: A pair of mating titanium alloy substrates for use in a gas turbine engine are provided, one of which comprises an aluminum bronze alloy wear resistant coating. The coating consists essentially of 9.0-11.0% aluminum (Al), 0.0-1.50% iron (Fe), and a remainder of copper (Cu). The wear resistant coating is disposed between the mating substrates and inhibits frictional wear between the mating substrates.

Abstract: A fan rotor assembly (30) for a gas turbine engine (10) comprises a fan rotor (32) and a plurality of radially extending fan blades (34). Each fan blade (34) has a root (64) which is retained in a corresponding groove (54) in the periphery of the hub (54) of the fan rotor (32). Each fan blade root (64) comprises four axially spaced root portions (66, 68, 70 and 72) which locate in a corresponding number of axially spaced groove portions (56, 58, 60 and 62). The flanks (94, 98) of two of the root portions (66, 70) are arranged at an angle to the axis of the fan rotor (32) and the flanks (96, 100) of the other two root portions (68, 72) are arranged at an angle in the opposite sense. The flanks (74, 76, 78 and 80) of the groove portions (56, 58, 60 and 62) are arranged at the same angle as the flanks (94, 96, 98 and 100) of the corresponding root portions (66, 68, 70 and 72). The arrangement provides improved retention of the fan blades (34) and reduces the weight of the fan blades (34) and fan rotor (32).

Abstract: A dovetail joint between a rotor wheel and a bucket includes a male dovetail component on the rotor wheel and a female dovetail component in the bucket. The male dovetail component has axially projecting hooks with slanted crush surfaces along generally radially inwardly directed surfaces. The slanted surfaces form included angles with neck portions parallel to a plane normal to the axis of rotation which are larger than 90.degree. and decrease in included angle from the radially outermost hooks to the radially innermost hook where the angle is 90.degree.. Compound fillets are also provided along the transition surfaces between the slanted crush surfaces and the neck surfaces. Thus, stress concentrations are minimized.

Abstract: A process for joining a base for an airfoil to a disk for an integrally bladed rotor stage in a gas turbine engine includes providing a disk having a radially outer rim with a slot defined therein by a recessed surface, and further includes providing a base having a longitudinally extending root portion facing opposite a longitudinally extending support portion for supporting the airfoil. The root portion has a root surface and the support portion has an outer surface. The process further includes bringing the root surface of the root portion into contact with the recessed surface bounding the slot, applying pressure and relative movement between the base and the rim to achieve during welding, substantial contiguity between the root surface and the recessed surface over the area of the recessed surface, resulting in a substantially continuous linear friction weld between the base and the rim.

Abstract: A pair of mating titanium alloy substrates for use in a gas turbine engine are provided, one of which comprises an aluminum bronze alloy wear resistant coating. The coating consists essentially of 9.0-11.0% aluminum (Al), 0.0-1.50% iron (Fe), and a remainder of copper (Cu). The wear resistant coating is disposed between the mating substrates and inhibits frictional wear between the mating substrates.

Abstract: A rotor for a turbomachine has slots formed therein at an angle to an axis of rotation of the rotor. Blades have at least one blade root with at least two regions of different rigidity that are adapted and preferably matched to different regions of rigidity of the slot. The blade roots can be fitted into the slots. A blade for a rotor is also provided. The invention is especially suitable for use in gas-turbine compressors, as a result of which local stresses at slots can be greatly reduced.

Abstract: A damper for the blades of a turbine rotor for attenuating the vibrations excited in a radial direction by configuring the damper to have at least two legs bearing against the blade and the disc supporting the blade for grounding the radial vibrations transmitted radially through the damper and offsetting the center of gravity of the damper from the point of contact with the blade.

Abstract: A rotor blade includes an airfoil, an integral platform, and a dovetail integrally joined to the platform. The dovetail includes a neck extending between a pair of dovetail lobes configured to engage a complementary slot in a rotor disk. The dovetail includes an undercut extending between an outer face of the dovetail lobes and the neck for introducing a stress concentration thereat. The stress concentration relieves stress at the outer face at the expense of the neck for increasing the high cycle fatigue life of the blade.

Abstract: The invention provides a blade-rotating device for rotating a plurality of radially-oriented blades having a blade base with two grooved base sides. The device has a portion, such as a disc or a rotor disc of a turbine, for rotating the blades with the portion having an outer cylindrical surface. The device also has a plurality of steel steeples each having a steeple base and two opposite grooved steeple sides extending therefrom. Each of the opposite steeple sides complements at least one of the grooved sides of at least one of the blade bases. The steeple bases are linear friction welded to the outer cylindrical surface in an approximately longitudinal direction such that the blade bases are engagable with the steel steeples.

Abstract: A turbine or compressor disk assembly comprises a plurality of blades attached to a central hub by means of the blade root of each blade engaging a corresponding slot in the hub. According to the principles of the present invention, a turbine or compressor disk assembly includes one or more locally bulging regions extending axially away from the surface of the disk in the vicinity of the bottom contact plane of the disk attachment firtree. The locally bulging regions reduce the peak Macke stress in the disk bottom fillet and blade attachment root. In an illustrative embodiment, two locally bulging regions are incorporated into the disk and the corresponding blades, one of which extends forward from the leading edge of the disk and the other of which extends rearward from the trailing edge of the disk. The rearward locally bulging region has an exaggerated extension which allows the stress reduction ring to form an aft flow discourager as well as functioning to reduce peak stress.

Abstract: A rotary turbomachine, and particularly a turbojet engine, has a compressor stage that is especially adapted for operation in the transonic region. The compressor stage includes a rotor (1) having a hub (2) and a plurality of compressor blades (3) extending radially therefrom, and a stator having a plurality of compressor blades extending radially between a stator hub and a housing. The hub (2) of the rotor (1), and/or the hub of the stator, and/or the housing of the stator, have a circumferential surface contour that is not continuously rotationally symmetrical. Namely, a concave contour (K) is provided in the circumferential surface of the hub near the base of each blade (3) on the pressure side (PS) thereof, while the circumferential surface on the suction side (SS) of the base of each blade has a contour that is linear, slightly convex, convex/concave, or slightly concave to a lesser degree than the concave contour (K) on the pressure side of each blade.

Abstract: A gas turbine engine fan blade includes an airfoil and dovetail having an integral aluminum core and a fatigue resistant reinforcing sheath.

Abstract: A fan blade (34) of a gas turbine engine (10) includes a pair of root rails (54) disposed on a bottom (56) of a dovetail root portion (44) thereof. The pair of root rails (54) has a wedge-shaped contour and is tapered in a decreasing cross-section from the leading edge (50) to the trailing edge (52). The root rails (54) provide a wide, stable base for the blade to minimize tangential motion during windmilling. A pair of slot rails (76) is provided on the base (70) of a dovetail-shaped slot (40) retaining the dovetail root portion (44) of the blade (34). The slot rails (76) of the dovetail slot (40) mate with the root rails (54) of the dovetail root portion (44).

Abstract: In an arrangement for fixing blades (5) to a rotor (1) or stator, a number of slots (7) which run helically on the circumferential surface of the body to be fitted with blades are provided, filling pieces (6) which serve to space apart the blades (5) and fix the blades (5) in a frictional manner being present between individual blade rows (3).

Abstract: An annular seal plate for an interstage air riding seal arrangement for the internal cooling system of a gas turbine engine is mounted on a rotatable disc by means of an annular mortise and tenon mounting. The contact faces of the mortise and tenon are angled relative to a radial plane of the disc so that centrifugally generated forces are reacted by the faces in a sense to counter axial tilt so that a seal face of the plate lies exactly in a radial plane, at a design rotational speed, thereby to eliminate convergence or divergence between the seal faces of the air riding seal.

Abstract: A fan for a gas turbine engine comprises a rotary disc carrying an annular array of radially extending fan blades. Each fan blade has a root portion which is radially constrained within a generally axially extending groove formed within a disc rim. The blade root comprises two portions each radially convergent toward one another with respect to the axis of rotation of the disc. The groove within the disc also comprises two base portions each being correspondingly radially convergent toward one another. Axial retention of the blade root within the groove is achieved by the centrifugal component of the blade load acting in an opposite direction to any applied axial load in the forward or rearward direction.

Abstract: A gas turbine engine rotor disc (20) is provided with a plurality of cooling air supply passages (32) which are generally radially extending but are inclined in the downstream direction. Each passage (32) has a cross-sectional configuration which renders the ends (34) of the passages (32) less likely than normal to act as the site of hoop-stress induced cracks.

Abstract: A vane for a fan including a platform, a root and a blade. The platform has opposite outer and inner surfaces, and opposite upper and lower surfaces. The upper surface has a camber shape, and the lower surface has a concave shape corresponding to the camber shape of the upper surface. The root is fixed to the inner surface of the platform and adapted to be fixed to a hub of the fan. The blade is fixed to the outer surface of the platform at a position shifted toward the lower surface from a center line between the upper and lower surfaces.

Abstract: A dovetail assembly for mounting blades with dovetail roots axially into disk dovetail slots around a disk periphery which is subject to a tensile stress field due to centrifugal and thermal growth induced forces generated by the rotor when the rotor is rotating. The dovetail assembly has stress risers each in the form of a transition fillet between a neck or an area of minimum width and a pressure face which is designed to be a contact surface between complementary blade dovetail roots and disk dovetail slots formed between adjacent disk posts. The region around the stress riser has deep compressive residual stresses imparted by laser shock peening and is radially adjacent the stress riser and extends inward from a laser shocked surface of the component along the transition fillet.

Abstract: Damping disposition for vanes (3) fitted with stilts (4, 5) engaged in the alveoles of the rotor disk (1). Damping elements (10) are added including a portion (11, 12) extending along the flanks (6) of the stilts (4) and including a portion (15) leaning against them. The beating vibrations of the vanes (3) and dampened by these elements able to be used on smooth vanes without any platform, such as large blower vanes. Application for turbo-engines.

Abstract: A root attachment for a rotor blade of a gas turbine engine comprises teeth on the root section of the rotor blade and teeth on the slot of the rotor. The teeth flank surfaces of the teeth on the root section of the rotor blade are arranged at a different angle to the teeth flank surfaces of the teeth on the slot of the rotor. As the load on the rotor blade increases the teeth deflect such that the area of contact between the teeth flank surfaces of the root section and the slot increases to a maximum at the fully loaded condition. The teeth flank surfaces are substantially planar. This enables rotor blade and slot teeth with different stiffnesses to be used for the rotor blades and rotor, for example titanium aluminide turbine blades and nickel base alloy turbine rotor.

Abstract: A rotor disc which is suitable for carrying the fan blades of a ducted fan gas turbine engine is constituted by two sub-discs which are maintained in axially spaced apart relationship by a plurality of spacer members. The spacer members are generally axially extending and at least partially define the operational air flow path over the disc.

Abstract: A turbine rotor assembly includes a rotor disk, rotor blades joined thereto, and a blade retention plate for axially retaining the blades on the disk. The disk includes an outer rim, an inner hub, and an annular web extending therebetween. An axial extension extends from the web and is spaced radially below a corner of the rim. The blades have dovetails axially disposed in respective dovetail slots in the disk rim. The retention plate includes an outer rim which abuts the disk rim and blade dovetails, an inner hub, and a web extending therebetween. The plate hub has an outer rabbet disposed in abutting contact with the rim corner, and an inner rabbet radially adjoining the axial extension for maintaining abutting contact therewith upon differential radial contraction between the retention plate and the disk which separates the outer rabbet from the rim corner.

Abstract: A lightweight high temperature multiproperty rotor disk and method of manufacture for use in a gas turbine engine. The rotor disk of a nickel super alloy and having a plurality of lightweight high temperature single crystal blade attachment lugs bonded thereto. The single crystal high temperature rotor blade attachment lugs being cast in a substantially rectangular shape and bonded to the circumference of the disk. A method of manufacturing the rotor includes positioning the plurality of single crystal attachment lugs on a positioning ring that has locating nests formed thereon. Subsequently, the rotor disk is placed within the bonding ring and a restraining ring is placed around the entire subassembly. Upon subjecting the assembly to high temperature the rotor disk, having a coefficient of thermal expansion greater than the coefficient of thermal expansion of the restraining member, forces the components into a mating relationship.

Abstract: A cobalt based coating is provided which is functional to reduce the coefficient of friction between two operatively, frictionally engaging titanium alloy surfaces.

Abstract: A rotor of a turbomachine includes blades having an improved blade root configuration for anchoring and centering the blade roots in axial slots of the rotor disk. Each blade root includes two portions or sections, namely a serrated section that has a single bilaterally uniformly shaped serration, and a terminal portion that has rounded surfaces and a narrower cross-section than the serrated root section. The serrated section provides anchoring against centrifugal forces. The terminal portion provides the predominant centering effect, for which it is radially movably received and located in a recess provided in the bottom of the slot. The special root configuration substantially avoids misfit and centering problems associated with manufacturing variances, whereby the blades and blade roots can be made of lightweight engineering materials.

Abstract: In a bladed rotor of an axial-flow turbomachine, the blades essentially consist of a blade body and a blade root. They are fastened with their blade roots in rows in encircling blade grooves having lateral supporting prongs. A shoulder serving to assemble the blades on the rotor is attached between blade body and blade root. The blade grooves are shaped symmetrically to the associated planes of symmetry by a plurality of radii (R1-R4), the preceding radius (R1) in each case being larger than the following radius (R2). The circular arcs formed by the radii (R1-R4) have a common tangent (T) at their contact points.

Abstract: In a drum rotor (1) for an axial-flow turbomachine, the blades are fastened with their roots (3) in rows in encircling blade grooves (10) having lateral supporting prongs. At least the first blade groove (10), on the inlet side of the working medium, is shaped in an axially asymmetrical manner by a plurality of radii (R1-R4). The preceding radius (R1) in each case is larger than the following radius (R2), and the circular arcs (21-24) formed by the radii (R1-R4) have a common tangent (T) at their contact points.

Abstract: An assembly of a plurality of composite blades including blade roots carried by blade root receiving slots in the rotor wherein the slot has a slot wall with a radially outward portion which, when assembled, diverges from a spaced apart juxtaposed blade root pressure face radially outer surface in an amount which is a function of a predetermined amount of centrifugal loading on the blade during operation of the assembly, to allow at least a portion of the radially outer surface of the root pressure face to be in contact with the slot wall radially outward surface during operation.

Abstract: Blades for use in combusion turbine engine are formed from a single crystal garnet body--suitably an aluminum garnet, desirably a rare earth aluminum garnet--having an epitaxial surface layer of a second garnet having a lattice constant larger than that of the body garnet so as a compressive strain, to thereby strengthen the blade.

Abstract: An arrangement for clipping stress peaks in the anchoring of a turbine blade including a root that is shaped like a Christmas tree, that includes tangs, and that is received in a correspondingly-shaped groove in a disk so as to anchor the blade thereto, wherein the outline of each tang, on the extrados side of the blade root, is such that the clearance for mounting the root in its groove, on the extrados side, is not uniform but rather it includes at least one zone in which the clearance varies, the zone corresponding to a zone of the tang in question where, if the disk-and-blade assembly were rotating at its rated speed of rotation, there would otherwise occur a stress peak relative to the mean stress along the tang if the blade were mounted with constant clearance along the entire length of the tang, the varying clearance varying along the at least one zone with the same sign as the stress peak.

Abstract: In a bladed rotor (1) of a turbo-machine, the blade roots (3) and blade grooves are made pinetree-shaped with a plurality of indentations (10, 21) and teeth (11, 20). A plurality of supporting surfaces (F) extending obliquely are formed thereby. The indentations (10, 21), adjoining the supporting surfaces, in the blade root (3) and in the blade groove are described essentially by a first curve radius (R1) and a second curve radius (R2).The first curve radius (R1) adjacent to the supporting surface (F) is smaller than the second curve radius (R2). The opening angle (W) between a tangent (T) to the circle arc (22), described by the first curve radius R1, at the end point (A, B) of the supporting surface and the supporting surface (F) amounts to at least forty degrees.

Abstract: In a steam turbine rotor wheel and bucket dovetail joint construction wherein the wheel dovetail includes a plurality of radially adjacent hooks interconnected by respective necks, each having an upper and lower fillet pair of substantially identical fillet, an improvement wherein each upper fillet has a compound radius.

Abstract: Turbomachinery rotor component, which may be a disk or drum rotor, having an annular rim disposed about a centerline for securing the blades to the rotor and that is subject to a tensile stress field due to centrifugal and thermal growth induced forces generated by the rotor when the rotor is rotating. An annular channel slot is cut around the rim, has outer annular overhanging rails, and has at least one stress riser in the form of a three sided slot cut in the radial direction through one of the rails. A region around rounded corners of the stress riser has deep compressive residual stresses imparted by laser shock peening and extends inward from a laser shocked surface of the component encompassed by the region.

Abstract: Cooling air delivery systems for gas turbine engines are used to increase component life and increase power and efficiencies. The present system increases the component life and increases efficiencies by better utilizing the cooling air bled from the compressor section of the gas turbine engine. For example, a flow of cooling air which is directed through internal passages of the engine and into a cavity and is used to cool the interface between a turbine disc and a plurality of turbine blades and further reduces ingestion of hot power gases into the internal portion of the gas turbine engine. The cavity has a generally tapered configuration and effectively cools the interface between the disc and the blade. Thus, increasing component life and increasing power and efficiencies of the engine.

Abstract: A blade support includes first and second abutting disks, with the first disk having a plurality of axial dovetail slots at one end thereof, and a circumferential dovetail first half-slot at an opposite end thereof, with the second disk having a circumferential dovetail second half-slot therein. A plurality of rotor blades have axial dovetails disposed in respective ones of the axial dovetail slots, and circumferential dovetails disposed in part in each of the first and second half-slots. The first and second disks are retained together so that the first and second half-slots collectively support the circumferential dovetails.

Abstract: An improved, highly stable blade-to-blade vibration damper configuration which provides substantially continuous blade vibration damping and sealing of an interplatform gap due to positional stability thereof. Disposed in a subplatform cavity, the damper is comprised of a generally triangular shaped body having a vertex thereof aligned with the interplatform gap. A primary damper load face abuts a first inclined platform load face and a secondary damper load face abuts a second inclined platform load face. Maintenance of sliding planar contact between primary and first load faces and sliding linear contact between secondary and second load faces is afforded by orienting inclined platform faces to have an included angle greater than that of the damper vertex and offsetting a damper center of gravity toward the primary load face. The damper may include one or more legs to orient the damper in the cavity and one or more extending tabs to discourage hot gas flow thereby.

Abstract: In the rotor of a turbomachine, especially the turbine of a gas turbine engine, each rotor blade is mounted in the rotor disk rim by a blade root having a serrated profile that engages a correspondingly formed axial groove in the rotor disk rim. A shroud band is constructed so that at a predetermined operating point of the turbine, adjacent blades are supported against one another so as to damp vibrations. Light weight and heavy weight rotor blades are mounted alternatingly around the circumference of the rotor disk. The blade root and corresponding axial groove associated with each light weight or heavy weight blade are differently dimensioned or shaped to correspond to the differing loads associated with the heavy weight blades and the light weight blades respectively. Different materials, having different densities, are used for the heavy weight blades and the light weight blades.

Abstract: A dovetail joint between a rotor wheel and a bucket includes a male dovetail component on the rotor wheel and a female dovetail component in the bucket, wherein the male dovetail component includes four laterally projecting hooks, each hook from an uppermost hook to a lowermost hook decreasing in radial thickness and increasing in axial length. Each hook has a substantially flat top surface to effectively increase the average thickness, and a neck between the respective hooks having enlarged fillets. The new dovetail joint reduces peak stresses so as to avoid or minimize stress corrosion cracking.

Abstract: A refrigerant or coolant compressor of the radial type with an impeller hng a plurality of vane elements can be used for compressing water vapor as a refrigerant or coolant under vacuum conditions. The impeller of the compressor is constructed to produce a high volume flow rate at the required compression ratio, in view of the low density of water vapor as the preferred flow medium. The impeller has sufficient strength to operate at the required high circumferential velocities. The impeller includes vane elements, disk elements, vane support elements and a hub. The vane elements are individually connected to the hub by the support elements. The support elements may be either ring-shaped elements connecting a rear surface of the disk elements to the hub, or may be pin-shaped insert members connecting the root of each vane element to the hub. The components of the impeller are made of a polymer composite material reinforced preferably with carbon fibers.

Abstract: In a turbofan engine a fan designed for reduced damage during a blade-out condition wherein the fan blade airfoil is disposed asymmetrically on the platform and includes a forwardly displaced leading edge. The platform is configured to make early contact with a following airfoil upon blade-out.

Abstract: The present invention relieves the stress at a fillet of a blade shank using at least one axially extending pocket on the shank near an axially extending fillet of a shank to root interface or a shank to platform interface on a pressure or suction side of the blade and the pocket is axially shorter than the fillet. The stress relief pocket has a radially extending pocket width that is substantially wider than a fillet width between generally axially extending edges of the fillet wherein the pocket width may be several times wider than the fillet width. The wider pocket shields the fillet thereby reducing the stress concentration at the fillet.

Abstract: A fan blade (34) of a gas turbine engine (10) includes a pair of root rails (54) disposed on a bottom (56) of a dovetail root portion (44) thereof. The pair of root rails (54) has a wedge-shaped contour and is tapered in a decreasing cross-section from the leading edge (50) to the trailing edge (52). The root rails (54) provide a wide, stable base for the blade to minimize tangential motion during windmilling. A pair of slot rails (76) is provided on the base (70) of a dovetail-shaped slot (40) retaining the dovetail root portion (44) of the blade (34). The slot rails (76) of the dovetail slot (40) mate with the root rails (54) of the dovetail root portion (44).

Abstract: The multi-lug attachment of the blades to the disk of a turbine rotor of a gas turbine engine, where the blades are air-cooled, optimizes weight reduction and tensile stress by attachment including a pair of lugs that transitions from a contoured airfoil to a pair of parallel rails releasably attached to complementary recesses formed in the disk. One lug is transitioned from the pressure surface to the attachment surface and the other lug is transitioned from the suction surface to the attachment surface. The lugs are tapered to optimize the attachment.

Abstract: A turbojet engine rotor, wherein the circumference of the disk is occupied y teeth and slots in an alternating form. The teeth have a variable rigidity in the axial direction, which leads to different shapes of their cross-section. The teeth have a twisted outer surface and oblique recesses are provided on the circumference of the rotor for receiving vane roots which alternate with the teeth for separating and retaining the vanes.

Abstract: A turbine blade having a preestablished rate of thermal expansion is attached to a turbine wheel having a preestablished rate of thermal expansion being greater than the preestablished rate of thermal expansion of the turbine blade. The turbine blade has a root portion having a pair of recessed portions thereon. The turbine wheel includes a plurality of openings in which the turbine blade is positioned. Each of the openings have a pair of grooves therein in which are positioned a pair of pins having a generally rectangular cross-section and a reaction surface thereon. A pair of cylindrical rollers interposed respective ones of the pair of reaction surfaces and the pair of recessed portions. The attachment system or turbine assembly provides an economical, reliable and effective attachment of a component having a preestablished rate of thermal expansion to a component having a greater preestablished rate of thermal expansion.

Abstract: A turbine compressor or turbine bladed disc assembly wherein the blades 6 include root portions 8 provided with opposed converging surfaces 9,10 and the disc 14 includes slots 11 provided with opposed diverging surfaces 12,13. In assembly the root portion 8 of a blade 6 is engaged in a slot 11 with abutting converging and diverging surfaces and the blade 6 is oscillated relative to the disc whilst pressure is applied radially to urge the root into close contact with the slot until melting temperature is reached whereupon the oscillatory motion is arrested and the pressure maintained as the converging and diverging surfaces coalesce. A hole 16 may be drilled through the assembled blade and disc assembly in order that communication may be made with cooling air passages in the blade.

Abstract: The present invention provides an improved adjusting shim used in a valve train for an internal combustion engine for an automobile. The adjusting shim produced from a base material consisting of a ceramic material containing 80 to 98 wt. % of silicon nitride and/or sialon and has a porosity of not more than 3%, a bending strength of not less than 1.0 GPa and an impact compressive elastic limit (Hugoniot elastic limit) of not less than 15 GPa, wherein the base material is provided on the surface thereof which contacts a cam with a ceramic surface layer having a composition different from that of the base material and a hardness lower than that of the base material. The adjusting shim of the present invention enables a power loss of a valve train to be minimized; the abrasion resistance thereof to be improved; and the fuel economy, the performance and durability of an internal combustion engine to be improved.