Abstract: A rotor vane for an aircraft turbine engine, the vane having an axis of rotation once it has been rigidly connected to a rotor and a stacking axis. The vane includes a blade extending between an internal platform and an external platform bearing at least one projecting lip. The external platform is configured to cooperate in a form-fitting manner with the complementary side edges of adjacent vanes. The rotor vane has wear-resistant covering of a lower side edge which extends over one wall of a substantially rectilinear first ridge of the platform and over one wall of a second ridge of the platform. The second ridge extends at least partially inside the lip and is inclined relative to the first ridge in a direction substantially parallel to a transverse axis of the lip.

Abstract: A bellows includes a flexible section including a series of convolutions extending between a first end and a second end. The convolutions are formed of a nano-grained aluminum alloy.

Abstract: Pulsed detonation engines (PDEs), or various components thereof, such as the detonation chamber and/or nozzle, can be economically constructed from materials having low thermal stability, such as plastics, composites, and light metals. During operation, the intermittent injection and detonation of reactants produces a motive force (e.g., thrust) over relatively short intervals useful in orbit control and the like. The relatively short intervals of operation prevent temperatures of the PDE components from reaching temperatures that would result in their dimensional failure.

Abstract: A method of making heat engine components, such as rocket nozzle throats, includes forming a body having at least one surface, wherein the body is made of a refractory metal, and compressing a ceramic powder between a die and the at least one surface of the body, with sufficient heat and pressure to densify and thus form a solid ceramic coating on the at least one surface of the body. Nozzle throats made according to the invention have bodies made of refractory metals, such as tantalum, and ceramic coatings on the inner surfaces of the annular bodies, wherein the ceramic coatings are made of ceramic materials such as tantalum carbides (TaC and Ta2C).

Abstract: A thrust chamber assembly for liquid fueled rocket engines and the method of making it wherein a two-piece mandrel wrapped with a silica tape saturated with a phenolic resin, the tape extending along the mandrel and covering the combustion chamber portion of the mandrel to the throat portion. The phenolic in the tape is cured and the end of the wrap is machined. The remainder of the mandrel is wrapped with a third silica tape. The resin in the third tape is cured and the assembly is machined. The entire assembly is then wrapped with a tow of graphite fibers wetted with an epoxy resin and, after the epoxy resin is cured, the graphite is machined to final dimensions.

Abstract: The present invention concerns an erosion resistant, low signature liner for solid propellant rocket motors and rocket motors having such a low signature liner. The liner combustion products in the exhaust plume have low visibility in the IR and visible light spectra which is essential for use in low signature solid propellant rocket motors. The liner also forms an in-situ protective layer at elevated temperatures for enhanced erosion resistance. These erosion resistant properties eliminate the need for supplemental case insulation which would contribute additional smoke and combustion products to the exhaust plume as well as add additional costs and inert weight to the motor.

Abstract: A combustor liner includes an inner layer for facing combustion gases, and an opposite outer layer for facing a cooling fluid. The outer layer has a greater coefficient of thermal conductivity than the inner layer for reducing temperature gradients in the liner. In a preferred embodiment, the outer layer significantly reduces temperature gradients in the liner which are caused by the varying cooling ability of impingement cooling air jets for more uniformly cooling the combustor liner and reducing the maximum temperature thereof.

Abstract: A compressor for a gas turbine, comprises a rotor including a plurality of rotor parts, and blades for a plurality of stages, e.g., twelve stages, implanted in the rotor parts, the rotor part in which the blades for at least two of the stages are implanted is made of a Ni--Cr--Mo--V low alloy steel, and the Ni--Cr--Mo--V low alloy steel has substantially an entire bainite microstructure comprising, in weight percent, C of 0.15-0.40%, Si not more than 0.1%, Mn not more than 0.5%, Ni of 1.5-2.5%, Cr of 0.8-2.5%, Mo of 0.8-2.0%, V of 0.1-0.35% and the balance being substantially Fe. Blades for at least two of the stages are implanted in one of the rotor parts.

Abstract: A nickel base superalloy, having either columnar or equiaxed grain structure, which has significantly improved resistance to hydrogen embrittlement, and to fatigue in air is disclosed. The superalloy consists essentially of, in weight percent, about 0.006-0.17 carbon, about 6.0-22.0 chromium, up to about 17.0 cobalt, up to about 9.0 molybdenum, up to about 12.5 tungsten, up to about 5.0 titanium, up to about 6.7 aluminum, up to about 4.5 tantalum, up to about 2.5 hafnium, up to about 18.5 iron, up to about 3.25 rhenium, up to about 1.25 columbium, remainder nickel. The microstructure of the superalloy consists essentially of a plurality of fine, discrete carbide particles, .gamma.' precipitates in a .gamma. matrix, and is essentially free of script carbides, .gamma./.gamma.' eutectic islands and porosity.

Abstract: A stiffened composite is provided exhibiting multi-directional reinforcement with continuous long fibers. The composite has a substrate and serpentine stiffeners adhered thereto. The stiffeners merge toward, join with and diverge from adjacent stiffeners to provide a stiffener matrix on the substrate. The serpentine stiffener permit the matrix to exhibit multi-directional reinforcement with the continuous long fibers within the serpentine ribs. The composite structures are useful as aircraft bypass ducts, and as other composite products.

Abstract: A composition consisting essentially of a polyphosphazene binder having the following repeating units: ##STR1## wherein n is from about 20 to about 50,000, A is selected from --O-- and --NH--, and R.sup.1 and R.sup.2 are independently selected from alkyl having from one to about 20 carbon atoms, aryl having from 6 to about 14 carbon atoms, alkaryl having from 7 to about 30 carbon atoms, any of the preceding moieties substituted with halogen, hydroxyl, alkoxy, aryloxy, or nitro groups, or any of the preceding moieties linked by ether or amine linkages. Combined with an organic fiber filler, this material provides an erosion resistant insulation composition for temporarily protecting the interior of a rocket motor combustion chamber from thermal damage while the rocket is firing.

Abstract: A combustor liner includes an inner layer for facing combustion gases, and an opposite outer layer for facing a cooling fluid. The outer layer has a greater coefficient of thermal conductivity than the inner layer for reducing temperature gradients in the liner. In a preferred embodiment, the outer layer significantly reduces temperature gradients in the liner which are caused by the varying cooling ability of impingement cooling air jets for more uniformly cooling the combustor liner and reducing the maximum temperature thereof.

Abstract: A wear reduction kit for a combustion system of a gas turbine includes:a) a plurality of U-shaped inserts for use with a gas turbine transition piece having a peripheral end frame at one end thereof, the peripheral end frame having a continuous peripheral slot therein, open in a radially outward direction wherein the plurality of U-shaped inserts are adapted for insertion within the peripheral slot, and further wherein the U-shaped inserts are formed of material having a hardness greater than the end frame;b) at least a pair of guide block elements for use in securing a transition piece of a gas turbine combustor to a transition piece support, wherein each of the guide block elements is secured to the transition piece and adapted for engagement with cooperating guide finger elements secured to the support; and at least a pair of covers for the guide finger elements, and wherein the guide blocks are formed of material having a hardness greater than the finger element covers; andc) cooperating combustor flow sl

Abstract: The present invention relates to a heat treated, gamma prime precipitation strengthened nickel base alloy having an improved resistance to hydrogen embrittlement, particularly crack propagation. The alloy has a microstructure which is essentially free of script carbides, gamma--gamma prime eutectic islands and porosity. The microstructure further includes a plurality of regularly occurring large barrier gamma prime precipitates and a continuous field of fine cuboidal gamma prime precipitates surrounding the large barrier gamma prime precipitates.

Abstract: Copper alloy rocket engine combustion chamber linings have been found to deteriorate when exposed to cyclic reducing/oxidizing (redox) environments which are a consequence of the combustion process. This deterioration, known as blanching, can be characterized by increased roughness and bum through sites in the wall of the combustion chamber lining and can seriously reduce the operational lifetime of the combustion chamber. The blanching problem can be significantly reduced by depositing a thin layer of Cu-30.sup.v /.sub.o Cr (a copper matrix with 30.+-.10 volume percent of chromium) on the inside wall of the combustion chamber. The microstructure of the Cu-30.sup.v /.sub.o Cr coating consists of finely distributed chromium (Cr) particles in a copper (Cu) matrix. When exposed to an oxidizing environment at high temperatures, the coating forms a protective chromium scale which is stable in hydrogen atmospheres (e.g., substantially unreduced by high pressure hydrogen). The Cu-30.sup.v /.sub.

Abstract: A high-temperature gas turbine includes a compressor, and a turbine coupled to the compressor, the turbine being capable of rotating at a high speed by a gas of firing temperature generated by a combustor. The compressor includes a combined type rotor shaft and blades of at least 15 stages which are implanted in the combined type rotor shaft. The compressor compresses air to a temperature of 400.degree. C. or above. The combined type rotor shaft is made of Ni--Cr--Mo--V low alloy steel having all bainitic microstructure whose composition consists, as analyzed in weight percent, of 0.15-0.40% C, 0.1% or less Si, 0.5% or less Mn, 1.5-2.5% Ni, 0.8-2.5% Cr, 0.8-2.0% Mo, 0.1-0.35% V and the balance substantially of Fe.

Abstract: A casing assembly for a ducted fan gas turbine engine comprises a casing around which extend two annular, axially spaced apart, frangible rails. Pieces of cloth woven from an aromatic polyamide fibre are positioned on the casing radially outer surface between the rails. Continuous layers of the cloth are wound around the casing radially outwardly of the rails and cloth pieces. The arrangement provided enhanced containment of any fan blades of the engine which may become detached from the gas turbine engine which carries the casing assembly.

Abstract: An improved rocket motor and a method for its manufacture are characterized by the provision of a thermoplastic elastomer binder in the outer casing, the insulation layer, and the propellant. When the motor is heated to the melting point of the binder, the bondlines between the propellant and insulation and between the insulation and casing are simultaneously fused to form a unitary motor structure.

Abstract: Self venting ablative carbon or graphite composite materials are prepared by interspersing low temperature degradable thread within the woven carbon or graphite cloth prior to impregnation with a phenolic or resin matrix. The degradable threads do not degrade during composite curing, but do degrade at a temperature less than the thermal decomposition (char) temperature of the cured composite material. Thus, as the threads decompose, they leave pores throughout the cured composite which are available to vent gaseous decomposition products produced as the resin matrix approaches the char temperature. Venting the resin decomposition products reduces internal pressure within the composite material thereby reducing the likelihood of spalling and ply lifting.

Abstract: A heat resistant steel of the present invention contains 0.05 to 0.2 wt. % of C, less than 0.5 wt. % of Si, less than 0.6 wt. % of Mn, 8 to 13 wt. % of Cr, 1.5 to 3 wt. % of Mo, 2 to 3 wt. % of Ni, 0.05 to 0.3 wt. % of V, 0.02 to 0.2 wt. % in total of either or both of Nb and Ta, 0.02 to 0.1 wt. % of N and the balance substantially Fe. Since a gas turbine of the present invention is constituted by members, such as discs, blades, shafts and so forth, made of alloys of this kind, the gas turbine has a structure in which it is possible to achieve a high level of creep rupture strength and Charpy impact value.

Abstract: An axial flow compressor for a bypass type jet engine has a low pressure stage which includes a one-piece, annular hoop shroud mounted to the airfoil of each of the rotor blades in the low pressure stage to divide the airflow entering the compressor into an outer, bypass airflow path and an inner airflow path directed to the higher pressure stages of the compressor. The hoop shroud is formed of circumferentially wound, parallel fibers such as silicone carbide fibers embedded in a metal matrix material such as titanium, wherein the volume ratio of the hoop shroud is approximately one-third fibers and two-thirds metal matrix material.

Abstract: The coated article includes a refractory metal substrate having an oxidation resistant intermetallic layer formed in situ thereon; e.g., a fused refractory metal silicide layer, and a ceramic layer applied on and adhering to the intermetallic layer to provide enhanced high temperature properties and improved resistance to premature catastrophic failure in high temperature oxidizing environments where dissimilar materials are present that may destructively react with the inermetallic layer.

Abstract: Composite polar bosses applied in rocket motor cases have very desirable characteristics. The lead time for a boss can be reduced from 6-12 months to 4-6 months. Weight savings for the boss/nozzle assembly are about 20-40 percent. A composite polar boss which attaches nozzle assemblies to solid fueled rocket motor cases is disclosed. This polar boss is a carbonized fabrication which sits within a circumferential indentation within the motor case, and has a threaded inner circumference which permits the nozzle assembly to be attached thereto. The materials of the boss are selected to permit a service temperature of up to 3,200 degrees F.

Abstract: A heat resistant steel of the present invention contains 0.05 to 0.2 wt. % of C, less than 0.5 wt. % of Si, less than 0.6 wt. % of Mn, 8 to 13 wt. % of Cr, 1.5 to 3 wt. % of Mo, 2 to 3 wt. % of Ni, 0.05 to 0.3 wt. % of V, 0.02 to 0.2 wt. % in total of either or both of Nb and Ta, 0.02 to 0.1 wt. % of N and the balance substantially Fe. Since a gas turbine of the present invention is constituted by members, such as discs, blades, shafts and so forth, made of alloys of this kind, the gas turbine has a structure in which it is possible to achieve a high level of creep rupture strength and Charpy impact value.

Abstract: Combustion chamber for rocket motor of the type comprising two domes, interconnected by a cylindrical shell, each of said domes and the intermediate shell being produced from a filamentary composite material, in which each dome comprises a slightly conical skirt of which the external face is joined over at least part of its length to the inner face of the shell, via a split cylindro-conical connection ring and at least opposite the conical skirt of the dome, one conical face having the same conicity as the conical skirt, the assembling of the three elements to be connected, i.e. dome, shell and ring, being achieved by adhesive bonding.

Abstract: In the cooling wall of a regenerative cooling rocket combustion chamber, the throat portion thereof in which the heat load is highest is formed of a porous material having a perspiring cooling action. The other portion in which the heat load is relatively low is designed to be cooled by conventional forced convection. By such a composite cooling system, a high cooling effect is provided and yet the pressure loss in the cooling passages is small and light weight and long life of the engine are attained concurrently.

Abstract: A solid fuel rocket engine comprises a casing which defines a combustion chamber with a solid propellant charge in the combustion chamber and an insulating layer of a water containing alkali silicate and fibers disposed between the casing and the charge. The layer is a type which will froth at a predetermined temperature which occurs during burn off of the propellant charge so that it increases in its thickness dimension in the area of the hot gases which develop and provides an adequate insulation. A method of producing the layer comprises heating the alkali silicate layer to a temperature in which it becomes plastic and sheathing the propellant charge with the plastic layer. Thereafter the sheath charge is inserted into the combustion chamber casing leaving an annular gap between this layer and combustion chamber casing. Thereafter heat is applied to the vicinity of the chamber to cause the layer to froth and expand to close the annular gap.

Abstract: A method of making a rocket motor case with an integral nozzle in which the rocket motor has a large length to diameter ratio and is made by applying a linear wrap of fibers as a first layer on a mandrel and then by applying radial wraps over said linear wrap and also conforming the linear fibers to the shape of the nozzle after applying the linear fibers on said mandrel to provide a structure that has sufficient stiffness to accomplish the desired result of a large length to diameter rocket motor case.

Abstract: A gas generator outlet manufactured of different metals, each possessing ferent thermal capability, welded together with a tapered internal insulation system. The generator outlet is designed to contain a high-pressure gas having a flame temperature of 3000.degree. F. The external surface of the gas generator outlet remains at or below 600.degree. F. while the outlet interfaces with a metal manifold having a temperature of approximately 3000.degree. F.

Abstract: A gas turbine propulsion engine for aircraft of the type in which thrust is provided only from flow of exhaust gases or a combination of flow of exhaust gases and by-pass air, the engine having a cone-shaped spinner with a solid pointed tip portion and a main body portion. The tip portion has at least a surface made of a different material from the substantially rigid material of the main body portion, the at least surface of the tip portion being flexible relative to the body portion at operational rotations of the spinner whereby ice is automatically shed from the spinner due to cracks being developed at the bond between the tip portion and the main body portion. The cracks in the ice developed at the bond are propagated by the surface deflection of the remainder of the tip portion causing any accretion of ice thereon to shed from the remainder of the spinner.

Abstract: A thrust nozzle for a reaction engine having a combustion chamber with a sustained flight discharge nozzle defined therein for the passage of thrust gases and a sustained flight thrust arrangement connected to the combustion chamber for supplying thrust gases through the combustion chamber, the thrust nozzle comprising a starting nozzle formed on an interior wall of the combustion chamber radially inwardly of the sustained flight discharge nozzle made of a plurality of layers of ablatable material which are ablated by the passing thrust gases to discharge vaporized or fragmentary parts of the layers of ablatable material.