273 results about "Damage tolerance" patented technology

Pre-impregnated composite material (prepreg) is provided that can be molded to form composite parts that have high levels of both strength and damage tolerance without causing any substantial negative impact upon the physical or chemical characteristics of the uncured prepreg or cured part. This is achieved by including in the matrix resin a substantial amount of a multifunctional aromatic epoxy resin that has at least one phenyl group that is meta-substituted.

Sub-arrays such as tiles or chips having pixel elements arranged on a routing layer or carrier to form a larger array. Through-chip vias or the like to the backside of the chip are used for connecting with the pixel elements. Edge features of the tiles may provide for physical alignment, mechanical attachment and chip-to-chip communication. Edge damage tolerance with minimal loss of function may be achieved by moving unit cell circuitry and the electrically active portions of a pixel element away from the tile edge(s) while leaving the optically active portion closer to the edge(s) if minor damage will not cause a complete failure of the pixel. The pixel elements may be thermal emitter elements for IR image projectors, thermal detector elements for microbolometers, LED-based emitters, or quantum photon detectors such as those found in visible, infrared and ultraviolet FPAs (focal plane arrays), and the like. Various architectures are disclosed.

The invention provides a fiber-toughened ceramic-based composite material with a ternary-layer-shaped MAX phase interface layer. A ternary-layer-shaped MAX phase material is introduced to serve as an interface layer, and therefore the anti-irradiation property, the heat-conducting property and the anti-oxidation property of the ceramic-based composite material can be effectively improved; in addition, through multiple fracture energy absorbing mechanisms of the MAX phase material, fracture energy can be effectively absorbed, extension of cracks in the ceramic-based composite material is hindered, and therefore the toughness and the damage tolerance of the ceramic-based composite material are improved. Accordingly, the application fields of the composite material are effectively widened, and the composite material has good application prospects in the fields of aerospace thermal structure materials, nuclear energy structural materials and the like.

A polymer composite structure having resin matrix interlayers reinforced with shape memory alloy (SMA) particles. In one preferred form Nitinol® alloy particles are used for the SMA particles. The Nitinol® alloy particles may comprise cylindrical, oval or spherical shaped particles and are intermixed in the resin of the resin matrix interlayer(s) of the composite structure. The SMA particles provide superelastic, reversible strain properties that significantly improve the damage resistance, damage tolerance (e.g. compression-after-impact (CAI) strength) and elevated temperature performance of the composite structure without negatively affecting the hot-wet compression strength of the composite structure. The polymer composite structure is ideally suited for aerospace and aircraft applications where lightweight and structurally strong materials are essential.

The invention belongs to the advanced composite material preparing technology and relates to a preparation method for carbon nanotube non-woven fabric interlayer modified fiber reinforced composite materials. The method includes: enabling carbon nanotube non-woven fabric to enter resin rich areas of the layers of fiber reinforced thermosetting resin base composite materials in direct intercalation mode, using an autoclave molding process or a liquid forming process, and preparing interlayer the interlayer modified composite materials according to an original curing process of matrix resin. The carbon nanotube non-woven fabric can enter weak interlayer areas of the composite materials through direct intercalation mode, has no influence on flow of the resin in the curing process and a liquid forming process of the matrix resin, and simultaneously a network structure formed by carbon nanotubes in an interlayer can effectively prevent interlayer microcracks from being formed and extending, thereby improving strengthening-toughening performance of interlayer of the composite materials, obtaining high impact damage resistance and high damage tolerance, and covering a temperature range of typical aerospace composite structure application, in particular to a more than 300 DEG C high temperature range.

An extruded structural member having improved damage tolerance containing a base section (6); a stiffening section having at least one pair of structural stiffeners (10), the structural stiffeners integral with the base section (6) and projecting outwardly thereof; and at least one intra-stiffener (90) area positioned between the pair of structural stiffeners (10), the intra-stiffener area (90) having a microstructure with intentionally increased amounts of fiber texture to reduce the rate of fatigue crack growth in the extruded structural member.

Pre-impregnated composite material (prepreg) is provided that can be cured/molded to form composite parts having high damage tolerance and interlaminar fracture toughness. The matrix resin includes a thermoplastic particle component that includes a mixture of polyamide 12 particles and polyamide 11 particles.

The invention relates to a high-damage tolerance type ultrahigh strength aluminum alloy and a preparation method thereof. The high-damage tolerance type ultrahigh strength aluminum alloy comprises the following components in percentage by weight: 8.5-10.0 percent of Zn, 1.0-2.5 percent of Mg, 1.0-2.0 percent of Cu, 0.06-0.20 percent of Zr, 0.02-0.05 percent of Ti, not more than 0.08 percent of Fe, not more than 0.06 percent of Si and the balance of Al and unavoidable impurities, wherein the Fe content is larger than Si content. The method comprises the following steps: firstly, smelting and casting; then carrying out homogenization on an alloy casting blank by stages, carrying out hot rolling on the homogenized casting blank and carrying out solid dissolution on a hot rolled plate; prestretching the plate in four hours after quenching; and finally immediately carrying out double-stage ageing processing on the prestretched plate. The obtained alloy has higher strength and damage tolerance and is suitable for being applied in an airplane part exposed in the atmospheric environment for a long time.

A vane assembly 10 suitable for a turbine engine features a refractory vane 12 with an internal cavity 20 and a pair of flexible metallic baffles 26 extending into the cavity from spanwisely opposite ends of the vane. A rigid fastener 48, such as a nut and bolt assembly applies a tensile load to the baffles. The tensile load is reacted out as a compressive load applied to the vane. In another embodiment, the baffle is relatively rigid but the fastener is flexible. The compressive loading exerted on the vane counteracts the brittleness customarily exhibited by refractory materials and imparts damage tolerance to the vane. The arrangement also allows the use of a metal baffle that can be easily secured to the vane and dispenses with any need for a potentially troublesome seal between the baffles and the spanwise extremities of the vane.

The invention discloses a low-cost high-efficiency forming process of a fiber reinforced thermoplastic composite material, comprising the following steps of: arranging a workpiece to be formed, which is prepared by laying fiber reinforced materials, in a closed space containing a mold template; laying demolding cloth and a diversion grid in sequence on the workpiece; laying or coating a demolding material under the workpiece; arranging an impregnating pipe on one side of the width direction of the workpiece; arranging a vacuum pipeline on the other side along the length direction of the workpiece and parallel to the side surface; arranging a breathable material between the vacuum pipeline and the workpiece; then pumping vacuum in the closed space through the vacuum pipeline, simultaneously raising the temperature of the mold template to the impregnating temperature of 50-250 DEG C; and finally filling molten thermoplastic resin or the precursor of the thermoplastic resin of which the temperature is 50-250 DEG C and the viscosity is 0.15-1.0Pa.s into the workpiece from the impregnating pipe, and then insulating for polymerization or reducing the temperature for cooling and forming. The forming process needs not to prepare a pre-impregnated belt in advance and has low cost and high efficiency, and the workpiece can be formed easily and has superior toughness and damage tolerance.

The invention pertains to airplane fatigue and damage-tolerance testing technologies, and relates to a damage-tolerance testing method for the whole wing spar of an airplane. The method comprises the steps of: (I) choosing a part with large load from the whole wing spar for serving as a testing part examining segment, (II) producing two same testing parts according to the examining segment, (III) reserving an interval of 20 to 30mm between the two back-to-back testing parts which are respectively provided with a coating in the upper part and the lower part (the coating is connected with an upper spar edge strip and a lower spar edge strip to form a case segment), (IV) fixing the root of the testing case segment completing assembly on a load-bearing wall, (V) exerting concentrated load at one end of the testing parts to simulate the bending moment of the spar at the testing segment and the shearing force of a web, without considering the influence of the shearing flow of the coating and (VI) using a finite element method to conduct damage tolerance analysis on the testing parts and comparing testing results. The method solves the difficulty that the crack expansion and remaining strength of the present airplane whole wing spar structure lack theoretical evidence; and the design of the testing parts adopts a method of simultaneously conducting tests to the two back-to-back testing parts, thereby eliminating the distortion of a single spar.

A novel materials technology has been developed and demonstrated for providing a high modulus composite material for use to 1000 DEG F. and above. This material can be produced at 5-20% of the cost of refractory materials, and has higher structural properties. This technology successfully resolves the problem of "thermal shock" or "ply lift," which limits traditional high temperature laminates (such as graphite/polyimide and graphite/phenolic) to temperatures of 550-650 DEG F. in thicker (0.25'' and above) laminates. The technology disclosed herein is an enabling technology for the nose for the External Tank (ET) of the Space Shuttle, and has been shown to be capable of withstanding the severe environments encountered by the nose cone through wind tunnel testing, high temperature subcomponent testing, and full scale structural, dynamic, acoustic, and damage tolerance testing.

The invention discloses a high-efficiency multifunctional foliage fertilizer and a production method thereof. The foliage fertilizer is prepared by chelating shrimp processing leftovers, which serve as the main raw material, with medium and minor elements. The foliage fertilizer is prepared from 10-30% of amino acids, 2-8% of medium and minor elements, 10-30% of organic nitrogen, 3-20% of phosphorus, 3-20% of potassium, and 10-30% of shrimp peptide protein, chitosan, chitosan oligosaccharide, chitin, glucosamine hydrochloride, astacin and shrimp cephalin. The foliage fertilizer disclosed by the invention enhances the low-temperature tolerance, drought tolerance, salt damage tolerance, chemical damage tolerance and other stress tolerances of crops on the premise of comprehensively and efficiently supplementing necessary nutrients for crops, and has the functions of restoring the wound in the root system, enhancing the disease resistance of the crops, inhibiting disease and pest and the like. The production technique can reduce the acidolysis consumption, avoids introducing abundant inorganic salts and chloride ions to destroy soil ecological equilibrium and influence plant growth, lowers the energy consumption and requirements for equipment strength, and can implement the clean production of complete utilization and zero pollution discharge.

The present invention relates to an extruded, rolled and/or forged product made of an aluminium alloy. Alloys of the present invention may comprise (by mass):

Zn 6.7-7.5% Cu 2.0-2.8% Mg 1.6-2.2%

at least one element selected from the group composed of:

Zr 0.08-0.20% Cr 0.05-0.25% Sc 0.01-0.50%

Hf 0.05-0.20% and V 0.02-0.20%

Fe+Si<0.20%

other elements ≦0.05 each and ≦0.15 total, balance aluminium. Products of the present invention in some embodiments have an improved compromise between static mechanical strength and damage tolerance.

The invention relates to an aircraft fuselage panel test device, in particular to a fuselage panel combined load test device. The fuselage panel combined load test device comprises an axial load applying assembly (100), a shear load applying assembly (200), an air pressure and circumferential load balance assembly (300) and a counterweight assembly (400), wherein single load can be applied among the assemblies, and the assemblies can be freely combined to apply load. Stretching/compressing, shearing and airtight load can be applied to a fuselage panel at the same time for static force and fatigue experiments, and static force, fatigue and damage tolerance tests related to the fuselage panel can be performed, so that the design level is promoted, and the design period is shortened.

The invention relates to a weave method of 2.5-dimensional overall braided multi-through pipe fabric. The weave method is characterized in that (1) the structure of the fabric adopts a 2.5-dimensional shallow-cross bended-connection structure; (2) the overall weaving of the fabric adopts core imitation-shaped weaving to ensure the internal molded surface dimension of the fabric; (3) the fabric is divided into two parts, namely a connecting body part and a through pipe part, wherein the connecting body is a part connected with the through pipes, while the through pipes are connected through the same connecting body; (4) the lower section of each through pipe is provided with warp yarns of the part in advance; (5) weft yarns between through pipes in the same plane are lined between the warp yarns layer by layer through adopting a weft insertion process; (6) when the horizontal section of the connecting body increases, warp yarns are increased through adopting a shift yarn adding method, and the method can effectively improve the hole phenomenon in yarn-adding points; and (7) when the horizontal section of the connecting body decreases, warp yarns are reduced through adopting a shift yarn reduction method. The 2.5-dimensional multi-through pipe fabric woven by the weave method has the advantages of better overall property, controllable shaping, high damage tolerance, excellent performances of impact resistance, delamination resistance, fatigue resistance, and the like.

A method is disclosed utilizing off the shelf constant cross section thickness sandwich panels comprised of Fiber Reinforced Thermoplastic (FRTP) resin skins and low density thermoplastic (TP) core material wherein the steps of selectively and controllably exposing the panels to heat and incrementally thermoforming the skin-core into a consolidated composite edge or intra-panel area in consideration of subsequent mating and attachment of the FRTP sandwich panel to other structures is achieved. The exact configuration of articles so thermoformed is design optimized to overcome manufacturing, assembly, weight, in-service and structural performance shortcomings of prior art and FRP sandwich panel structures. Further disclosed is an improved, load-bearing, modular design container structure assembled from such thermoformed FRTP sandwich panels in which is utilized the unique core-skin edge configuration of the present invention in consideration of improved: load bearing performance, useful load volume, reduced manufacturing costs, structural weight savings, impact and damage tolerance and repair and replace issues.

The invention relates to a method for improving the damage-tolerance property of a 2,000-type aluminium alloy plate material. In the method, preferable hot-rolling outlet temperature, high-temperature softening treatment and smaller cold-rolling deformation in softening treatment process can be controlled, so that intercrystalline energy storage of the material is reduced, recrystallization degree is controlled and a malleableized texture is formed. By the method, the service life of the 2,000-type material such as 2,024, 2E12 and the like is prolonged by 2 to 4 times; the crack propagation rate is reduced to 1/4 to 1/2 of the original crack propagation rate; the production efficiency is improved; and the energy is further saved.

The present invention is weaving process of three-dimensional conic-casing fabric and its product. The weaving process includes the following steps: 1. pre-setting warp yarns based on the fabric structure and in tubular fabric weaving process; 2. introducing weft yarns in planar helical wefting method and interweaving with preset warp yarns; 3. setting warp cutting points on two sides based on the planar shape of the fabric and performing copying warp cutting operation; and 4. repeating the steps 2 and 3 until finishing the weaving of the three-dimensional conic-casing fabric. The present invention has short weaving period, low cost, simple operation, high efficiency, high adaptability and other features, and the fabric has high integrality, high damage tolerance, high impact resistance and other advantages.

Aircraft fuselage section (32) subjected to impacts of external bodies, the aircraft fuselage having a curved shape with at least a vertical symmetry plane (A-A) and a central longitudinal axis and comprising a skin (35) and a plurality of frames (37) arranged perpendicularly to said longitudinal axis (33), the aircraft fuselage section also comprising at least an inner reticular structure (51, 53) mounted on a supporting structure (41, 43) comprising longitudinal beams (39) attached to the skin (35) and interconnected with said frames (37), said inner reticular structure (51, 53) being arranged for creating at least one closed cell (75) with the skin (35) for improving its resistance and its damage tolerance to said impacts. Said inner reticular structure (51, 53) can be formed by panels (61), rods (65, 65′), cables (67, 67′) or belts (69, 69′).

The invention belongs to a technology for preparing composite materials and relates to a method for preparing rigid three-dimensional crystal whisker interlaminar modified continuous fiber composite materials. The interlaminar modified thermosetting resin composite materials are prepared by the following method by adoption of a micron-scale heteropical rigid micro 3-3 structure and adoption of a four-pin zinc oxide crystal whisker as an interlaminar reinforcement: firstly, the four-pin zinc oxide crystal whisker which is subjected to vacuum drying for 2 hours at a temperature of 100 DEG C is deposited on the surface of fabrics by means of a mechanical vibrating screen or electrostatic adherence or deposition through a fluidized bed or powdering through a drum-type silk screen, and modified fabrics are obtained; and secondly, after the modified fabrics are laid as required, the resin transfer molding technology or the resin membrane permeation technology is utilized for preparing the composite materials by the prior matrix resin solidifying technology. The method can improve the strong interlaminar toughening performance of the composite materials, obtain high surge impedance and high damaged tolerance, and cover the temperature range of structural application of typical aerospace composite materials and particularly cover the high temperature range of more than 300 DEG C.