112results about "Fuselage insulation" patented technology

A structure system for installation into a fuselage structure in an interior of an aircraft is proposed, the system comprising at least one first carrier element, at least one second carrier element, at least one elongate frame element having a first end and a second end, at least one first attachment device for attaching the first carrier element to the fuselage structure or to an element coupled with the fuselage structure, at least one first coupling device mounted to the first end for coupling the first carrier element to the at least one frame element, and at least one second coupling device mounted to the second end for coupling the second carrier element to the at least one frame element, wherein the at least one frame element is adapted for holding at least one component to be installed in the interior, wherein the at least one first carrier element is designed for coupling with the at least one first coupling device, wherein the at least one second carrier element is designed for coupling with the at least one second coupling device, and wherein one of the at least one first coupling device and the at least one second coupling device is adapted to snap into the respective carrier element, to be moved relative to the respective end of the frame element and to be arrested in a desired relative position thereto.

The invention relates to the technical field of layered products, and particularly discloses a heat-insulating and wave-absorbing skin material as well as a preparation method and application thereof. The heat-insulating and wave-absorbing skin material is of an integrated layered structure based on an aluminum oxide-mullite system, comprising a functional gradient layer with double gradients of fiber density and aperture, an impedance matching layer with a gradient dielectric constant, a resonance wave absorbing layer with a periodic dot matrix formed by silicon carbide and aluminum oxide core-shell microspheres, and a heat insulation layer with a nano-micron double-scale pore structure. According to the preparation method, a green body is constructed mainly through tape casting and silk-screen printing, and then co-sintering integrated forming is carried out. The material has the excellent characteristics of high temperature resistance, toughness, efficient heat insulation, broadband wave absorption and the like, and is suitable for extreme environments such as hypersonic aircraft skin and the like.

The application discloses an aircraft cargo hold bottom structure. The aircraft cargo hold bottom structure comprises a machine body, a floor, a heat insulation layer and a connecting piece. The floor is located in the machine body and connected with the machine body, and the floor and the machine body enclose a first cabin. The heat insulation layer is located in the first cabin and connected with the floor through the connecting piece and is spaced apart from the machine body, so that the heat insulation layer is not directly attached to the machine body, the heat insulation layer can avoid absorbing water on the machine body, thereby slowing down aging of the heat insulation layer, and water on the machine body can be normally evaporated and removed, the machine body can be prevented from being in a humid state frequently, and thus corrosion of the machine body is reduced.

The invention provides an aircraft composite heat insulation plate based on welding forming, a flap lower wall plate and an assembling method, and the aircraft composite heat insulation plate based on welding forming comprises an upper frame and a lower frame which are sequentially stacked to form a frame assembly; the outer edge of the outer skin and the frame assembly are aligned and stacked, the lower surface of the outer skin abuts against the lower frame, and the upper frame, the lower frame and the outer skin are formed at a time in a welding mode; the aerogel layer is arranged on the lower surface of the outer skin and located in the frame assembly. The invention has the beneficial effects that the nano porous structure of the aerogel limits the motion of gas molecules and reduces convective heat transfer, and the complex nano skeleton increases the heat conduction path, improves the heat radiation absorption, reduces the solid heat conduction and radiative heat transfer, and can tolerate the high temperature of 650 DEG C in the jet flow area of an engine.

The invention relates to a sound absorption structure and an aircraft. The sound absorbing structure comprises at least one honeycomb structure (52) comprising at least one partition enclosure (62) fixed in one of the cells (60) of the honeycomb structure (52) and connected to at least one partition (58) delimiting said cell (60), the partition enclosure (62) separates an inner region (ZI) located inside the partition enclosure (62) from an outer region (ZE) located in the unit (60) and outside the partition enclosure (62), the partition enclosure (62) comprising at least one through-hole (70) configured to communicate the inner region (ZI) with the outer region (ZE). This configuration makes it possible to simply define several regions in the same cell, thereby helping to increase the spectrum of attenuating acoustic waves. The aircraft comprises at least one such sound absorbing structure.

The subject matter of the invention is an acoustic absorption structure comprising a unit structure interposed between an acoustic resistive layer and a reflective layer, the unit structure comprising walls delimiting cells, and for at least one cell, the unit structure comprising: at least one and at most three L-shaped partitions, the at least one and at most three L-shaped partitions are impermeable and spaced apart from each other, and define a straight compartment and at least one L-shaped compartment; and an angled, acoustically resistive porous separator in each L-shaped compartment, the angled, acoustically resistive porous separator having a perforation rate of less than or equal to 15%, and dividing the L-shaped compartment into a first region and a second region, thereby forming two types of resonators. This arrangement makes it possible to increase the frequency range of the processed acoustic waves. Another subject of the invention is an aircraft comprising at least one such sound absorbing structure.

An acoustic absorption structure including a cellular structure which comprises tubular cells, closed at each of their ends by an acoustically resistive layer and a reflective layer, and at least one individual partitioning system, placed in a single cell and including a partition configured to separate the cell into first and second compartments, at least one opening passing through the partition, several feet independent of the conduit, having a first end connected to the partition and a second end in contact against the reflective layer, and at least one conduit which has a first end rigidly secured to the partition surrounding the opening and a second end spaced apart from the reflective layer.

A swirl recovery vane with acoustic treatment including a leading edge and a trailing edge opposite chordwise from the leading edge; an attachment region opposite spanwise from a tip region; a span dimension extending between the attachment region and the tip region; a chord dimension extending between the leading edge and the trailing edge; a pressure side opposite a suction side of the swirl recovery vane; an acoustic zone integrally formed within the swirl recovery vane. A gas turbine engine includes the swirl recovery vane with acoustic treatment. A process forms the swirl recovery vane with acoustic treatment.

The invention provides a heat prevention and insulation structure of an aircraft base body and the aircraft base body, and belongs to the technical field of aircraft heat prevention and insulation. The heat insulation structure comprises a flexible connecting layer covering the surface of the aircraft base body; the heat insulation units are located on the flexible connecting layer, the first ends of the heat insulation units are connected with the flexible connecting layer, and the heat insulation units are arranged at intervals; the plurality of armor flakes are positioned on the heat insulation units and are in one-to-one correspondence with the plurality of heat insulation units, and the second ends of the heat insulation units are connected with the armor flakes; the reflecting coating covers the upper surfaces and the side surfaces of the armor flakes; the armor flakes on the adjacent heat insulation units are partially overlapped in a one-by-one pressing mode, and the tail portion of the front armor flake in every two adjacent armor flakes is in lap joint with the head portion of the rear armor flake, so that a stacked structure of the flakes is formed. According to the heat prevention and insulation structure, a labyrinth type heat flow transfer path is constructed, heat resistance is increased, and heat insulation efficiency is improved.

The invention relates to an acoustic absorption structure comprising a honeycomb structure (62), interposed between an acoustically resistive layer (64) and a reflective layer (66), which includes walls (68) delimiting cells (70) and, for at least one cell (70),: - at least one and at most three L-shaped partitions (72, 72'), impermeable and spaced apart, which delimit a straight compartment (70.1) and at least one L-shaped compartment (70.2, 70.3), - in each L-shaped compartment (70.2, 70.3), an inclined acoustically resistive porous partition (80, 80') which has an openness ratio less than or equal to 15% and divides the L-shaped compartment (70.2, 70.3) into first and second zones (80.1, 80.2, 80.1', 80.2'). forming two types of resonators. This arrangement makes it possible to increase the frequency range of the processed acoustic waves. The invention also relates to an aircraft comprising at least one such acoustic absorption structure.

The application relates to the technical field of fairings and thermal protection of hypersonic aircraft, in particular to an aircraft antenna cover connecting structure. The aircraft antenna cover connecting structure comprises an antenna cover, nuts and heat insulation screws corresponding to the nuts. The antenna cover comprises a connecting end for connecting with a rear cabin section. An inner wall of the connecting end is circumferentially spaced apart with at least three nut holes. A bolt hole in communication with the nut holes is further arranged in the connecting end. The nuts corresponding to the nut holes are arranged in the nut holes. The heat insulation screws corresponding to the nuts are arranged to pass through connecting holes on the rear cabin section and bolt holes on the connecting end to connect with the nuts, so as to connect the rear cabin section and the connecting end. The application can solve the problem that the adhesives in the prior art have the risk of falling off even if the temperature of the adhesives is improved.

The application belongs to the field of acoustics, and discloses a low-frequency and wide-band noise reduction acoustic structure and a design method of the acoustic structure. The acoustic structure comprises a plurality of acoustic structural members and an embedded member. At least one embedded member is arranged in the plurality of acoustic structural members. On the same embedded member, the acoustic structural members in communication with the two open ends of the embedded member are arranged in pairs and jointly form a first unit. The acoustic structural member not in communication with the embedded member is a second unit. The first unit and the second unit are arranged in parallel. Compared with the prior art, when sound waves pass through the acoustic structural members arranged in pairs, the coupling resonance generated in the acoustic structural members can widen the noise reduction frequency band at a certain resonance frequency in the low-frequency range. The added embedded member can generate a new noise reduction frequency band. The weak coupling effect between the parallel units can further widen the noise reduction frequency band. Finally, the low-frequency and wide-band noise reduction effect is achieved.

The invention relates to a sound-absorbing structure comprising a honeycomb structure (32) which includes tubular cells (40), sealed at each end by an acoustically resistive layer (34) and a reflective layer (36), and at least one individual partitioning system (42), positioned within a single cell (40) and comprising: - a partition (44) configured to separate the cell (40) into first and second compartments (40.1, 40.2), - at least one orifice (46) passing through the partition (44), - several feet (48) independent of the conduit (54), which has a first end (48.1) connected to the partition (44) and a second end (48.2) in contact with the reflective layer (36), - at least one conduit (54) which has a first end (54.1) integral with the partition (44) surrounding the orifice (46) and a second end (54.2) spaced from the reflective layer (36).

A swirl recovery vane with structurally capable acoustic treatment includes a leading edge and a trailing edge opposite chordwise from the leading edge; an attachment region opposite spanwise from a tip region; a span dimension extending between the attachment region and the tip region; a chord dimension extending between the leading edge and trailing edge; a pressure side opposite a suction side of the swirl recovery vane; a structural feature formed within the swirl recovery vane extending at least one of spanwise through the swirl recovery vane between the attachment region and tip region and / or chordwise between the leading edge and the trailing edge; an acoustic receiver formed within the structural feature extending at least one of spanwise through the swirl recovery vane between the attachment region and the tip region or chordwise between the leading edge and the trailing edge; and an acoustic panel inserted into the acoustic receiver. A gas turbine engine includes the swirl recovery vane with structurally capable acoustic treatment. A process forms the gas turbine engine with the swirl recovery vane with structurally capable acoustic treatment.

The application belongs to the field of protective coating, and particularly discloses a composite laser protective coating and a preparation method thereof. The protective coating comprises, in sequence, a high-temperature oxidation-resistant layer, a ceramic thermal insulation layer, a ceramic reflective layer and a metal reflective layer which are arranged on a substrate in a laminated manner. The porosity of the ceramic thermal insulation layer is not less than 10%. The material of the high-temperature oxidation-resistant layer comprises at least one of NiCrAlY, CoCrAlY and NiCoCrAlY. The material of the ceramic thermal insulation layer comprises zirconium oxide and an oxide doped in the zirconium oxide. The material of the ceramic reflective layer comprises zirconium oxide and a rare earth element doped in the zirconium oxide. The protective coating has a low thermal conductivity, and the average temperature on the side of the substrate away from the protective coating is lower than 150 DEG C after the substrate is continuously irradiated by a laser with a power density of 8050 W·cm ‑2 -2 for 300 s, and the bonding strength of the coating is 35-50 MPa.

A vehicle panel includes two sidewalls and a resonator disposed between them on one of the cell walls. The resonator includes a network of cells, each cell defines an internal space to absorb and dissipate noise and vibrations within a specific frequency range. Certain cells or subsets of cells may be interconnected to define a larger internal cell space. An airgap between the resonator and one of the sidewalls may be partially or completely filled with dampening material to absorb high frequency noise. The dampening material may also provide thermal insulation.

The present application belongs to the technical field of aircraft skin, and relates to an aircraft elastic porous medium skin, a plurality of hole bodies are prefabricated on the skin, an organ pipe type nozzle is installed in the hole body, the organ pipe type nozzle is connected with a jet control system for providing a heat medium for the organ pipe type nozzle; the organ pipe type nozzle comprises an inlet section, a resonance section, an intermediate section and an outlet section connected in sequence; the inlet section is provided with an incidence hole, the resonance section is provided with a resonance hole, the intermediate section is provided with an intermediate hole, and the outlet section is provided with an emission hole which is a diverging hole; the incidence hole, the resonance hole and the intermediate hole are all cylindrical holes, and the diameters of the incidence hole, the resonance hole and the intermediate hole gradually decrease. The elastic porous medium skin is used to induce the slip and micro-cavitation effect of unsteady boundary layer fluid, and the geometric parameters of the control hole are controlled, so that the self-adaptive boundary layer flow state is greatly reduced, the frictional resistance is reduced, and the outer lining has the characteristics of noise reduction, and has the scattering and absorption effect on sound waves.

The utility model relates to aviation aircraft technical field, concretely is an aviation aircraft with floating anti -sinking material quality shell, including aviation aircraft main part, be equipped with fuselage frame body on the aviation aircraft main part, and the fuselage frame body upper end is fixed with fuselage framework, the fuselage framework outside cover is equipped with mutually sealed clamping nose shell, backrest shell and chassis shell, and the inside of fuselage framework, backrest shell and chassis shell is clamped with lining shell respectively, and sets up sealable gas filling layer in lining shell, the bottom of fuselage frame body is equipped with chassis protection ring. The aviation aircraft with floating anti -sinking material quality shell whole frame body adopts lightweight design, and through setting up lining shell with gas filling layer in the nose shell, can provide floating dust sink function under the condition that aviation aircraft self load is not increased, guarantees the buoyancy of aviation aircraft when flying, improves the security of aviation aircraft near water area near execution task.

This invention relates to the field of layered product technology, specifically disclosing a heat-insulating and microwave-absorbing skin material, its preparation method, and its application. The heat-insulating and microwave-absorbing skin material is an integrated layered structure based on an alumina-mullite system, comprising a functionally gradient layer with dual gradients in fiber density and pore size, an impedance matching layer with a gradient in dielectric constant, a resonant microwave-absorbing layer composed of a periodic lattice of silicon carbide@alumina core-shell microspheres, and a heat-insulating layer with a nano- to micron dual-scale pore structure. The preparation method mainly involves constructing a green body through casting and screen printing, followed by co-sintering to form an integrated structure. The material of this invention possesses excellent properties such as high temperature resistance, toughness, high-efficiency heat insulation, and broadband microwave absorption, making it suitable for extreme environments such as hypersonic vehicle skins.

The embodiment of the invention discloses a piezoelectric stack clamping device for active vibration reduction of a bulkhead, relates to the technical field of structural vibration control, and can implement faster active control on the vibration response of the bulkhead on the premise that the structure is simple and the stability is high. The device is characterized in that a pair of bases are installed on the bulkhead in a non-punching mode, and the size and the shape of the bottom of each base are matched with the size and the shape of the bulkhead; one end of the piezoelectric stack actuator is fixedly connected with a base of the clamping device through a bearing seat, and the other end of the piezoelectric stack actuator is in threaded connection with the flexible hinge; the flexible hinge is fixedly connected to the inner side of the other base; the piezoelectric stack actuator serves as a source of excitation force, excitation voltage generates displacement response after being input into the piezoelectric stack actuator, the displacement response is transmitted to the clamping device through the flexible hinge, the displacement response is converted into the excitation force under the limitation of the clamping device, and therefore a pair of force which is equal in magnitude and opposite in direction is output to the base, and excitation is formed.

The present disclosure relates to a spacer (100) for covering a gap between a sidewall (220) arranged in a longitudinal direction and a monument wall (200) arranged in a transverse direction in an aircraft. The spacer comprises an upper mount (111), and a lower mount (113), wherein the lower mount (113) restricts a relative movement between the monument and the spacer in the longitudinal direction, and allows a relative movement in the transverse direction.

An acoustic absorption structure including at least one cellular structure between an acoustically resistive layer and a reflective layer. The cellular structure includes a first face in contact with the acoustically resistive layer, a second face in contact with the reflective layer, and a multitude of cells each discharging at the level of the first and second faces, each cell being delimited by at least one partition, wherein the cellular structure comprises at least one partitioning enclosure positioned in one of the cells of the cellular structure and connected to at least one partition delimiting the cell. The partitioning enclosure separates an interior zone inside the partitioning enclosure and an exterior zone situated in the cell and outside the partitioning enclosure. The partitioning enclosure includes at least one through-orifice configured so that the interior and exterior zones communicate.

The invention relates to a sound-absorbing structure comprising at least one honeycomb structure (52) which includes at least one partition enclosure (62) positioned within one of the cells (60) of the honeycomb structure (52) and connected to at least one partition (58) delimiting said cell (60). The partition enclosure (62) separates an inner zone (ZI) located inside the partition enclosure (62) from an outer zone (ZE) located within the cell (60) and outside the partition enclosure (62). The partition enclosure (62) includes at least one through-hole (70) configured to connect the inner and outer zones (ZE, ZI). This configuration allows for the simple delimitation of several zones within the same cell, thereby increasing the frequency spectrum of the attenuated sound waves.The invention also relates to an aircraft comprising at least one such sound-absorbing structure.

A panel comprises a panel interior surface, a foam layer, and a composite layer. A first end of the panel is oriented at a first end angle relative to a foam layer interior surface, wherein the first end angle is not 90 degrees. A foam layer interior surface is operatively arranged at the panel interior surface. The composite layer comprises a composite layer interior surface operatively engaged with the foam layer exterior surface. An aircraft assembly comprises a fuselage and at least a first panel and a second panel. A first end angle of a first panel first end corresponds to a second end angle of a second panel second end, and the first panel first end operatively engages the second panel second end. A method comprising operatively engaging a first panel interior surface of a first panel with a fuselage exterior surface of a fuselage.

The application discloses a ceramic matrix composite product labyrinth cavity component and a flow guide deposition method, which comprises a horizontal skeleton and a vertical skeleton constituting an internal frame, and process holes for CVI flow guide deposition are formed in the horizontal skeleton and the vertical skeleton; a plurality of hole groups for CVI flow guide deposition are formed on one side of the frame structure, the number of the plurality of hole groups is the same as the number of rows or columns of the internal frame, and the plurality of hole groups correspond to a plurality of frame rows or a plurality of frame columns of the frame structure one by one; the hole group comprises an air outlet hole and a flow guide hole; the flow guide hole is in sliding fit with a flow guide pipe used in CVI flow guide deposition; and a plugging structure is arranged at the air outlet hole and the flow guide hole. Through the labyrinth cavity structure, the determined skeleton part structure form, the assembly scheme and the interval distance between adjacent skeleton parts, the application has the advantages of simple assembly, only one CVI flow guide deposition process, shortened time cost, and facilitation of production promotion.

An active noise control system (100) and an aircraft (200) having such an active noise controls system (100) are disclosed. The active noise control system (100) comprises a foil (110), configured to be mounted to an interior panel of an aircraft (200). The foil (110) comprises a plurality of conductive tracks (111), a driver arrangement (112), a plurality of actuators (113), and a plurality of sensors (114). Each of the plurality of actuators (113) is connected to the driver arrangement (112) via at least one of the plurality of conductive tracks (111). Each of the plurality of sensors (114) is connected to the driver arrangement (112) via at least one of the plurality of conductive tracks (111). Each of the sensors (114) is configured to sense vibrations and / or acoustic sounds of the foil and to send measurement signals (115) to the driver arrangement (112) The driver arrangement (112) is configured to analyze the measurement signals (115) and to generate and send control signals (116) to the actuators (113) so that the actuators generate vibrations and / or acoustic sounds that counteract the vibrations and / or acoustic sounds sensed by the sensors (114).

The invention discloses an active flexible heat-proof material as well as a preparation method and application thereof. The active flexible heat-proof material comprises an active flexible heat-proof material heat-insulating layer, a metal reflecting layer and a fiber fabric layer which are sequentially stacked, the active flexible heat-proof material heat insulation layer is prepared by mixing methyl hydrogen liquid ceramic resin, methyltrimethoxysilane, a catalyst, a diluent, a graphite substrate, ceramic chopped fibers and silicon dioxide hollow microspheres and then calcining. The material is low in overall density and excellent in heat insulation performance, the material structure is stable in an extreme high-temperature environment, the heat insulation performance is not attenuated, and the safety of internal equipment is ensured. The method is suitable for a spacecraft surface sensitive to weight or a high-speed aircraft thermal protection system, the structural load can be reduced, and the energy efficiency can be improved.

The invention provides a long-heat-conduction-path dot-matrix type thermal protection structure filled with a heat insulation and phase change material. The long-heat-conduction-path dot-matrix type thermal protection structure comprises a structure outer plate, a dot matrix layer, a heat insulation layer, a phase change material layer and a structure bottom plate. The dot matrix layer is arranged between the structure outer plate and the structure bottom plate and is composed of supporting columns and supporting rods, the supporting columns connect the structure outer plate on the upper side and the structure bottom plate on the lower side to the supporting rods, a certain distance is arranged at the position of the connecting point, and the heat conduction path of the dot matrix structure is prolonged; gaps of the dot matrix layer are filled with the heat insulation layer and the phase-change material layer, the heat insulation layer reduces conduction of heat of the outer plate to the inside, the phase-change material layer absorbs heat conducted to the structure bottom plate at a specific temperature by means of the energy storage characteristic of the phase-change material layer, and the good heat insulation and temperature control effect is achieved. The lattice structure can significantly reduce the weight of the structure, the long heat conduction path design and heat insulation can significantly slow down the rapid temperature rise of the structure bottom plate caused by the thermal short circuit effect through filling of the phase change material, and the structure has excellent thermal protection performance.