323results about "Fuselage insulation" patented technology

The invention discloses a heat-proof skirt which is of a single-layer or multi-layer structure, each single-layer material is formed by compositing quartz fiber cloth and coating when multiple layers are arranged, the heat-proof skirt is made of 6-8 single-layer materials, and the skirt edge is in a wave shape. The coating is prepared from the following raw materials in parts by weight: 4 to 6 parts of vinyl polysiloxane, 15 to 20 parts of hydrogen-containing silicone oil, 15 to 20 parts of ammonium polyphosphate, 10 to 15 parts of aluminum hydroxide, 5 to 8 parts of silicon dioxide, 0.005 to 0.02 part of a platinum catalyst and 30 to 40 parts of a solvent. The heat-proof skirt has contractility, and the influence on normal swinging resistance of the engine is reduced to the maximum extent.

The invention discloses a composite heat insulation coating structure and method for a high-temperature component of an unmanned aerial vehicle. The structure comprises a heat insulation base material layer made of a high silica glass fiber material and a reflecting layer made of metal foil, and the heat insulation base material layer and the reflecting layer are compounded into a whole. The reflecting layer is used for reflecting heat radiation, and the heat insulation base material layer is used for blocking heat conduction. The problems that a traditional ceramic fiber material is prone to embrittlement and short in service life in a high-temperature vibration environment are effectively solved through a double heat insulation mechanism of reflection-blocking. The coating structure has excellent heat-resistant cyclicity, good flexibility and process adaptability, can perfectly fit special-shaped structures such as a bulkhead and a tail cover of the unmanned aerial vehicle, and remarkably improves the reliability and prolongs the service life of a thermal management system of the unmanned aerial vehicle.

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.

There is provided an acoustic black hole comprising: in a first axis, a variation from a first characteristic to a second characteristic, along a line; in a second axis different to the first axis, a variation from a first characteristic to a second characteristic, along the second axis.

An acoustic panel for an aircraft includes a skin, a plurality of first cells and a plurality of second cells interspersed with the first cells. Each first cell includes a first tube, a first endwall and a first cavity. The first tube projects vertically out from the skin to a distal end of the first tube. The first endwall closes off the first cavity at its distal end. The first cavity extends vertically through the first tube from the skin to the first endwall. Each second cell includes a second tube, a second endwall and a second cavity. The second tube projects vertically out from the skin to a distal end of the second tube. The second endwall closes off the second cavity at its distal end and is offset from the first endwall. The second cavity extends vertically through the second tube from the skin to the second endwall.

The invention discloses a reusable light heat-proof and heat-insulating material and application thereof, and the heat-proof and heat-insulating material is designed according to the anti-radiation / heat-proof / heat-insulating gradient thought for a thermal environment profile of a reusable aircraft. The surface of the material is a ceramic-based heat-proof coating with high emissivity, so that radiation heat flux permeation on the surface of the material is reduced; the interior of the material is a silicon-based heat-proof material, so that the material has a non-ablation / low-ablation effect; the innermost layer of the material is a light thermal insulation material to reduce the structural temperature of the aircraft. According to the thermal protection material spraying integrated forming method, the binding force between the material and the surface of the metal matrix is high, the interface is stable, the process implementation performance is good, the material has the capacity of being repeatedly used for a certain number of times, and the thermal protection and insulation requirements of repeatedly used aircrafts are met.

A thermal protection structure of a high-temperature and supersonic aircraft belongs to the technical field of thermal protection of aircrafts under high heat flux and comprises a water absorption layer, a front edge area and two side areas, and the front edge area comprises a first evaporative cooling porous medium layer, a cross-scale porous medium layer and a high-temperature-resistant wall surface which are sequentially laid on the outer side of the front edge of the water absorption layer; the two side areas comprise a second evaporative cooling porous medium layer and a third evaporative cooling porous medium layer which are laid on the outer sides of the two sides of the water absorption layer respectively, and a coolant conveying pipeline communicated with the front edge and the two side areas is arranged on the inner side of the water absorption layer; fluid is sucked to the high-temperature-resistant wall face through the capillary acting force of micron-scale and nano-scale pores of the first evaporative cooling porous medium layer and the cross-scale porous medium layer, the millimeter-scale pores of the cross-scale porous medium layer provide steam escape channels, and the high-temperature-resistant wall face blocks direct disturbance of high-temperature main flow to the fluid in the porous medium. Steam outlets on two sides of a front edge stagnation point discharge steam; the device has the advantages of no need of an additional liquid supply power device, high temperature resistance and high cooling efficiency.

A variable fuselage fuel tank structure of a high-speed aircraft comprises a fuselage fixing structure, an external flexible heat insulation skin and an internal skin bearing framework. The machine body fixing structure comprises a lower wall plate and an upper wall plate; the edges of the lower wall plate and the upper wall plate are connected, and the upper wall plate is provided with a fuel circulating hole; the outer flexible heat insulation skin is connected with the edge of the upper wall plate; the external flexible heat-insulating skin comprises a flexible fabric heat-proof outer layer and a microporous aerogel heat-insulating inner layer; the inner skin bearing framework is connected to the inner side of the outer flexible heat insulation skin and comprises reinforcing longitudinal beams, elastic elements and high-temperature-resistant rubber. The longitudinal beams are connected to the inner side of the outer flexible heat insulation skin in the longitudinal direction. A plurality of groups of elastic elements are respectively connected between two adjacent longitudinal beams, are distributed along the longitudinal direction, and are connected to the inner side of the external flexible heat insulation skin; and the high-temperature-resistant rubber is filled between the longitudinal beam and the elastic element, is filled in the elastic element and is adhered to the inner side of the external flexible heat-insulating skin.

The invention discloses a thin-wall sweating cooling structure and an application method thereof. The sweating cooling structure is composed of a plurality of sweating units arranged in an array mode. Each sweating unit sequentially comprises a flow distribution layer, a pressure control layer and a porous sweating surface from bottom to top. The flow distribution layer internally comprises a sweating working medium inlet and a main flow channel for distributing flow; the sweating working medium inlet and the main runner are shared by the sweating units which are arranged in an array manner; the pressure control layer internally comprises a branch flow channel for distributing flow, a porous medium pressure control structure and a buffer area, the branch flow channel is connected with the main flow channel, and the porous medium pressure control structure is arranged parallel to the sweating wall surface; the porous sweating surface is composed of a plurality of sweating micropores and a non-permeable solid wall surface, and the number, the size and the angle of the sweating micropores can be adjusted according to flow regulation and control requirements. The structure has the remarkable characteristics that the structure thickness is small, the sweating flow distribution can be adjusted at will, the porous array structure on the sweating surface can be adjusted at will, and the use efficiency of the sweating working medium is high.

A composite article comprises a body comprising a radial region, a tapered region, and a cavity region. The cavity region comprises a cylindrical shape and an inner diameter of the cylindrical shape decreases between the radial region and the tapered region. A nosecone region is adjacent to the tapered region of the body and at an opposing end of the body to the cavity region. A thermal protection structure is on the body and comprises a base layer, an insulating layer is on the base layer, and an erosion resistant layer is on the insulating layer. The insulating layer exhibits a specific gravity of from about 0.01 to about 1.0. Methods of forming the thermal protection structure and of thermally protecting an article are also disclosed.

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.

Thermal barrier curtain systems include a thermal barrier curtain and port and starboard anchor / hoist strap assemblies. The thermal barrier curtain includes forward and aft flexible air-impervious panels which are sealed about peripheral edges thereof so as to define a central thermal barrier region and separator elements positioned between and attached to the forward and aft panels within the central thermal barrier region so as to define a dead air space therebetween. Each of the port and starboard anchor / hoist strap assemblies includes an elongate strap, upper and lower attachment members positioned at opposed upper and lower ends of the elongate strap, an upper attachment loop adapted to receive and be attached to a respective connection element of the thermal barrier curtain, and an intermediate retaining loop which is adapted to receive and be attached to the lower attachment member of the anchor / hoist assembly.

A heat shield for protecting a part of an aircraft is disclosed including superposed layers providing thermal protection, and one or more layers providing fire protection. The heat shield also may include other layers implementing other functions.

The present invention proposes a modular ceramic-based composite thermal insulation tile and a design method thereof. The thermal insulation tile includes a ceramic-based outer shell, which is filled with a plurality of hollow topologically configured particles. The plurality of hollow topologically configured particles form a lattice structure particle layer and are integrally formed with the ceramic-based outer shell; the gaps between adjacent hollow topologically configured particles and the cavities inside the hollow topologically configured particles are filled with lightweight, low-thermal conductivity materials and together constitute a filling layer. The method uses finite element simulation to calculate the temperature difference between the upper and lower surfaces of the modular ceramic-based composite thermal insulation tile model, and uses it as a basis for optimizing the thickness of the filling layer. The present invention is applicable to ultra-high temperature thermal protection structures, and the thermal insulation performance is significantly improved while taking into account the goal of lightweight thermal protection structures; at the same time, the finite element method is used for structural parametric design to reduce the time period from design to application of the thermal protection structure, thereby achieving rapid optimization of the ultra-high temperature thermal protection structure.

The invention provides a honeycomb sandwich wall plate structure and a preparation method thereof, the honeycomb sandwich wall plate structure comprises a first panel, a second panel, a honeycomb core layer, a honeycomb insert block, a heat insulation plate and a fastener, and the first panel and the second panel are oppositely arranged; the honeycomb core layer is arranged between the first panel and the second panel, one side of the honeycomb core layer is connected to the first panel, the other side of the honeycomb core layer is connected to the second panel, and embedded holes are formed in the honeycomb core layer; the honeycomb embedded blocks are arranged in the embedded holes and are connected to the honeycomb core layer; the heat insulation plate is arranged on the side, away from the second panel, of the first panel. And the fastener is connected with the heat insulation plate and the honeycomb insert block. The wallboard structure provided by the invention has the characteristics of high temperature resistance, vibration absorption, high durability, low cost and light weight.

The invention discloses a testing device of an active and passive combined thermal protection system, and relates to the technical field of thermal protection of hypersonic aircrafts. The micro-channel cold plate comprises a micro-channel cold plate assembly, the top of the micro-channel cold plate assembly is fixedly connected with a passive thermal insulation layer, and a plurality of thermocouples are installed between the passive thermal insulation layer and the micro-channel cold plate assembly; a gear pump is installed at an inlet of the micro-channel cold plate assembly through a fuel oil pipeline, a fuel oil tank is installed at the inlet end of the gear pump, and an adjusting valve is installed at the outlet end of the gear pump. By adjusting the flow of the micro-channel cold plate in real time, the change of an external pneumatic thermal environment is well matched, so that efficient thermal protection of a large-area fuselage of a high-speed aircraft under long-time and high-heat-flow conditions is realized, and the consumption of a coolant is effectively reduced; and when the control module adjusts the flow, constant-pressure supply of the cooling working medium is achieved by changing the opening degree of the pressure adjusting valve, the working stability is improved, and the thermal failure problem of the active and passive thermal protection system is effectively avoided.

An apparatus is provided for an aircraft. This apparatus includes an acoustic panel, and the acoustic panel includes a face skin, a back skin and a cellular core disposed between and connected to the face skin and the back skin. The back skin laterally overlaps the face skin and includes an aerogel layer. The cellular core includes a plurality of cavities that extend through the cellular core from the face skin to the back skin. Each of the cavities is fluidly coupled with one or more perforations in the face skin.

An actively-cooled heat shield system includes a heat shield, a tank, a pump, a heat exchanger, and a turbine. The heat shield defines a windward side of a vehicle. The tank stores a coolant. The pump receives the coolant from the tank and outputs a pressurized coolant. The heat exchanger is integrally connected with the heat shield. The heat exchanger receives the pressurized coolant from the pump, transfers heat from the heat shield to the pressurized coolant to generate a heated fluid, and outputs the heated fluid. The turbine includes an inlet, a shaft, and an outlet. The inlet receives the heated fluid output from the heat exchanger. The shaft is coupled to the pump and includes turbine blades. The shaft rotates and powers the pump when the heated fluid received from the heat exchanger acts on the turbine blades. The outlet outputs the heated fluid.

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 relates to a sound-absorbing structure comprising a honeycomb structure (52) which includes at least one cell (60, 60') delimited by at least one wall (58') and at least one partitioning system (62). The latter comprises at least one subassembly (64, 64') positioned within the cell (60, 60') and designed to divide it into at least two cavities, as well as at least one tab (78, 78') connected to the subassembly (64, 64') and having a hook shape (84) positioned astride one of the first and second end edges (58.1', 58.2') of the wall (58'). This tab (78, 78') facilitates the placement of each partitioning system (62) and its retention within the cell (60, 60'). The invention also relates to an aircraft comprising at least one such structure and a method for assembling such a structure.

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.

A shielded, multi-layer heatshield material for hypersonic flight applications includes an ablative / insulative structure having a thickness and providing both thermal insulation and a first ablation resistance, and a separate shield layer bound to an outer surface of the ablative / insulative structure, the shield layer being thinner than the ablative / insulative structure while providing a second higher ablation resistance greater than the first ablation resistance. In one arrangement the ablative / insulative structure includes a carbon / carbon composite layer attached to a syntactic carbon foam layer, using carbon fiber stitched loops for reinforcement, and the shield layer is a carbide outer layer. In another arrangement the ablative / insulative structure includes a silica composite layer attached to a silica insulative foam layer, using stitched loops for reinforcement, and the shield layer is a ceramic outer layer.

The invention belongs to the field of aircraft noise reduction, and discloses an aircraft cabin active noise reduction control method and related equipment, and the method comprises the steps: collecting the vibration, rotating speed, cabin sound pressure and other multi-modal data of an aircraft in real time, and carrying out the flight state classification through a pre-trained multilayer neural network model; and the current accurate working condition is determined by combining state machine conversion logic, and finally, the parameters of the feed-forward filter are dynamically adjusted through a minimum mean square algorithm to realize noise reduction. According to the method, the high-frequency component suppression capability in a multi-source coupled noise scene is effectively improved, the noise reduction efficiency is kept stable in the whole flight period, the convergence speed is improved compared with a traditional scheme, and the local optimum problem is avoided through a state self-adaptive feed-forward mechanism.

The utility model provides a pressurizing calandria sandwich skin structure and an aircraft, the pressurizing calandria sandwich skin structure comprises an upper panel, a lower panel and at least one layer of pressurizing calandria limited between the upper panel and the lower panel, each layer of pressurizing calandria is formed by tightly arranging a plurality of pipe fittings with circular sections, and the pipe fittings are filled with gas with set pressure. According to the pressurizing calandria sandwich skin structure, loads are transmitted through gas, the whole structure is evenly stressed, the interior of the sandwich skin is basically hollow, the light weight requirement can be fully met, the pressurizing calandria sandwich skin is adopted to replace a traditional solid structure, the structural weight is reduced, and meanwhile the structural rigidity requirement is met.

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 invention discloses an aerogel heat insulation pad with a heat conduction layer and a preparation method of the aerogel heat insulation pad. The heat insulation pad is formed by clamping a heat conduction layer between two aerogel layers, and the heat conduction layer can disperse heat to the whole plane from a local concentration point and dissipate the heat to the air through the aerogel layers. By optimizing the material and structural design of the heat conduction layer, the heat dissipation and soaking capacity of the aerogel layer in the cross section direction is remarkably improved, and meanwhile the heat insulation performance of aerogel is reserved. The heat insulation pad has excellent heat management performance and is suitable for the fields of electronic equipment heat dissipation, aerospace heat management, building heat preservation and the like.

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.