426results about "Fuselage insulation" patented technology

The invention discloses a micro-channel device for phase change sweating and an application method, and belongs to the technical field of thermal protection of aircrafts. The structure sequentially comprises a solid inner wall, a transportation phase change layer and a sweating outer wall surface from inside to outside, and is mainly composed of a bottom plate, a porous rib wall, a solid rib wall, a solid outer wall, an inlet flow channel, a main flow channel, a branch flow channel and the like. The bottom plate is a solid inner wall and is used for supporting and conveying; the flow channel can be divided into an inlet flow channel, a main flow channel and a branch flow channel, and a cooling medium sequentially passes through the inlet flow channel, the main flow channel and the branch flow channel; a cooling medium is injected into an external flow field through the porous rib walls which are periodically arranged; the solid rib walls isolate the adjacent main runners, so that a cooling medium flows out through the porous rib walls; the outer surfaces of the porous rib walls of the solid outer wall are flush to form a sweating outer wall surface; the device has the advantages of being integrated in structure, small in thickness, uniform in sweating and efficient, and a new way can be provided for long-time cross-domain sailing thermal protection of high-speed aircrafts.

The thermal protection structure comprises a first porous medium layer arranged on the arc-shaped wall face of the front edge of the aircraft and a plurality of second porous medium layers arranged on the upper wall face and the lower wall face of the front edge of the aircraft at intervals in the flow direction. Solid walls are arranged between the first porous medium layers and the second porous medium layers and between every two adjacent second porous medium layers; the porosity of each second porous medium layer is firstly increased and then reduced along the flow direction, and the porosity of the first porous medium layer is smaller than that of the second porous medium layer at the most upstream. The method is applied to the field of thermal protection of aircrafts, the problem of local overheating existing in a traditional sweating cooling technology is effectively solved through the synergistic effect of the intermittent sweating layout and the gradient porosity material, and the situation that the cooling efficiency is reduced due to overheating of a cooling agent is avoided.

The invention discloses a medium-high temperature heat-power conversion working medium for a high-speed aircraft as well as a preparation method and application of the medium-high temperature heat-power conversion working medium. The medium-high temperature heat-power conversion working medium comprises the following components in percentage by mass: 10-90% of ammonia, 10-70% of multi-component alcohol and the balance of water, the multi-component alcohol comprises at least two of ethanol, isopropanol, ethylene glycol, n-propyl alcohol and n-butyl alcohol; and the mass of each alcohol in the multi-component alcohol is 5-40% of the total mass of the medium-high temperature heat-work conversion working medium. Ammonia, multi-component alcohol and water are combined for the first time, a double-bed catalyst is combined, physical heat absorption can be utilized, heat absorption can be enhanced through chemical reaction, the heat sink can reach 7.0 MJ / kg or above, and the cooling agent can be used for pneumatic heat absorption of a high-speed aircraft with the speed higher than 5 Mach and is high in cooling capacity and low in cost. The method is suitable for a thermal protection system of a pneumatic heating component with higher heat flux and continuous long time; meanwhile, high-temperature and high-pressure small-molecule gas is generated, thrust can be generated to do work, and the maneuverability of the high-speed aircraft is effectively enhanced.

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.

The invention discloses a fiber-reinforced resin-based composite board, a composite board component and a preparation method.The fiber-reinforced resin-based composite board comprises a first preformed body and a second preformed body, and the first preformed body and the second preformed body are matched to form a hollow structure with a cavity; the first preformed body and the second preformed body are formed by laying fiber prepregs layer by layer in different directions, so that fibers in the first preformed body and the second preformed body are distributed in different directions. According to the fiber-reinforced resin-based composite board prepared by adopting a two-step method, the porosity in the composite board can be greatly reduced, the fibers are in a straightened arrangement state, and the bearing performance or mechanical performance of the fiber-reinforced resin-based composite board is superior to that of a laminated board, a plate rib structure composite board and a pultrusion hollow structure composite board under the same thickness. The problem that in the prior art, bearing capacity, light weight and sound insulation performance of a plate, especially a vehicle body plate are difficult to consider at the same time is solved.

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.

The utility model discloses a self-adaptive variable acting force heat shield. The heat shield is mainly formed by an inner flexible metal dense net, an outer flexible metal dense net, a flexible heat insulation material and a supporting framework through a forming mold. The inner flexible metal dense net and the outer flexible metal dense net wrap and clamp the flexible heat insulation material to form the flexible main body of the heat insulation cover, so that the heat insulation cover is good in heat insulation performance, strong in damp heat resistance, mould resistance and salt mist resistance, moderate in toughness and rigidity, and capable of generating corresponding displacement and certain elastic deformation according to borne tension, compression and bending acting force; and the device can be suitable for strong wind blowing in any direction and can be used for a long time in a high-temperature environment.

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 rotary wing aircraft with a fuselage that forms an aircraft interior region, the fuselage comprising an upper primary skin that separates the aircraft interior region from an aircraft upper deck arranged above the fuselage, the upper primary skin being mounted on top of a predetermined number of supporting beams; wherein the aircraft upper deck comprises an engine accommodating region with a firewall arrangement, the engine accommodating region accommodating at least one aircraft engine within the firewall arrangement such that the firewall arrangement defines a fire proof separation at least between the at least one aircraft engine and the aircraft interior region; and wherein the firewall arrangement comprises at least a funnel-shaped lower firewall arranged between the at least one aircraft engine and the upper primary skin of the fuselage.

The invention provides an air rudder, and belongs to the technical field of air rudders. The control surface skin is arranged outside the rudder skeleton and surrounds the rudder skeleton, the control surface skin is connected with the rudder skeleton, and a mounting cavity is formed between the control surface skin and the rudder skeleton; the rudder shaft is connected with the rudder framework and located at the bottom of the rudder framework; the rudder skeleton, the rudder surface skin and the rudder shaft are integrally formed through three-dimensional printing. The air rudder has the beneficial effects that the rudder framework, the rudder surface skin and the rudder shaft are integrally formed through three-dimensional printing, connection of the air rudder and an aircraft can be achieved, no screw or rivet or other connecting structures exist among the rudder framework, the rudder surface skin and the rudder shaft, the rigidity of the air rudder can be enhanced, the assembling clearance of the air rudder is reduced, the machining and assembling procedures are reduced, and the cost is reduced.

The present invention discloses a multi-layer structure high temperature resistant heat insulating composite material and a preparation method thereof, belonging to the technical field of thermal protection materials. The multi-layer structure high temperature resistant heat insulating composite material includes, from top to bottom, a high-density ceramic composite material layer, a fiber felt reinforced ceramic composite material layer, an aerogel heat insulating composite material layer and a resin-based composite material layer. The upper panel of the multi-layer structure high temperature resistant heat insulating composite material has excellent high temperature resistance and ablation resistance, and the lower panel has excellent bonding performance. Each layer structure can give full play to its own performance advantages, and has broad application prospects in the fields of aerospace and the like. In addition, its preparation process is simple, the cycle is short, and the cost is low, which is conducive to mass production.

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 invention belongs to the technical field of acoustics, and discloses a low-frequency and broadband noise reduction acoustic structure and a design method of the acoustics structure.The acoustics structure comprises a plurality of acoustics structural parts and embedded parts, at least one embedded part is arranged in the multiple acoustics structural parts in a penetrating mode, and the multiple embedded parts are arranged on the same embedded part. The acoustic structural parts communicated with the openings in the two ends of the embedded part are arranged in pairs and form a first element together with the embedded part, the acoustic structural part not communicated with the embedded part is a second element, and the first element and the second element are arranged in parallel. Compared with the prior art, when sound waves pass through the acoustic structural members which are arranged in pairs, coupling resonance generated in the acoustic structural members can widen the silencing frequency band at a certain resonant frequency in a low-frequency range, the additionally arranged embedded members can enable the acoustic structural members to generate a new silencing frequency band, and the weak coupling effect existing between the parallel elements can continue to widen the silencing frequency band. Finally, the low-frequency and wider-band noise elimination and reduction effect is achieved.

The invention provides a forming method and a forming device for a high-temperature-resistant hollow leading edge structure. According to the forming method, a V-shaped structure prefabricated body, an H-shaped structure prefabricated body, a flat plate structure prefabricated body, a corresponding V-shaped structure forming mold, a corresponding H-shaped structure forming mold and a corresponding flat plate structure forming mold are assembled to form a plurality of combined molds; the method comprises the following steps: firstly, preparing a plurality of combined molds, then performing densification treatment on the plurality of combined molds by adopting a precursor impregnation cracking process, then removing the molds, processing, preparing a coating on the surface of a component to obtain a V-shaped structure, an H-shaped structure and a flat plate structure, and finally assembling the V-shaped structure, the H-shaped structure and the flat plate structure to form the hollow front edge structural component. According to the method, preparation of the hollow leading edge structural member made of the carbon fiber reinforced ceramic matrix composite can be achieved, mold removal is convenient, and the mold can be repeatedly used.

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.

The invention relates to a honeycomb structure for an acoustic panel comprising: - at least one acoustic cell comprising a cavity delimited by a peripheral wall, and - at least one first partition (24) arranged in the cavity to divide the cavity in two, the first partition (24) comprising a first acoustically permeable membrane (25), characterized in that the first partition (24) further comprises first lugs (26), each first lug (26) being received in a respective first slot formed in the peripheral wall to mechanically fix the first partition (24) in the corresponding cavity. Abbreviated figure: Figure 5