75 results about "Nose cone" patented technology

A method of directing particles entrained within an airflow disposed to enter a gas turbine engine of an aircraft is provided. The gas turbine engine includes a nose cone, a fan section, a compressor inlet, a compressor section, and a turbine section. The nose cone is fixed for rotation with the fan section. The method includes: providing an entrained particle removal (EPR) system configured to inject a fluid outward from the nose cone from a plurality of nozzles engaged with the nose cone, wherein the nozzles are spaced apart from one another around a circumference of the nose cone; and controlling the EPR system to inject the fluid from the plurality of nozzles into the airflow. The particles wetted by the injected fluid are subject to centrifugal force in and aft of the fan section and are directed radially outward of the compressor inlet.

A payload container includes a nose cone electronics hub assembly, a collar assembly, a canister assembly, and a grid fin system. The nose cone electronics hub assembly has a nose cone housing electronic components in support of flight and detonator systems. The collar assembly includes a collar coupling the nose cone electronics hub and canister together. A canister of the canister assembly houses separate parts when in a shipping configuration that can be attached to an exterior of the payload container in a deployment configuration. A fin assembly facilitates vertical descent of the payload container when dropped from an unmanned aircraft.

The invention relates to a cross-medium aircraft variant nose cone with dynamically adjustable surface wettability, and belongs to the field of cross-medium aircrafts. Comprising a variant skeleton and a flexible skin. Under the heating condition, the variant framework can drive the flexible skin to deform, the flexible skin keeps the deformed state after normal temperature is recovered, and therefore curvature distribution and section characteristics of the nose cone structure are changed. The surface of the flexible skin is covered with a wettability intelligent response coating. After the wettability intelligent response coating is irradiated by ultraviolet rays, the surface wettability can be converted into a super-hydrophilic state, and the wettability intelligent response coating can be recovered into a super-hydrophobic state after being heated. The head configuration and the surface wettability layout of the cross-medium aircraft are adjusted according to the specific water entry working condition, the hydrodynamic force and the hydrodynamic moment borne by the aircraft in the water entry process are changed through the influence effect of the configuration and the surface wettability on water entry cavitation evolution, the impact load is reduced as needed, and the attitude and trajectory stability is improved; the purposes of being compatible with various launching platforms and guaranteeing the water entry efficiency and safety are achieved.

A lighter-than-air airship has an exoskeleton constructed of spokes and hubs to create a set of connected hexagrams comprised of isosceles triangles wherein the spokes flex and vary in length to produce the slope of said airship's surface. In one embodiment, the exoskeleton connects to a nose cone that includes a cockpit cabin for controlling the airship's operation from a single location that can be physically separated from the exoskeleton in response to catastrophic events and for autonomous and / or remotely piloted operation. An improved means is also provided for landing and unloading cargo, and through use of unmanned aerial vehicles in another embodiment, the airship is configured for remote pickup, transport, delivery and return of payloads such as packages. In yet another embodiment, the airship provides a communications platform for beam form transmission and satellite signal relay, including in combination with the foregoing disclosed attributes.

A multifunctional integrated Mars reentry detector relates to the technical field of Mars detection and mainly comprises a detector body, a rotor wing, a nose cone cover, a rotating device, a heat shield and the like. During launching, the rotor wings are in a compressed and folded state, and the detector body, the nose cone cover and the heat-proof cover serve as a whole device fairing. In the initial stage of reentry, the heat shield is used for protecting the rotor blade from pneumatic heating, can be further unfolded to a certain angle after descending at a certain speed, and serves as a speed brake for further speed reduction. And when the speed is reduced to subsonic speed, the rotor blades are unlocked and released, the blades are unfolded through an internal telescopic mechanism, the blades are driven to rotate by utilizing relative incoming flow, and rotation kinetic energy is converted into electric energy to be stored while lift force is provided for deceleration. In the landing stage, the rotor wings are switched to an active working mode, stored electric energy is used for providing rotating power for the rotating device, and autonomous hovering and obstacle avoidance are achieved; and the heat shield is completely unfolded to provide stable support and buffer for the detector body.

A treatment device includes a housing having a longitudinal axis extending between a proximal end and a distal end, a striking element disposed within the housing and moveable along the longitudinal axis, a tip disposed adjacent the distal end, and a nose cone disposed about at least a portion of the tip, the tip being moveable within the nose cone.

A snap fit nose cone comprising a body having a tapered nose and an edge with a snap fit lip configured such that the nose cone can be snapped into place on an inlet housing and such that the edge of the inlet housing abuts the edge of the inlet housing body.

This invention discloses an integrated heat dissipation and power generation device for high-speed aircraft, comprising: a heat dissipation pulsating heat pipe arranged in the nose cone of the high-speed aircraft; the heat dissipation pulsating heat pipe includes an annular pulsating heat pipe condensation section, several U-shaped pulsating heat pipe evaporation sections, and two intermediate sections connecting each U-shaped pulsating heat pipe evaporation section to the annular pulsating heat pipe condensation section; at least one branch pipe communicating with the intermediate section is arranged at intervals between the two intermediate sections; a heat extraction pipe for heat exchange with the annular pulsating heat pipe condensation section of the heat dissipation pulsating heat pipe; and a power generation module for converting the heat energy acquired by the heat extraction pipe into electrical energy. This integrated device provides thermal protection for the aircraft while converting high-temperature aerodynamic heat into electricity, alleviating the aircraft's demand for onboard battery power, reducing the need for carrying cooling fluid, and improving energy utilization efficiency. It is of great significance for aircraft operating in harsh thermal environments for extended periods.

The invention discloses a method for reducing storage and calculation amount in electromagnetic scattering non-singularity solution of a rotating body, and belongs to the field of electromagnetic engineering. According to the method, on the basis of a moment method (MoM) theoretical framework, an auxiliary function is introduced to eliminate the influence of Green function singularity in a surface integral equation on numerical calculation, and meanwhile, the physical constraint that the electric field divergence of a passive area is zero is further combined, so that effective order reduction of an unknown quantity is achieved; a traditional four-variable integral equation set is simplified into a three-variable coupled wave equation, the order and storage requirements of an impedance matrix are remarkably reduced, and the solving efficiency is greatly improved while the calculation precision is guaranteed. A specific electromagnetic rotating body device model is used for simulation, numerical verification shows that the method has excellent performance in calculation of an electromagnetic scattering field of a typical rotating body structure, and an efficient and reliable technical means is provided for rapid electromagnetic calculation of engineering equipment such as spherical communication satellites and missile nose cones.

The invention discloses a method for preparing a ceramic-based composite material through ultra-fast infiltration and pyrolysis and application, and belongs to the field of ceramic-based composite material preparation. The preparation method comprises the following steps: dipping a porous C / C composite material blank or a fiber preform in a ceramic precursor, carrying out cross-linking curing at a certain temperature, and then under a vacuum or inert atmosphere, adopting a Joule heating method to make the porous C / C composite material blank or the fiber preform subjected to ultra-fast heating (gt; and carrying out heat preservation at the temperature of 10 DEG C / s to finish a precursor-ceramic phase conversion process, and then repeating impregnation, cross-linking curing and precursor-ceramic phase conversion to realize the ultra-fast preparation of the ceramic-based composite material. The method can effectively solve the problems of long period, high energy consumption and the like of an impregnation cracking process. The method can be applied to ultra-fast preparation of high-temperature components such as hypersonic aircraft nose cones, wing leading edges and engine exhaust nozzles, and the mechanical property and the heat-proof performance of the high-temperature components can be improved.

The application relates to a preparation method of a high-density graphite-based dissipation heat-resistant composite material. In order to solve the problems of high cost of the existing dissipation heat-resistant composite material and poor ablation resistance of the graphite dissipation heat-resistant composite material, the preparation method comprises the following steps: mixing multi-element dissipation agent raw materials, uniformly mixing the multi-element dissipation agent raw materials, loading the multi-element dissipation agent raw materials into a graphite crucible, processing a high-density graphite base into a shape to be prepared, mounting the high-density graphite base on a lifting rod of a vacuum infiltration furnace, preparing a dissipation agent alloy melt, and immersing the graphite base in the dissipation agent alloy melt to obtain the high-density graphite-based dissipation heat-resistant composite material. The high-density graphite-based dissipation heat-resistant composite material is prepared by using the vacuum infiltration method, high-purity graphite is used as the base, the composition of the dissipation agent alloy is adjusted, the dissipation agent is wetted with the graphite base, and the ablation resistance of the workpiece is improved. The application can be used for preparing aircraft engine nozzle throat liners, aircraft nose cones and wing leading edges and the like components.

The invention discloses a heat insulation-phase change composite thermal protection device which comprises a heat insulation unit and a heat absorption and removal unit. The heat insulation unit comprises an anti-ablation layer, a buffer heat insulation layer and a main body heat insulation layer; the heat absorbing and discharging unit comprises a high-temperature phase change layer, a medium-temperature phase change layer and a sensible heat recovery layer, and the high-temperature phase change layer and the medium-temperature phase change layer communicate with the sensible heat recovery layer through exhaust pipelines; the anti-ablation layer, the sensible heat recovery layer, the buffer heat insulation layer, the high-temperature phase change layer, the main body heat insulation layer and the medium-temperature phase change layer are sequentially arranged from outside to inside; the heat insulation performance of the materials selected by the buffer heat insulation layer is superior to that of the main heat insulation layer; and after being gasified, the phase-change materials of the high-temperature phase-change layer and the medium-temperature phase-change layer enter the sensible heat recovery layer through the exhaust pipeline to absorb and take away part of heat penetrating through the anti-ablation layer. According to the composite thermal protection device, external heat flow can be effectively blocked, internal heat can be timely dissipated, and therefore efficient heat insulation and temperature control of the nose cone part of the aircraft are achieved.

A nose cone 100 for reducing drag at supersonic speeds, comprises a frontal right conical body 112 and a rear cylindrical body 114. The nose cone has a pointed front end 116 and a blunt rear end 118. The conical body 112 includes a plurality of air intake apertures 122; and a network of air inlet passages 134 leading into the air inlet passages, terminating in convergent nozzle sections. The cylindrical body 114 defines frontal and rear aerodynamic surfaces 120, 120.1, respectively, over which airflows. The cone 100 includes a plurality of air venting apertures 124; and a network of air outlet passages 126 in flow communication with the air venting apertures 124 and the air inlet passages 134. The air outlet passages 126 define divergent nozzle sections adjacent the air inlet passages. A throat 128 is defined between each air inlet passage 134 and an associated air outlet passage 126, at which airflow is momentarily choked. The air inlet and outlet passages and the throat together comprise a convergent-divergent nozzle.

The present application relates to the technical field of aircraft reentry characteristic model verification, in particular to a kind of ball nose cone target high Mach number reentry radiation characteristic simulation model verification method, comprising: the head shock detachment distance of the ball nose cone target is calculated;Using iterative method to solve the final value of incoming flow parameter, wherein, incoming flow parameter includes incoming flow density and incoming flow Mach number;The head point heat flow of the ball nose cone target is calculated;Error analysis is carried out to the head point heat flow of the ball nose cone target calculated, to modify the final value of incoming flow parameter;The final value of the modified incoming flow parameter is input into the preset ball nose cone high Mach reentry radiation characteristic simulation model, and the reentry radiation characteristic data of the ball nose cone target is solved;Ball nose cone target high Mach number reentry radiation characteristic simulation model is verified.The verification method provided by the present application can reduce the influence of incoming flow parameter error on simulation model verification, improve the accuracy of model verification.

A system for applying a braid to a tubular member includes an infeed clamp, an infeed guide, and a nose cone. The infeed clamp is configured to prevent rotation of a workpiece, a first steering member, and a second steering member. The infeed guide is configured to maintain the first steering member and the second steering member in a desired axial location along the workpiece. The nose cone includes a proximal plate, a distal plate, and a radially expanding inner guiding member. The proximal plate and the distal plate form a recess when joined together, and the radially expanding inner guiding member is located in the recess.

Disclosed here are systems for detachable airframe components including detachable nose cones, propeller assemblies and motors. In some example embodiments, the assemblies include a nose cone with a connection receiver, a motor assembly with a rotatable section, where the rotatable section includes torque arms configured to secure with the nose cone connection receiver, and a propeller assembly, configured to connect to the nose cone.

The invention discloses a preparation method of a high-performance ablation-resistant thermal insulation material phenolic resin glass fiber plate, and belongs to the technical field of composite material preparation. The method comprises the following steps: mixing phenolic resin and absolute ethyl alcohol in proportion to prepare a prepreg; impregnating the high-silica glass fiber cloth in the prepreg layer by layer, then overlapping, and placing the high-silica glass fiber cloth between the PET films coated with the release agent; and carrying out hot press molding at specific temperatures (170 DEG C, 180 DEG C and 190 DEG C) and pressure (10kN) for 2 hours by using a hot press, cooling, demolding, and finishing to obtain the phenolic resin glass fiber plate. The prepared plate has excellent ablation resistance and heat insulation performance and high mechanical strength, is suitable for high-temperature environments such as spacecraft heat insulation covers and missile nose cones, and is simple in process, low in cost and suitable for batch production.

The disclosure relates to a nose cone (160) for a vibrometry measurement system. The nose cone comprises a cone body (165) having a base portion (210) and a top portion (220); the top portion (220) comprising a plurality of ports (230, 332, 334) on a top surface (225), wherein at least one light guiding element (144) is coupled to at least one of the plurality of ports (230, 332, 334). A speaker duct (336) configured to transmit an acoustic stimulus from a speaker (152); and a microphone duct (338) configured to relay an acoustic pressure to a microphone (154). The invention further relates to a vibrometry measurement system and a method for manufacturing a nose cone.

The application belongs to the technical field of steel rolling, and discloses a kind of automatic coil collecting device and method for high-speed wire production, including conveying roller way, coil collecting drum, nose cone, supporting claw, core rod, supporting plate and detection module;Coil collecting drum is arranged at the end of conveying roller way along the conveying direction, and nose cone is arranged in it, and coil falls into coil collecting drum and is sleeved on the outer periphery of nose cone;Supporting claw is arranged on the outer periphery of nose cone and can be opened and closed, and when closed, it is used to support the coil on the nose cone, and when opened, it allows the coil to fall;Core rod is coaxially arranged below the nose cone;Supporting plate is arranged on the outer periphery of core rod and can be lifted along its axial direction, and is used to support the coil;Detection module is arranged beside conveying roller way and is communicatively connected with supporting claw, and is used to detect the head end and tail end of coil, and the detection module is configured to: when detecting that the head end of a group of coils reaches the detection position, control the supporting claw to open after delaying for a first time T1;When detecting that the tail end of a group of coils leaves the detection position, control the supporting claw to close after delaying for a second time T2.

A treatment device includes a housing having a longitudinal axis extending between a proximal end and a distal end, a striking element disposed within the housing and moveable along the longitudinal axis, a tip disposed adjacent the distal end, and a nose cone disposed about at least a portion of the tip, the tip being moveable within the nose cone.

The application relates to a variant nose cone based on a rigid-flexible coupling bionic scale skin, which comprises an internal skeleton unit (1) for providing support for the variant nose cone and driving deformation of the variant nose cone, which is a rigid-flexible coupling multilayer parallel folding and unfolding mechanism; a rigid skin unit (2) covered on the top and outer side of the rigid-flexible coupling multilayer parallel folding and unfolding skeleton unit (1), which is composed of a plurality of bionic scales (21); and a rigid-flexible coupling connecting unit (3) for connecting the internal skeleton unit (1) and the rigid skin unit (2), which transmits the deformation movement of the internal skeleton unit (1) to the rigid skin unit (2) and drives the relative movement of the plurality of bionic scales (21), so that the variant nose cone is switched between a two-dimensional flat state and a three-dimensional cone state. Compared with the prior art, the application has the advantages of realizing large-scale deformation, achieving excellent aerodynamic shape continuity, environmental adaptability and lightweight target and the like.

The sub-orbital airship comprises a single-stage rocket and an airship body, the airship body is installed at the top of the single-stage rocket and comprises a load cabin, a propelling cabin is arranged at one end of the load cabin, a consignment cavity is formed in one end of the load cabin, and a supporting cavity is formed in the other end of the load cabin. A consignment cavity is formed in the base, one end of the consignment cavity is connected with an airship nose cone, an air-drop platform is arranged in the consignment cavity, the air-drop platform comprises a first compression releasing machine, one end of the first compression releasing machine is in transmission connection with an air pressure lifting platform, and the airship nose cone comprises a landing cabin body. According to the method, safe recovery and reuse after launching can be achieved, intact recovery of airships and rockets is ensured, the launching cost is greatly reduced, the resource utilization rate is increased, and a solid technical foundation is provided for wide application of sub-orbital aircrafts in the fields of commercial spaceflight, scientific research, emergency response and the like.

The invention provides an integrated active cooling and propelling system of a hypersonic gliding aircraft, a nose cone assembly and an aircraft, and belongs to the technical field of thermal protection and propelling of hypersonic aircrafts. The problems of insufficient working medium heat sink, single energy utilization and incompatibility with propulsion in traditional active cooling are solved. The passive thermal protection layer is attached to the outer side surface of the aircraft nose cone and used for blocking external extreme aerodynamic heat flow. The active cooling flow channel structure is arranged between the passive thermal protection layer and the nose cone; the ammonia conveying module comprises a first ammonia storage tank, a throttling valve and a second check valve which are sequentially connected through a pipeline, and an outlet of the ammonia conveying module communicates with an inlet of the active cooling flow channel structure; an outlet of the active cooling flow channel structure is communicated with an inlet of the auxiliary propulsion spray pipe module; and the auxiliary propulsion spray pipe module is used for spraying out the working medium from the active cooling flow channel structure in an accelerated manner. The method is mainly used in the field of aerospace equipment.

An apparatus for stabilizing a fan assembly of a gas turbine engine during transport includes a ring configured to engage a nose cone of the fan assembly; an outer bracket connected to the ring and configured to mount to a fan casing of the fan assembly; and a tensioning assembly between the ring and the outer bracket to move the ring along an axis of a shaft of the fan assembly. An assembly with the gas turbine engine and a method are also disclosed.

The invention provides a vehicle nose cone transonic speed reverse design method based on a bending shock wave theory and a vehicle nose cone. The method comprises the steps that incoming flow parameters are determined and designed according to a flight envelope; determining a curved shock wave surface shape function and a discrete shock wave molded line according to head cone geometric constraints, and solving discrete point aerodynamic parameters and first-order derivatives based on a curved shock wave theory; solving a subsonic velocity flow field control equation by adopting a time domain characteristic line method; and obtaining external flow field information by simultaneous equations, and reversely generating a nose cone wall surface. According to the method, through the innovative logic of'presetting the shock wave form-reversely deriving the wall surface ', the bending shock wave theory and the time domain characteristic line method are fused, the transonic shock wave structure is accurately controlled, the wall surface pressure distribution is optimized, the flow separation and shock wave oscillation are inhibited, the design efficiency is improved while the design precision is ensured, and the method is suitable for large-scale popularization and application. Reliable technical support is provided for rapid optimization design of the nose cone of the vehicle, and the method is suitable for aerodynamic design of nose cones of various transonic vehicles.

This invention provides a composite material truss-type unmanned aerial vehicle (UAV) fuselage structure and the UAV itself, employing an inverted triangular layout. It includes a main tube, support tubes, segmented multi-port joints, wing-body joints, a nose cone, and a tail cone. The main tube consists of three carbon steel tubes with the same outer diameter and varying wall thickness, divided into multiple sections based on process feasibility. The support tubes include four types: transverse support tubes, upward-sloping support tubes, two-sided inclined support tubes, and two-sided reinforced support tubes, all with the same outer diameter. The segmented multi-port joints include four types: four-way joints, five-way joints, six-way joints, and seven-way joints, each divided into two segments and press-bonded to the main tubes and support tubes. A multi-port joint below the nose cone connects to the fuselage, while upper lugs connect to the wings and tail. The nose cone and tail cone are contoured. This invention improves the overall structural rigidity and instability resistance by coordinating the load distribution across the truss components, and allows for flexible installation of onboard equipment.

This invention discloses a sweating cooling structure with an inclined cylindrical cavity array on the leading edge sidewall, comprising a gradient-aperture porous medium substrate, an inclined cylindrical cavity array, and a coolant supply chamber. The gradient-aperture porous medium substrate is divided into at least three pore size gradient regions along the normal direction from the back-heating surface to the heated surface, with the pore size decreasing progressively along the coolant flow direction. Simultaneously, a pore size gradient matching the heat flow distribution is also set along the spanwise direction. The inclined cylindrical cavity array is located on the surface of the heated surface, with its axis forming an angle of 15°–45° with the incoming flow direction, utilizing the negative pressure effect of the incoming flow to enhance steam discharge. This invention completely solves the steam blockage problem through a dual gradient pore size design in both the normal and spanwise directions, achieves efficient gas-liquid separation through the negative pressure suction of the inclined cylindrical cavities, and enhances the stability of the cooling film through the cavity turbulence effect. It is suitable for active thermal protection in high-heat-fluidity areas such as the nose cone sidewall, rudder leading edge sidewall, and wing leading edge sidewall of hypersonic vehicles traveling at Mach 3–11.5.

A heat pump engine comprising a main body shell, a nose cone fairing 2, first front generator 3, evaporating pipe 6, liquid flow core pipe 7, gas compressor rotor 8, gas compressor stator 9, radiator

The invention relates to a near space aircraft adjustable head cone and an adjusting method thereof, and belongs to the technical field of mechanism science and the technical field of aerospace. The adjustable nose cone comprises a first-stage actuator, a second-stage actuator, a third-stage actuator, an aircraft nose cone shell and an aircraft nose cone skin. The aircraft nose cone shell is covered with the aircraft nose cone skin, and the aircraft nose cone shell is connected with the first-stage actuator, the second-stage actuator, the third-stage actuator and the aircraft nose cone skin in a bearing mode and provides structural support for the aircraft nose cone skin. The first-stage actuator, the second-stage actuator and the third-stage actuator are connected in series, can stretch out and draw back along the central axis of the aircraft nose cone and deflect by a preset angle relative to the central axis of the aircraft nose cone, so that the angle and the stretching length of the aircraft nose cone are adjustable.

The application provides a hypersonic vehicle head drag reduction and heat reduction adaptive device, which combines a drag reduction and heat reduction device of an aerodynamic rod and a sweat cooling device; when the vehicle is in supersonic flight, the aerodynamic rod pushes the shock wave away from the wall surface of the nose cone front edge to achieve the effect of drag reduction, reduces the pressure and heat flow at the nose cone front edge, and makes the sweat cooling medium more easily flow out of the ceramic-based porous material, thereby solving the problem of low stagnation point cooling efficiency in sweat cooling; the cooling medium water is in liquid state in the connecting pipeline, the coolant micropore storage tank and the internal cavity of the nose cone shell, enters the ceramic-based porous material, absorbs heat in the ceramic-based porous material, is warmed and phase-changed, flows out of the porous material, then continues to flow downstream, and forms an air film on the surface of the nose cone shell area, thereby achieving the effect of heat reduction; and with the change of the incoming flow pressure, the aerodynamic rod can be self-adaptively adjusted in the length of extension to adjust the effect of drag reduction, and the sweat cooling can also be self-adaptively adjusted in the flow of the coolant to adjust the effect of heat reduction.