111 results about "Nose cone" patented technology

The invention discloses an anchoring device and an aerostat anchoring and releasing method. The anchoring device comprises an anchoring tower body, a nose cone, a guide assembly and a winding assembly. The nose cone is arranged at the top of the anchoring tower body through the mounting frame, is located on the mounting frame and can be opened and closed relative to the mounting frame so as to be switched between a release state and a locking state; the guide assembly comprises a rail and a sliding block arranged on the rail in a sliding mode, and the rail inclines downwards from the end, provided with the nose cone, of the installation frame and extends to the bottom of the anchoring tower body to form a triangular structure with the anchoring tower body. The winding assembly comprises a pull rope and a locking block fixed to the pull rope, and the pull rope is connected to the two ends of the sliding block so as to drive the sliding block to reciprocate along the track. A driving channel is formed in the center of the nose cone, the pull rope penetrates through the driving channel, and the locking block can move along the track along with movement of the pull rope and gradually enter the driving channel to push the nose cone to enable the nose cone to be converted into a locking state from a releasing state so as to lock the aerostat.

The present invention discloses a porous material transpiration cooling leading edge structure, which includes multiple layers of porous material layers and a cooling cavity provided on one side of the multiple layers of porous material layers. A coolant supply channel is provided in the cooling cavity, and the coolant supply channel is used to supply coolant to the cooling cavity. The multiple layers of porous material layers include an outer layer, an intermediate layer, and an inner layer, and the material of the inner layer is a shape memory alloy porous material. The present invention can adaptively adjust the porosity of the porous material, automatically adjust the usage amount of the coolant, effectively improve the usage efficiency of the coolant, and further enhance the thermal protection effect in the high-temperature tip region. Moreover, the shape memory alloy material can be reused. The porous material transpiration cooling leading edge structure can be used for structures such as the nose cone of an aircraft, the leading edge of a rudder, and the leading edge of a wing that are subjected to more high-temperature heat fluxes.

The invention belongs to the field of embedded atmosphere data system fault identification and elimination calculation, and particularly relates to a flying wing aircraft seven-pressure-measuring-point embedded atmosphere data system fault identification and calculation method, which comprises the following steps: step 1, pressure measuring points are grouped: each group of pressure measuring points comprises nose cone lower surface pressure measuring points, left and right symmetrical pressure measuring points and upper and lower symmetrical pressure measuring points, the pressure measuring points are used for solving the attack angle alpha and the sideslip angle beta, one pressure measuring point is reused, the pressure measuring points are divided into four groups under different attack angles, and each group of pressure measuring points is composed of four pressure measuring points; step 2, identifying the fault type of the pressure measuring point and the fault point: according to the attack angles alpha 1, alpha 2, alpha 3 and alpha 4, the sideslip angles beta 1, beta 2, beta 3 and beta 4 obtained by resolving the four groups of pressure measuring points under different attack angles, and the pressure measuring point used for resolving the attack angle alpha and the sideslip angle beta, identifying the fault type of the pressure measuring point, and further identifying the fault point; and step 3, removing fault points, and calculating and outputting atmospheric parameters including an attack angle alpha and a sideslip angle beta.

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.

An assembly for an aircraft propulsion system includes a propulsor and an imaging assembly. The propulsor includes a propulsor disk, a plurality of propulsor blades, and a nose cone. The plurality of propulsor blades are circumferentially distributed about the propulsor disk. Each propulsor blade of the plurality of propulsor blades extends radially between and to a root end and a tip end. The root end is disposed at the propulsor disk. The propulsor disk and the plurality of propulsor blades are configured to rotate about the rotational axis. The nose cone is disposed axially adjacent the propulsor disk. The imaging assembly includes an imaging device disposed on the nose cone. The imaging device includes a camera. The camera is configured to capture image data of each propulsor blade of the plurality of propulsor blades as the plurality of propulsor blades rotate about the rotational axis.

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.

The invention discloses a tungsten-containing ultrahigh-temperature high-entropy rare earth boride ceramic material as well as a preparation method and application thereof. The preparation method of the tungsten-containing ultrahigh-temperature high-entropy rare earth boride ceramic material comprises the following steps: 1) preparing a green body from metal oxide powder and boron carbide powder; 2) performing electric field sintering, crushing and screening on the green body to prepare tungsten-containing ultrahigh-temperature high-entropy rare earth boride ceramic powder; and 3) pressing the tungsten-containing ultrahigh-temperature high-entropy rare earth boride ceramic powder into a green body, and then carrying out spark plasma sintering. The tungsten-containing ultra-high-temperature high-entropy rare earth boride ceramic material has excellent ultra-high-temperature oxidation resistance, is wide in component space range and low in cost, has a very wide application prospect in ultra-high-temperature thermal protection components such as the leading edge, the nose cone and the control surface of a hypersonic aircraft, and has a wide application prospect. And the preparation method has the advantages of simple operation and process flow, short reaction time, low synthesis cost and the like, and is suitable for large-scale industrial production.

The utility model relates to the technical field of steel rolling equipment, in particular to a wire coil collecting and collecting device capable of reducing head scratches of wire coils, which comprises a coil collecting drum, a nose cone, a temporary coil supporting arm, a lifting coil supporting plate and a core rod, and is characterized in that a lifting support for lifting the wire coils is arranged on the lifting coil supporting plate; a winding drum notch for preventing the winding drum from interfering with the lifting support is formed in the lower part of the winding drum; a heightening support corresponding to the lifting support in height is arranged at the lower part of the core rod; the lifting support is fixedly connected with the lifting coil supporting plate through a plate inserting pin; the heightening support is fixedly connected with the core rod through the inserting plate pin; and a pin opening for the insertion plate to pass through is formed in the insertion plate pin. According to the utility model, by introducing the lifting support, the wire coil is lifted up after reaching a preset height, so that the bottom of the wire coil is prevented from being contacted with the temporary coil supporting arm, and the problems of falling impact and head scratch possibly occurring when the wire coil is separated from the temporary coil supporting arm in the prior art are solved.

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.

The present invention discloses a nose cone mooring structure for an airship without a rigid frame, comprising a base, a bottom ring, straight beams, and arc beams. The base is provided with several straight beams, and the ends of the straight beams away from the base are connected to the arc beams via a four-way structural member. The arc beams are provided with through slots that cooperate with fixing loops on the airship capsule. The bottom ring is provided between the four-way structural members. The nose cone mooring structure for an airship without a rigid frame of the present invention provides a simple nose cone fixing and mooring method for mooring an airship without a rigid frame. This solution can evenly distribute the tension during the mooring of the airship, does not require special reinforcement of the airship capsule material, and does not significantly increase the amount of capsule material used. At the same time, the nose cone can also ensure that the head portion, which is subject to greater air resistance, does not deform during the flight of the airship, thereby increasing the flight resistance.

A cryogenic insulated pipe having a conformally-bonded aerogel paper layer on an outer surface of the pipe. The conformally-bonded aerogel paper layer is conformally-bonded to the pipe outer surface in a pre-applied resin layer. An insulative blanket is applied over the aerogel paper, and a breathable protective layer is applied over the insulative blanket. Also disclosed is a cryogenic fuel tank for retrofitting a conventional fossil fuel-powered aircraft, or for a purposely built aircraft to run on hydrogen has an aerodynamically shaped outer surface including an ogive shaped nose cone, and a tapered tail cone, wherein the tapered tail cone includes actively adjustable elements for adjusting aerodynamic characteristics of the cryogenic fuel tank. The cryogenic fuel tank is configured to be attached below wings of the aircraft, through support pylons, which include sensors configured to measure forces applied by the cryogenic fuel tank to the airframe. The cryogenic fuel tank includes a nozzle and valve configured to vent gas from the cryogenic fuel tank by expansion through the nozzle in the event that the cryogenic fuel tank is j ettisoned from the aircraft.

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.

The present invention provides a fully automatic gluing and pressing device for the nose cone of an active touch pen, comprising a body, a gluing module, a pressing module, and a curing module; the pen tip clamping assembly of the pressing module comprises a movable bracket seat, a clamping bracket rotatably connected to the bracket seat, and a three-claw cylinder provided on the clamping bracket. When the clamping bracket rotates, the gluing needle of the gluing module can evenly apply the glue on the pen tip clamped by the three-claw cylinder along the circumferential direction. The operation is convenient, the gluing effect is good, and the safety is good. The pen tip loading assembly of the pressing module comprises a pen tip seat with a locating pin on the side and a floating pressure head for correcting and pressing the pen tip. When loading the pen tip, the locating pin and the locating hole can cooperate to ensure that the pen tip is inserted in a set posture and the alignment is accurate; the axis of the nose cone placement groove on the nose cone seat of the nose cone clamping assembly coincides with the axis of the three-claw cylinder. After the gluing is completed, the support seat can be moved to squeeze the pen tip into the nose cone. The pressing consistency is good and the yield rate is high.

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.

An integrated unit for de-icing the propeller blades and the nose of a turbomachine including: a rotary transformer including a static part and a rotary part; a DC-to-AC voltage converter connected at its output to the static part of the rotary transformer and intended to be connected at its input to a DC electrical power supply, and a power switch connected to the rotary transformer and configured to transmit electrical power to at least one pair of propeller blades and to the nose cone of the turbomachine.

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 utility model discloses an aerostat nose cone lock strength test device, which comprises a vertical loading device and a guiding device which are fixed on a heavy frame, a transverse loading device is arranged on the vertical loading device, the transverse loading device can move along the vertical direction under the driving of the vertical loading device, and the guiding device is arranged on the vertical loading device. The actuating cylinder in the transverse loading device can also move along a horizontal plane perpendicular to the axis direction of the nose cone lock test piece, the nose cone lock test piece is fixed on the heavy frame, the guide device is located below the nose cone lock test piece, and a pulley adapter plate in the guide device can move along a horizontal plane parallel to the axis of the nose cone lock test piece. And a guide pulley is arranged on the pulley adapter plate. The test device is convenient to install, can realize rapid and quantitative tensioning and fixing, and better simulates the boundary condition of the installation state of the test piece.

The present invention discloses a suborbital carrier rocket with an end cover and a first stage that can be recovered as a whole, comprising a first stage body and a rocket cover, the rocket cover comprising a barrel cover, an end cover shell segment and an end cover, the barrel cover is arranged at the front end of the first stage body, the barrel cover is used to place a suborbital payload, an adaptive load support is arranged at the bottom of the barrel cover, a slide rail assembly is arranged on the inner wall of the barrel cover, an electrically controlled locking device for locking the suborbital payload is arranged on the adaptive load support, the slide rail assembly is connected to the suborbital payload in a sliding manner; the end cover shell segment is arranged at the front end of the barrel cover, the end cover is arranged at the front end of the end cover shell segment, the end cover is composed of multiple petals hinged to the end cover shell segment, the multiple petals are hemispherical when closed, and the opening and closing of the end cover is controlled by an end cover actuator; a reverse thrust rocket is also arranged at the front end of the inner wall of the end cover shell segment; the present invention discloses a suborbital carrier rocket with an end cover and a first stage that can be recovered as a whole 100%, can be reused after inspection and evaluation, and has higher reuse value.

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.

The invention discloses an ultrahigh-temperature high-entropy boride ceramic material as well as a preparation method and application thereof. The preparation method of the ultrahigh-temperature high-entropy boride ceramic material comprises the following steps: 1) mixing metal oxide powder and B4C powder to prepare mixed powder; 2) pressing the mixed powder into a green body, performing electric field sintering, crushing and screening to obtain ultrahigh-temperature high-entropy boride ceramic powder; and 3) pressing the ultrahigh-temperature high-entropy boride ceramic powder into a green body, and then carrying out spark plasma sintering. The superhigh-temperature high-entropy boride ceramic material has excellent oxidation resistance and ablation resistance under the superhigh-temperature condition of 3000 DEG C or above, and can be used for superhigh-temperature thermal protection components such as the leading edge, the nose cone and the control surface of a hypersonic aircraft, and the preparation method has the advantages of being simple in operation and process, short in reaction time, low in cost and the like, and is suitable for industrial production. The method is suitable for large-scale industrial production and application.

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.

The invention provides a shark type unmanned cruise boat, and relates to the field of ships. The shark type unmanned cruise boat comprises a boat body assembly, an edge strip assembly, a peripheral balloon assembly and a bow cabin filling assembly. The edge strip assemblies are arranged on the outer sides, protruding outwards, of the middles of the two sides of the boat body assembly, and standing platforms are arranged on the upper sides of the edge strip assemblies. The peripheral balloon assemblies are arranged on the two sides of the boat body assembly and the streamline nose cone position of the boat body assembly. The standing platform of the shark-shaped unmanned cruise boat provides a standing position for an operator and also has the capability of improving the stability of the boat body assembly, and the peripheral balloon assembly has the capability of improving the stability of the boat body assembly while playing a buffering role in approaching a main boat, so that the stability of the boat body assembly is improved. And in cooperation with the bow tank filling assembly filled in the fore peak tank and the arc-shaped cover for sealing, the rollover resistance of the boat body assembly is improved, and unmanned remote operation of the boat body assembly is more convenient.

The present invention relates to a wing-rudder type structure and a manufacturing method thereof, and specifically to a wing-rudder head cone and hanging point integrated structure and a manufacturing method thereof, in order to solve the shortcomings in the prior art that when the cross-sectional area of the head cone changes greatly, the continuity of the composite material is difficult to ensure, and at the same time, it is difficult for the solidified material to enter the middle position of the cross-section with a larger thickness, resulting in uneven distribution of the solidified material in the middle position, which has a great impact on the bearing performance of the head cone. The wing-rudder head cone and hanging point integrated structure of the present invention includes a head cone formed by splicing a first semi-cone and a second semi-cone, an inner and outer wrapping rings arranged on the bottom surface of the head cone, and a rudder body fixed to the side of the head cone; the bottom surface of the head cone is coaxially provided with an annular boss, an inner and outer wrapping rings, and the rudder body is fixedly connected to the outer wrapping ring through the first and second hanging points; the manufacturing method of the present invention effectively improves the uniformity of the material density of the integrated structure and reduces the manufacturing cycle.

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.