95 results about "Cruise missile" patented technology

The invention relates to a remote high precision autonomous integrated navigation and positioning method, which is characterized in that a Strapdown Inertial Navigation System (SINS) is used as a main navigation system during the whole flight course of the aircraft, assisted by 3D high precision position and attitude angle information provided by celestial navigation system (CNS) based on the least square differential correction in boost phase (or middle segment). In reentry phase (terminal), using the characteristics of synthetic aperture radar (SAR), such as strong penetration capability, high resolving precision and all-weather, the SINS can be corrected through accurate location information and course information provided by SAR scene matching after motion compensation when the aircraft reentry into atmospheres, so the impact point (hit) accuracy of the aircraft can be increased and the invention has remarkable effects of eliminating or decreasing non-guidance error. The invention has advantages of autonomy and high precision, which can be used for improving remote ballistic missile, remote cruise missile, navigation and positioning accuracy of remote aircraft, such as long-endurance unmanned aerial vehicle, etc.

A missile, such as a cruise missile, has a nose payload portion having a frangible nose cover and a relatively hard target penetration nose cone. The nose cone may have a liquid fuel tank within, and a chemical energy explosive charge, such as a shaped charge, aft of the liquid fuel tank. The target penetration nose cone enables perforation of certain types of targets prior to detonation of the chemical energy explosive and the liquid fuel. The frangible nose cover is configured to be easily perforated or otherwise removed by the explosive force of the chemical energy explosive charge when the missile system is utilized for the attack of hard targets. The nose payload portion may have a fragmentation case, with one or more features designed to enhance fragmentation during detonation of the explosive and/or the liquid fuel.

A small gas turbine engine for use in an UAV such as a cruise missile, the gas turbine having a combustor forming a primary burn zone and a secondary burn zone, and in which fuel is injected into both the primary and the secondary burn zones by either a rotary cup injector or a plurality of fuel injector nozzles. The secondary burn zone with separate fuel injection allows for the diameter of the engine to be reduced in size but still allow for adequate power and efficiency to be reached for powering the vehicle. Air flow from the compressor is used to cool the combustor walls before being injected into the combustor, and to pass through and cool the guide nozzles and a main bearing located near the hot section of the combustor prior to being introduced into the combustor.

The invention discloses a method and a device for identifying target types. The method comprises steps that a sound signal of a target is acquired through a planar sound transmission array, the target includes any one of a low-altitude aircraft, a light aircraft, a power delta wing, a cruise missile and a power paraglider, the direction of the sound signal is evaluated, spatial filtering processing on the sound signal after direction evaluation is carried out, after spatial filtering processing, characteristics of the sound signal, characteristic vectors corresponding to the characteristics of the sound signal and characteristics of the target are extracted, type confidence of the target is determined according to the characteristics of the sound signal, the characteristic vectors of the sound signal and the characteristics of the target, the type confidence is identified through the fusion identification technology, and thereby the type of the target is determined. Through the method, type identification accuracy of the target is improved.

The invention relates to a passive radar detection method by use of synergistic or asynergistic satellite signals to detect low altitude targets, belonging to the technical field of space technology, which comprises a satellite and a ground passive radar system; wherein, the system comprises a plurality of detection stations and a data processing station, and interactive communications are conducted between the detection stations and the data processing station by means of radio or optical fiber transmission method. The signals, expanded frequency measurement codes and the signals reflected on low altitude aircrafts by and of the satellite, especially high radiation power satellite, are received by the detection station receivers of the passive radar with high sensitivity. After correlation, the path difference between the direct radiation signals of the satellite and the signals from the same source reflected by the low altitude targets can be derived; then, a group of passive detection equations can be established; the solutions can be obtained simultaneously. And a three-dimensional position coordinate of the low or super low altitude targets can be thus acquired. The invention has the advantages of huge detection area, small blind area, the capability of detecting low or super low aircrafts like cruise missiles and super low altitude airplanes passively at long range, and extensive application value for military use.

A cruise missile recovery system, for capturing a flying cruise missile and for guiding the cruise missile to a landing spot. The cruise missile recovery system has an intercept device for capturing the flying cruise missile, a radio-controlled parafoil connected to the intercept device for allowing the intercept device and cruise missile to be guided to the landing spot during descent of the intercept device and cruise missile, after capture of the flying cruise missile by the intercept device, and a position-stabilized suspension vehicle for suspending the radio-controlled parafoil device and the intercept device at a stable location above the earth, prior to capture of the flying cruise missile by the intercept device.

An optical impact system controls munitions termination through sensing proximity to a target and preventing effects of countermeasures on false munitions termination. Embodiments can be implemented on in a variety of munitions such as small and mid caliber that can be applicable in non-lethal weapons and in weapons of high lethality with airburst capability for example and in guided air-to-ground and cruise missiles. Embodiments can improve accuracy, reliability and lethality of munitions depending on its designation without modification in a weapon itself and make the weapon resistant to optical countermeasures.

The invention discloses a troposphere intermittent gliding flight control method and belongs to the flight control technical field. The troposphere intermittent gliding flight control method is particularly applied to an aircraft to increase flight time or range of the aircraft by reasonably using updrafts in a troposphere region. The method includes the following steps that: step one, heat updraft modeling and judgment are carried out; step two, aircraft modeling is carried out, and a control law when updrafts do not exist is designed; and step three, intermittent gliding flight control is carried out. According to the control method of the invention, atmospheric energy is utilized to perform intermittent gliding flight, and therefore, the flight time and range of the aircraft can be improved, and the control method has important significance for the aircraft to realize long-flight time, long-range and attack patrol operational goals. The control method of the invention can be widely applied to small-sized unmanned aerial vehicles and small-sized cruise missiles, and can effectively improve the combat effectiveness of the unmanned aerial vehicles and cruise missiles.

The invention discloses an amphibious cruise missile based on a throat offset type pneumatic thrust-vectoring nozzle and a posture control method of the amphibious cruise missile. The amphibious cruise missile comprises a missile body, a switching section and the throat offset type pneumatic thrust-vectoring nozzle. The advancing direction of the missile is changed and jumping out of water is achieved by controlling the tail throat offset type pneumatic thrust-vectoring nozzle. Compared with a traditional missile, the amphibious cruise missile has the beneficial effects that a rudder and other structures are removed, smaller radar reflection area is achieved, and therefore the stealth performance is better. When the throat offset type pneumatic thrust-vectoring nozzle works underwater, water serves as a working medium, and the beneficial effects that no gas is exhausted, no sailing path exists, the noise is small and the stealthiness is good are achieved; and when the throat offset type pneumatic thrust-vectoring nozzle works in the air, high-temperature fuel serves as a working medium of the thrust-vectoring nozzle, and the beneficial effects that reaction is sensitive, the flight speed is high, and the voyage is far are achieved.

The invention brings forward a speed control modeling method of a cruise missile. Through combination with a description of dynamic features of a ramjet engine, a composition and modeling flow of a speed control loop is given. The method specifically comprises modeling of the ramjet engine, a speed loop total model for simulation and a small-disturbance model for control. A foundation is laid for performance evaluation of the cruise missile and design of a controller. The speed control loop modeling method for simulation, brought forward by the invention, effectively reflects interface relations among a projectile body, the ramjet engine, atmosphere environment, a measurement apparatus and the controller, and provides a basis for realizing engine performance and total performance evaluation and technical indexes. The small-disturbance model modeling method for control, provided by the invention provides a reasonable design model for the design of a speed controller and can effectively improve the control quality.

The invention belongs to the technical field of scene matching and navigation, and specifically relates to a method for automatically selecting a scene matching area based on image information entropy, and aims to provide a selection method for realizing effective and automatic selection of a scene matching area for cruise missiles. It includes the following steps: the selection of scene information, the selection of feature stability and the selection of regional uniqueness. The present invention adopts the selection steps of scene information, feature stability and regional uniqueness, adopts the strategy of hierarchical selection and reverse elimination, reduces the calculation amount of scene matching area selection, and realizes the automatic, fast and effective scene matching area selection. The self-matching success probability of the selected scene matching area is >97%, the error variance is <10, and the selected matching area can meet the adaptability requirements.

The invention discloses a living body verification method, which comprises the following steps that: (1) in a visible spectrum imaging picture, obtaining position data where a face image is positioned; (2) calculating the position data of the face image in an infrared spectral imaging picture; (3) calculating a face image GLCM (Ground Launched Cruise Missile) feature parameter, and adopting a SVM(Support Vector Machine) to judge whether a face is a real face or not. The invention also discloses a living body verification system, which comprises a visible spectral imaging module, a module forobtaining visible spectral image position data, a switching position data module, an infrared structure light projection module, an infrared spectral imaging module, a GLCM feature parameter module and a SVM. The method and the system have the advantages that a copycat can identity hostile attacks when the copycat uses pictures, 3D masks and the like to carry out the hostile attacks on a face-based identity verification system, in addition, a verified person does not need to actively cooperate, so that verification time is shortened, and verification accuracy is improved.

Disclosed are supersonic missile anti-drag wings. Two sets of missile wings are distributed on two sides of a missile body symmetrically. Each missile wing comprises an upper wing piece and a lower wing piece. Root chord lengths of the upper wing pieces and the lower wing pieces are both 0.52m, and stretching lengths of the upper wing pieces and the lower wing pieces are both 0.52m. The upper reflex angle of each upper wing piece is -9 degrees, and the upper reflex angle of each lower wing piece is 9 degrees. The leading edge backward sweep angles of the upper wing pieces and the lower wing pieces are both 63.6 degrees. Each upper wing piece and each lower wing piece are respectively connected with a driving mechanism. When a Mach number of a supersonic cruise missile is lower than a design point, the upper wing pieces on the two sides of the missile body deflect downwards, the lower wing pieces deflect upwards so that the area ratio of the maximum thickness portion of the wings and the leading edge inlet is 1. Upper stream passes between the wings so that occurrence of jamming is avoided. When the Mach number continues to increase and reaches the designed Mach number, the upper wing pieces on the two sides of the missile body deflect upwards, the lower wing pieces deflect downwards, and the Busemann double-wing outward position is realized. Shock wave resistance can be reduced by utilizing wave interference effect of shock wave weakening and shock wave expanding.

The invention discloses a fabrication method for a cover body of a superhigh temperature ceramic matrix composite material ablation head antenna cover. The method comprises the following steps of weaving an ablation head flat plate fabric by a carbon fiber three-phase orthogonal mode to obtain an ablation head woven fabric; immersing and compounding to obtain C/SiC-ZrC composite ceramic ablation head rough blank, and carrying out profiling weaving on the cover body fabric by a needling mode to obtain a cover body woven fabric; mechanically processing the cover body woven fabric to required product sizes to obtain an ablation head and the cover body; and connecting and forming the ablation head and the cover body by nuts, thereby obtaining the cover body of the ablation head antenna cover. According to the method, the C/SiC-ZrC composite material and a C/SiC-HfC composite material which have excellent temperature resistance reaching 2,000 DEG C and ablation resistance are used; and through optimal design on a prefabricated element of the ablation head and control on volume content of annular fibers, zero expansion of the ablation head composite material in an annular direction can be achieved, and thus, the obtained cover body of the antenna cover can be applied to intermediate and remote land attack cruise missiles with high mach (greater than 10Ma), long endurance (greater than 1,000 seconds) and high accuracy.

A navigation system for a swarm of guided projectiles, such as cruise missiles, having three or more projectiles launched at a target is provided. The projectiles can be in secure communication with one another and can operate as a single swarm using a fusion of angle of arrival and time difference of arrival of a detected signal to direct the projectiles to a point of impact on a target that is offset from an emitter generating the detected signal. The projectiles can further use knowledge of their relative position to maintain and/or adjust their flight path and/or speed to maintain their course to the point of impact in areas of GPS denial or unavailability.

The invention discloses an ultralow-thermal-conductivity stable-high-temperature-phase neodymium-cerium compounded zirconate thermal barrier coating material and a preparation method thereof. The thermal conductivity of the material is 60% lower or more than 60% lower than that of an existing thermal barrier coating material, and is also lower than the thermal conductivity of a rare earth zirconate material of the same kind; and on the other hand, no phase changes are generated within the interval between the room temperature and 1600 DEG C. The material is more beneficial to increase of the thermal simulation effect and working temperature of the thermal barrier coating, and can be used for the aerospace fields of aviation turbine engines, gas turbines, hypersonic-speed cruise missile hot end parts and the like.

The invention relates to a large-caliber supersonic target projectile for testing or training. The large-caliber supersonic target projectile comprises a rocket engine and a target projectile, wherein the rocket engine comprises a combustion chamber, a jet tube and a plurality of fins; the front end of the combustion chamber is connected with the target projectile, and the back end of the combustion chamber is connected with the jet tube; the fins are distributed on the outer circumference of the jet tube; the target projectile comprises a shell, a counterweight body and a counterweight ring; the shell has a hollow structure; the counterweight body is embedded and fixed at the inner front end of the shell; and the counterweight ring is arranged on the middle part of the shell, and is fixed by using a preset separation blade. According to the large-aperture supersonic target projectile, on one hand, the gunpowder loading structure and gunpowder loading scheme of the rocket engine are improved, so that the thrust force of the engine is increased; on the other hand, the weight of a head structure is reduced, so that a high launching speed is achieved for the complete projectile; and target characteristics such as geometry, speed, infrared rays, radar reflecting area and the like at the final flying stage of a cruise missile can be well simulated, so that the requirements of testing or training on the similarity, economy and repeatability of a simulated target are met.

The invention discloses a new casting method, wherein, a casting mold with a pressurizing device is immerged in the inside of molten metal (8) for casting; the pressuring device is driven to carry out pressurization by utilizing lifting force and the gravity of the casting mold or other power. The molten metal in a die cavity (12) is filled, fed and solidified under necessary pressure during the lifting process, and then the casting mold is immerged in water for chilling so as to improve the mechanical property of castings. An upper mold (4) and a lower mold (10) of the casting mold are respectively provided with a plunger hole which is communicated with the die cavity (12) by a feeding cavity; the feeding cavity is slantingly arranged at the thickness changing site of the castings, and a driving mechanism is connected on the plunger to be matched with the plunger hole. The molten metal in the die cavity (12) is pressurized under the action of the driving mechanism, thus causing the castings to be filled, fed and solidified under single directional, double directional and multi directional high-pressure pressurizing. The method is suitable for the casting of various complicated aluminum alloy castings such as the cylinder of an auto engine, wheel hub and extra large thin wall aluminum alloy castings such as cruise missile hull with 0.8m in diameter, 6m in length and 6-8mm in thickness and various common castings.

The invention discloses a distributed storage structure-oriented handling method. An intensive computing-oriented five-layer parallel compiling frame is adopted and comprises an intensive computing application compiling layer for generating a redundancy-removed multi-atom application parallel area, an atom application layer based on a data parallel model, a data parallel compiling layer used for translating an application instruction set of the atom application layer into SPMD parallel codes based on a multi-core runtime system, a parallel runtime layer for finishing task scheduling and data flow scheduling, and a local compiling layer and an operating system layer for supporting the parallel runtime layer. Based on a model of the handling method, embedded type multi-core MATLAB can be developed, a programming effect of 'what you see is what you get' is implemented, and a simple and efficient multi-core programming platform is provided for complex algorithms such as matrix operation, pattern recognition and machine learning for rapid R&D and deployment of new-generation cruise missiles, air defense missiles and smart unmanned aerial vehicles and the like.

The invention provides a 3D printing cruising missile lightweight design and manufacturing method comprising the steps that (1) data collection is conducted on a part, and treatment and model characteristic fitting are conducted on collected data; (2) reverse reconstruction is conducted on a solid model through a B-Rep method in three-dimensional design software to obtain a three-dimensional digital model of the part; (3) structural performance analysis is conducted on the SLM formed part to verify related performance indexes; (4) on the basis of a constant strength principle, part inner structure redesign is conducted on the reconstructed model through a skin-skeleton structure; and (5) the final lightweight model is used for generating an STL model, an SLI model is generated after part data hierarchy is conducted on the STL model, the SLI model is imported into SLM equipment, corresponding technological parameters are set, and the final lightweight solid part is obtained after machining through selective laser melting forming technology. Through the 3D printing cruising missile lightweight design and manufacturing method, weight and manufacturing time of the SLM formed part can be reduced, the lightweight high-performance aeronautical part is obtained, and meanwhile machining time is reduced.