118 results about "Launch control" patented technology

The invention discloses a fire fighting truck applicable to high-rise and super high-rise building fire suppression, comprising a truck chassis (1), a command control equipment (2), an launch control equipment (3), an equipment compartment (4), an launching apparatus (5), an ejection device (6), a fire extinguishing bomb (7), a turret control apparatus (8) and a photoelectric detection equipment (9), wherein, the command control equipment (2) and the launch control equipment (3) are arranged in the co-pilot position of the cab of the truck chassis (1); the equipment compartment (4) is arranged behind the cab and is fixed on the truck chassis (1) with bolts; the launching apparatus (5) is arranged on the truck chassis (1) and is fixed with screws; the ejection device (6) is arranged and fixed on the launching apparatus (5); the fire extinguishing bomb (7) is arranged within the ejection device (6); the turret control apparatus (8) is distributed on the truck chassis and the launching apparatus for accomplishing the actions of leveling the truck body and controlling the launching apparatus.

The present invention relates to a launch control unit for selectively engaging a revised shift control strategy. The revised shift control strategy is operable to control an internal combustion engine and / or an automatic transmission of a motor vehicle. The launch control unit can be configured to receive a brake signal from a foot brake and a torque demand from a throttle pedal. In use, the launch control unit is operable to engage said revised shift control strategy upon receipt of a torque demand greater than or equal to an upper torque threshold within a predefined time period of receipt of a brake signal less than or equal to a lower brake threshold.

The invention is applicable to the technical field of rocket control, and provides an omni-directional launch control method and device for a carrier rocket as well as computer equipment and a storagemedium. The method comprises the following steps that an initial attitude angle is acquired, moreover, an initial quaternion is calculated, wherein the initial attitude angle comprises an initial rolling angle; a real-time angle increment of the current control period is acquired, and the real-time quaternion is calculated; a standard quaternion is calculated according to the initial rolling angle; an attitude angle deviation is calculated on the basis of the real-time quaternion and the standard quaternion, and a target attitude control signal is obtained according to the attitude angle deviation and is output to an execution mechanism. The initial quaternion is calculated through the initial attitude angle to obtain the standard quaternion and the real-time quaternion, the attitude angle deviation is obtained through calculation and is resolved to obtain the target attitude control signal for controlling the rocket, omni-directional launch control of the rocket is realized without rotating an indexing mechanism, the indexing mechanism is omitted, so that the cost is saved, the time for rotating the rocket for alignment is shortened, and the rapid response capability and the automatic level of the rocket are improved.

The present invention relates to launch control of a vehicle having an automated manual transmission and an automated clutch selectively engaging an output of an engine to an input of the transmission. When a request for vehicle launch is detected, an output speed of the automated clutch and a desired clutch output torque are obtained. The automated clutch is partially engaged to allow for slip between an input to the clutch and an output from the clutch, and a desired amount of the slip between the input to the clutch and the output from the clutch is determined. A desired engine speed is calculated based on the desired amount of slip, with the engine torque automatically adjusted to obtain the desired engine speed.

A method for controlling the launch of a vehicle having a powertrain that includes a first power path for driving a first wheel set and including an engine, a transmission having an input driveably connected to a crankshaft of the engine, a current gear, an input clutch associated with the current gear and an output connected to first wheel set, and a second power path including an electric machine for driving a second wheel set. The method includes determining a desired magnitude of wheel torque to be produced at the first wheel set, using the desired magnitude of wheel torque to be produced at the first wheel set to determine a magnitude of crankshaft torque, producing a predetermined magnitude of torque capacity at the input clutch, determining a crankshaft speed error, determining a change in magnitude of torque at the crankshaft that will reduce the crankshaft speed error, and operating the engine to produce a desired engine torque, whose magnitude is determined from the change in magnitude of torque at the crankshaft and the magnitude of desired crankshaft torque.

The invention provides a launching method of a guided missile safe launching system. The safe launching system comprises a launching control system and a launching box / barrel. The launching control system comprises an ignition power source and a launching control unit. The launching control unit comprises a combat launching command relay, an engine ignition relay and a resistor. The launching box / barrel comprises a napalm bomb mechanism microswitch. After an initiating explosive device is connected with the resistor in parallel, one end of the circuit is sequentially connected with the napalm bomb mechanism microswitch, a first set of contact switches of the engine ignition relay and a first set of contact switches of the combat launching command relay in series; the other end of the circuit is grounded and then is sequentially connected with a second set of contact switches of the engine ignition relay and a second set of contact switches of the combat launching command relay in series. According to the safe launching system, the safety performance in missile launching can be improved.

The embodiment of the invention provides a carrier rocket test system, and relates to the technical field of rocket test launching. The system comprises an execution module used for receiving a parsed test launch control instruction sent by a parsing module and executing the parsed test launch control instruction; the execution module comprises at least one execution device; the execution module comprises a rear-end measurement and control device, and the rear-end measurement and control device is used for receiving the analyzed test launch control instruction, determining the instruction type of the analyzed test launch control instruction, determining a target execution device for executing the analyzed test launch control instruction from at least one execution device according to the instruction type, and sending the test launch control instruction to a target execution device; and a target execution device which is used for executing the analyzed test launch control instruction and returning an execution result to the rear-end measurement and control device. According to the embodiment of the invention, automatic testing in the test launch control process of the carrier rocket is realized, and a large amount of manual operation and manual interpretation are avoided.

The invention discloses a launcher for a meteorological sounding rocket. The launcher for the meteorological sounding rocket comprises a connecting rod, a shaft pin A, a shaft pin B, a first gasket, afirst nut, a second nut, an orientator, a supporting seat and a traction chassis. The launcher for the meteorological sounding rocket is stable in overall structure and suitable for operation under severe weather conditions; a launching angle adjusting mechanism adopts a connecting rod mechanism for driving a screw rod to transmit by a motor; the orientator adopts four guide rails to be combinedwith a middle ring A and a middle ring B to form a launching rail, the launching caliber is 122 mm, the maximum angle of the high-low launching angle is not less than 80 degrees, the height of the launcher in the transportation state is not larger than 3 m, and the stability is high; a supporting part adopts the motor to drive a trapezoidal lead screw shaft to rotate to drive the connecting rod tomove front and back so as to achieve the purpose of adjusting the launching angle, and a manual mode can be quickly switched; and a rocket operation automatic parameter recording instrument or a wireless launching controller can be used for control, and a high-power lithium battery is carried, so that the field power supply problem is solved.

The invention discloses a test launch control system and method. The system comprises a test launch control rear end, a telemetering device and a launch device; the test launch control rear end comprises a main controller, a data server and a switch, the main controller and the data server are both connected with the switch, and the main controller is used for obtaining data of the data server through the switch; the telemetering device is connected with the switch and is used for performing wireless communication with a target object and storing the data of the target object in the data server through the switch; and the launch device is used for launching the target object. A rocket test launch control system in the prior art seriously depends on a launch site, copper core cables need to be arranged due to the fact that the distance between a front end and a rear end is long, and the manpower and material resource cost is high. With the test launch control system and method of the invention adopted, the above technical problem is solved.

The embodiment of the invention provides a ground test launch control system. The ground test launch control system comprises: a test launch control computer in which a computer program for controlling rocket test and launch is pre-stored; and a universal interface adapter which is in data communication with the test launch control computer through an Ethernet communication bus. The universal interface adapter comprises a back plate and rocket-ground communication interface modules arranged on the back plate, each rocket-ground communication interface module is in data communication with the back plate through a back plate bus, and each rocket-ground communication interface module is further used for being in data communication with electrical equipment on a rocket. The universal interfaceadapter is used for converting an instruction sent by the test launch control computer into a control signal of the on-rocket electrical equipment and sending the control signal to the on-rocket electrical equipment through the rocket-ground communication interface modules, and is used for receiving a working state signal fed back on the rocket through the rocket-ground communication interface module and sending the working state signal to the test launch control computer. The ground test launch control system provided by the embodiment of the invention realizes miniaturization and multifunction of ground test equipment.

The invention discloses an aircraft and ground cooperative test launch control system and control method, and relates to the technical field of aircraft test launch control. The system comprises a remote launching controller which is used for generating a third control instruction for controlling the launching of an aircraft, a ruggedized computer which is used for sending a first control instruction and a second control instruction, and forwarding the third control instruction generated by the remote launching controller to a test-launch-control all-in-one machine, a ground power supply whichis used for supplying power to the test-launch-control all-in-one machine, and the test-launch-control all-in-one machine is used for sending a test instruction to other equipment on the aircraft andsending obtained test data to the ruggedized computer; and when the ruggedized computer judges that the test data are normal, the test-launch-control all-in-one machine is further used for carrying out launch control on the aircraft. According to the invention, an integrated aircraft and ground cooperative test launch control architecture which takes tests on the aircraft as a main part and takesground tests as an auxiliary part is constructed, so that overall miniaturization, light weight and portability are realized, and the operation complexity and the use and maintenance cost of the system are reduced.

The invention provides a multi-orbit design method for carrier rocket fire-striving launching and a fire-striving launching control method. The fire-strike launching multi-orbit design method comprises the steps that task data of P Mars orbit entrances specified by a fire-strike launching task are obtained, the task data of each Mars orbit entrance comprises a launching window where a detector is located at the Mars orbit entrance, and P is a natural number larger than 2; an iterative optimization method is adopted, P fire-strike launching orbits where the carrier rocket carries the detector to fly are designed, and target points of the P fire-strike launching orbits are in one-to-one correspondence with the P Mars orbit entrances; and when the carrier rocket is controlled to fly according to any one of the P fire-striving launching orbits, the carrier rocket carries the detector to the Mars orbit entrance corresponding to the fire-striving launching orbit. According to the design method, the design efficiency is improved, the complexity of flight data is reduced, and reliable launching and accurate orbit injection of the large low-temperature carrier rocket can be realized.

The invention discloses a carrier rocket digital power supply and distribution control system and a test method thereof. The system comprises a plurality of switches. The plurality of switches receivedigital instructions sent by the ground test launch controller through the Ethernet so as to execute closing or opening actions; the carrier rocket digital power supply and distribution control system sends collected test data to the ground test launch control system through the Ethernet for analysis. Digital instructions and test data are transmitted through the Ethernet, and the problem that acable used for ground test launch control is not used any more after a rocket takes off and becomes redundant load on the rocket is solved.

A scalable and distributable software architecture for use in conjunction with various different weapons control system and launch systems is disclosed. The architecture discards the proprietary and non-open protocols and services that characterize in favor of open source adaptive and middleware components. In the illustrative embodiment of the invention, the inventive architecture is implemented as a Launch Control Module that separates different layers of responsibility within a launch controller (e.g., LCU) and exposes its variation points.

An electro-pneumatic control system for a motorcycle transmission. A pneumatic clutch cylinder used to disengage the clutch by pulling in a clutch lever via a clutch cylinder rod. Optionally, manual override of the clutch control may be provided for by linking the clutch lever to a clutch cable by a through rod and slip device. To change gear a pneumatic gear change cylinder, attached to the gear change lever, push-pulls lever to upshift and downshift transmission. The system is equipped with launch control that allows the motorcycle to start from a standstill.

The invention provides a launch control device and method suitable for dual-missile launch. The launch control device comprises: a first transceiver module, a second transceiver module, a DSP processor module, a gating module, and an upper computer module; the first transceiver module is connected with the DSP processor module; the upper computer module and the DSP processor module are connected in series through the second transceiver module; the gating module is connected in series with the DSP processor module and the upper computer module, respectively, and the gating module outputs the information of any one information channel or any number of information channels to the upper computer module according to a gating instruction and real-time information output by the upper computer module. The launch control device provided by the invention is reasonable in design and is convenient to use, provides an information channel gating function, especially a video gating function, in a missile launch process, which is beneficial to controlling the missile launch in a missile weapon system with an optical imaging guide device.

The invention relates to a launch control unmanned aerial vehicle and belongs to the field of aircrafts. The launch control unmanned aerial vehicle comprises an unmanned aerial vehicle body, an undercarriage, a high-pressure gas cabin and a controller. The undercarriage is arranged below the unmanned aerial vehicle body. The controller is arranged on the unmanned aerial vehicle body. The high-pressure gas cabin is arranged on the unmanned aerial vehicle body and connected with the controller. An internal through hole of the undercarriage is connected with the high-pressure gas cabin. Accordingto the launch control unmanned aerial vehicle, the principle is simple, the structure is simple, launch control is achieved, the head-start effect is achieved, a power supply is saved, and the higherendurance ability is achieved.

The invention discloses an automatic launch control method for a carrier rocket low-temperature power system. According to the method, the whole launch process is subjected to stage planning accordingto the characteristics of different time period processes of a launch day; each stage is divided into a plurality of items of work according to the independent degree of the work content, and each item of work is controlled and executed by an independent automatic control program; and due to the fact that the interpretation data is more, the algorithm is complex, the fault detection and positioning difficulty is high, and the processing requirement is high, a data processing knowledge base for fault diagnosis is constructed while the automatic launch control is executed, the important data inthe flow is interpreted, the rapid fault positioning is realized, and the fault processing is automatically executed.

The invention discloses an integrated measurement, launch and control method and device for a liquid carrier rocket. The device comprises a first measurement and control combination and a second measurement and control combination which are connected with each other through a cable. The first measurement and control combination is in network connection with the rear-end control computer, is in cable connection with the liquid carrier rocket, and is used for receiving control of the rear-end control computer and performing related detection and control on an electrical system on the rocket in the measurement and launching process of the liquid carrier rocket; the second measurement and control combination is in network connection with the rear-end control computer, is in cable connection with a ground controlled object, and is used for receiving the control of the rear-end control computer and performing related detection and control on the ground controlled object in the measurement and launch process of the liquid carrier rocket; therefore, the ground controlled object and the rocket electrical system are matched with each other in the testing and launching process of the liquid carrier rocket.

The invention provides a figure rocket launching safety locker based on MCU microprocessor control. The figure rocket launching safety locker comprises a controller, and an operation module, a launching control module, a communication module and a power supply module which are connected with the controller, and a shell for placing hardware equipment. The controller is an MCU microprocessor and isin communication connection with eahc module through UART. The controller is in two-way communication connection with the operation module, the launching control module and the communication module; the power supply module supplies power to all other modules in a one-way manner; the launching control module comprises a digital delay tube rocket launch controller and a capacitor discharge type rocket launch controller. The ignition voltage is not greater than 80v; a safety locking function is integrated on a traditional rocket launching device, so that rocket launching by non-operators can be prevented, information such as launchers and launching conditions can be recorded and uploaded to a server, unified management is facilitated, and the launching authority of the device can be releasedin time through a remote instruction.

The application discloses an ignition safety system of a multistage rocket engine and a detection method of the ignition safety system. The ignition safety mechanism of a multistage rocket engine comprises a ground testing and launching control system, an N-stage ignition safety mechanism and an N-stage intelligent detection combination, wherein the N-stage intelligent detection combination is used for detecting and recording the operating states of corresponding all stages of ignition safety mechanisms, and transmitting the operating states to the ground testing and launching control system;and the ground testing and launching control system is used for detecting the insurance states and the operating states of the ignition safety mechanisms and receiving the recorded operating state ofthe ignition safety mechanisms, detected by the N-stage intelligent detection combination, and according to the insurance states and / or the operating states of the ignition safety mechanisms, detecting the safety states of the ignition safety mechanisms. Through the adoption of the ignition safety system of the multistage rocket engine and the detection method of the ignition safety system disclosed by the invention, all the stages of the ignition safety mechanisms are connected in series, so that the design complexity of a cable network is simplified, and the weight of the cable network is reduced.

The invention relates to the technical field of space flight and aviation, and discloses a carrier rocket control system and method. The system comprises a ground test launch control system, a carrier rocket body, a sensor, a central computer, an output controller and an execution mechanism, and the output controller is connected with the ground test launch control system, the central computer and the execution mechanism. The carrier rocket body is connected with the sensor and the execution mechanism, the sensor is connected with the central computer, and the control process of the carrier rocket body comprises a ground test stage, a pre-launching preparation stage and a flight stage. The technical scheme of the embodiment of the invention has the advantages of strong real-time performance, high reliability, multitask processing, easy expansion and the like.

The invention discloses an artificial weather modification catalytic operation device as well as a control system and control method thereof, and belongs to the technical field of artificial weather modification safe operation. The operation device comprises a parachute compartment, an instrument compartment, a control compartment, a battery compartment, a rotor compartment, a catalyst scatteringcompartment and an engine compartment which are arranged in sequence in a rocket body as well as a tail wing arranged at the tail of the rocket body, wherein a control circuit, a navigation circuit and a safe lunch circuit are arranged in the control compartment, and the control circuit controls rocket ignition, device flight and catalyst scattering. The novel catalyst scattering operation deviceis improved and designed by organic combination of a rotor unmanned aerial vehicle and a rocket, and safe lunch of the operation device is carried out through an improved lunch controller; and compared with the prior art, the operation device improves the accuracy, effectiveness and safety of artificial weather modification catalytic operation, realizes reuse and greatly reduces the cost of artificial weather modification operation.

An unmanned aerial vehicle launching frame is characterized in that the unmanned aerial vehicle launching frame comprises a base frame, a front supporting assembly and a rear supporting assembly; the front supporting assembly is arranged at the front end of the base frame to support the front portion of an unmanned aerial vehicle, and the rear supporting assembly is arranged at the rear end of the base frame and can support the tail portion of the unmanned aerial vehicle; the rear supporting assembly comprises a rear supporting column, a U-shaped support and a self-locking spring, the rear supporting column is fixed to the base frame, and the two ends of the U-shaped support are hinged to the rear supporting column; the self-locking spring is connected with the base frame and the U-shaped support, the self-locking spring acts the elastic effect on the U-shaped support which is further provided with a placement frame matched with the an unmanned aerial vehicle tail cross beam, and the top end of the placement frame is provided with a U-shaped groove; and the unmanned aerial vehicle tail cross beam can be placed in the U-shaped groove. The unmanned aerial vehicle launching frame is of a pure mechanical structure, a control circuit does not need to be controlled, the structure is simple, the weight is low, carrying is convenient, the use range is wide, the requirements for unmanned aerial vehicle launching under different bad conditions can be met, and a launching controller is not needed.

The present invention relates to a control system for a vehicle powertrain (3). The vehicle powertrain (3) includes a transmission (27), a launch control clutch (31), and a transfer case (39) selectively operable in a high range and a low range. The control system includes a monitor (53, 55) for monitoring one or more operating parameters of the launch control clutch (31) and / or the transmission (27). The control system comprises a controller (49) configured, in dependence on said monitoring means determining that said one or more operating parameters are above a predefined operating threshold or outside a predefined operating range, to output a notification to the user to select said low range; and / or to output a transfer case control signal to the transfer case (39) automatically to select said low range.

The invention discloses a wireless temperature measurement emission controller for the low-frequency acquisition of the surface temperature of a piston of an internal combustion engine and belongs to the technical field of engines, electronic circuits and communication. The wireless temperature measurement emission controller comprises a temperature signal acquisition module, the output end of the temperature signal acquisition module is connected with an AD conversion module, and the output end of the AD conversion module is connected with a wireless emission module. According to the controller of the invention, the temperature signal collector and the AD module are designed, so that the whole structure of the controller is small and light, the installation of the controller in a narrow and complex environment in an engine cylinder is facilitated, the temperature measurement requirements of reliability and thermal load tests in an engine are met, power supply to the controller is realized through a power supply module, the anti-static capability of the controller is improved, abnormality caused by transient interference events such as static electricity is prevented, the stability of the working state is improved, the controller effectively adapts to the complex motion environment of the piston, and improves the reliability of real-time measurement.

The invention discloses a movable launching and recycling integrated platform for a fixed-wing unmanned aerial vehicle. The platform comprises hydraulic lifting launching bows, launching rail frames, an unmanned aerial vehicle mounting base, a quick disconnector, a launching control device and a traction power cable mechanism. Each hydraulic lifting launching bow is composed of two connecting ends and a driving end. The driving end of each hydraulic lifting launching bow is connected with a driving mechanism and used for providing lifting force and launching angles for the unmanned aerial vehicle. The launching rail frames are arranged on the left side and the right side respectively and connected with the four connecting ends of the front hydraulic lifting launching bow and the rear hydraulic lifting launching bow. A sliding groove is formed in each launching rail frame and used for movably installing an unmanned aerial vehicle installation base. The quick disconnector is mounted at the top end of the unmanned aerial vehicle installation base and used for fixing or disengagement of the unmanned aerial vehicle; the traction power cable mechanism is used for pulling the two unmanned aerial vehicle installation bases to move along the launching rail frames. According to the integrated platform, the work of taking-off launching and landing recovery of the unmanned aerial vehicle can be completed through one device.

In order to solve the problem of poor portability in the existing rocket test and launch control system adopting a remote measurement software based on a windows platform to monitor flight data, the disclosure provides a Qt-based liquid rocket test and launch control system time sequence monitoring method to improve the portability of the rocket test and launch control system. The method comprisesthe steps that an ignition command is sent to a flight controller on the rocket; flight data issued by the flight controller on the rocket are received; the time sequence position of the flight datais monitored, and flight data when the time sequence position changes are displayed; the first data are distributed to a QCustomPlot drawing assembly, wherein the first data are the data required forbeing drawn in the flight data; and the QCustomPlot drawing assembly is controlled to perform real-time drawing according to the first data. The technical solution of the disclosure is based on the QCustomPlot drawing assembly to draw to monitor the data issued by the flight controller on the rocket, so that the liquid rocket test and launch control system has higher portability.

The invention discloses an electromagnetic gun control method and device adaptive to multiple bullets. The electromagnetic gun control device comprises a keyboard system, a main control MCU, a Boost circuit and a launch control circuit, wherein a controller inputs a launch distance, a launch angle and bullet mass through the keyboard system, the main control MCU acquires launch voltage of the bullets according to the launch distance, launch angle and mass of the bullets, the Boost circuit is regulated according to the launch voltage so that output voltage of the Boost circuit reaches the launch voltage, and the launch control circuit is switched on to generate launch electromagnetic force so as to launch the bullets. The method can adapt to different bullets, the launch voltage can be quickly calculated according to the bullet mass, and the bullets are launched precisely to hit a target.