127 results about "Near space vehicle" patented technology

The invention discloses a fault-tolerant sliding-mode control method for a near-space vehicle. According to the fault-tolerant sliding-mode control method for the near-space vehicle, for the situation that the order of the magnitude of external disturbance in a quick loop and slow loop system is greatly larger than the order of the magnitude of uncertain times of the system, the nonlinear disturbance observer technology is used for processing hybrid disturbance, and unknown hybrid disturbance is estimated by a disturbance observer through known system information; in order to solve the problem that saturation of the control surface of the near-space vehicle is limited, the upper bound of the deflection angle output of a steering engine is applied to design of a control law, it is guaranteed that the input is within a certain range, auxiliary variables are designed, the deflection angle output of the steering engine is automatically adjusted through the self-adaption law, and therefore the situation that when the upper bound of the deflection angle is too large, the output is too large is avoided; a compensator is established through a radial basis function neural network and is used for fault-tolerant compensation when the steering engine breaks down, and therefore the problem that the steering engine of the near-space vehicle breaks down is solved. By the adoption of the fault-tolerant sliding-mode control method for the near-space vehicle, under the conditions of system uncertainties, unknown external disturbance, limited input saturation and a fault of the steering engine, the near-space vehicle has good control performance.

The invention discloses a hypersonic aerocraft and air inlet internal and external waverider integrated design method, and relates to a near space aerocraft. An aerodynamics characteristic is firstly appointed, and then a design scheme meeting the characteristic is inferred backwards; a three-dimensional shock wave curved surface in a complex shape is appointed, the change rule of the transverse curvature center is obtained, and a series of basic flow fields meeting the needs of the waverider design are inferred backwards according to the change rule; flow lines of different curvature centers and different radial positions are traced in every basic flow field in the circumferential direction; a waverider device capable of producing the appointed complex three-dimensional shock wave curved surface is obtained finally, namely the integrated design scheme is obtained. The advantages of a waveriders and an internal waverider air inlet are kept, the integrated design of the two high-performance devices is achieved, the waverider model with high lift-drag ratio and the scheme of the air inlet with full-flow capture can be obtained at the same time, and accordingly the overall performance of the aerocraft is improved.

The invention discloses a near space dynamic plasma sheath channel modeling and simulating method. The method comprises the steps that channel resources are partitioned according to large-scale space-time-frequency characteristics to construct a large-scale space-time-frequency channel resource matrix, wherein the height corresponds to large-scale space domain characteristics, the speed corresponds to large-scale time domain characteristics, and frequency points correspond to large-scale frequency domain characteristics; for all channel resource matrixes, dynamic small-scale channel parameters are obtained; small-scale channel model simulating is achieved according to given small-scale channel parameters; large-scale channel model simulating is achieved by repeating small-scale channel models at multiple states on a near space flight path. The method is suitable for channel modeling under different dynamic rules, random-amplitude phase probability density distribution which may occurs under a plasma sheath is simulated, and channel modeling feasibility is guaranteed; the method is suitable for near space aircraft plasma sheath channel modeling and spaceflight reentry-return spacecraft plasma sheath channel modeling.

The invention discloses a dynamic-actuator aircraft attitude distributed type fault-tolerant control system. The system comprises an actuator fault detection and identification unit, an auxiliary system based on an observer and an inversion fault-tolerant control algorithm based on order filtering. According to the distributed type fault-tolerant control system, the actuator fault detection and identification unit is used for obtaining fault information of actuators in real time, the self-adaptive sliding-mode observer is obtained according to an attitude angular speed ring, the observer has strong robustness, an undetermined or interference upper bound does not need to be known, and control surface damage fault information an interference information are both hidden in the observer; a fault-tolerant controller is obtained based on an observer model. Robustness fault-tolerant control over various faults of different types under the multi-fault condition is achieved, the fault-tolerant control system is applied to near-space vehicle attitude stable control and tracking control under the condition of faults of the actuators and the control surface, flying attitude robustness fault-tolerant control is achieved, and good control performance and effect are achieved.

The invention relates to a method for apperceiving and reconstructing a non-vision structural form of a near space vehicle model. The method comprises the following steps: on the basis of a distributed embedded dispersed bragg fiber grating sensing array, sensing and detecting strain information of each measuring point of a test model, and utilizing a certain curvature conversion algorithm to convert the strain information into curvature information; on the basis of a curvature interpolation, calculating the related coordinate value of each point of wings and frame of the test model according to a three-dimensional space curve fitting algorithm of orthogonal curvature information; utilizing a coordinate fuse algorithm to obtain an absolute coordinate value of each point of the model; and lastly, reconstructing the deformation state and the low-frequency vibrating form of the test model according to a computer graphic processing technology and realizing the three-dimensional visual real-time display for the change in the model form. An experimental system for the method comprises a simulated near space vehicle model, a fiber grating network analyzer with a network service function, a high-performance computer and a display, wherein the fiber grating sensing array is embedded into the surface of the simulated near space vehicle model.

The invention discloses an aircraft attitude fault-tolerant control method based on dynamic control redistribution. The method includes the steps that firstly, an attitude angle rate loop dynamic equation of a near-space aircraft X-33 is given, and a control surface jamming and damage fault model is established; then according to the control surface damage fault model of an attitude angle rate loop, a self-adaptive sliding-mode observer is designed; the attitude angle loop and the designed observer dynamic equation are combined, and an angle loop controller and an angular speed loop controller are designed. Accurate damage fault information, interference information and the like are not needed, the information is implied in the designed self-adaptive sliding-mode observer, the implicit information is fed back to a pseudo control law and online distribution algorithm LMI in real time, and then robust fault-tolerant control is achieved. Finally, the designed method is applied to attitude stability control and tracking control of the near-space aircraft under the control surface fault condition, flight attitude robust fault-tolerant control is achieved, and good control performance and effects are achieved.

The invention discloses an integration design method for a hypersonic slender body air vehicle and a three-dimensional inward rotation air inlet channel, and relates to near space aerocrafts. The method comprises the following steps that a three-dimensional shock wave hook face is designated according to the design requirement; a series of basic flow fields are derived reversely according to the relation between the discrete shock wave curve and the curvature center; a slender waverider vehicle lower surface molded line and the outlet section of the three-dimensional inward rotation air inlet channel are designed, and streamline tracing in the counter flow direction is conducted in the basic flow fields; the two-dimensional projection shape of a lip of the three-dimensional inward rotation air inlet channel is designed, and a three-dimensional configuration of the lip of the three-dimensional inward rotation air inlet channel is obtained according to the shock wave relation; and the hypersonic slender body air vehicle is subjected to geometric construction based on a compression molded face. The performance of an outer waverider vehicle front body and the performance of the three-dimensional inward rotation air inlet channel are simultaneously considered by the generated slender body hypersonic air vehicle and the air inlet channel, the lift drag characteristic is high, full-flow incoming flow capture can be ensured, engine thrust is increased, outflow resistance is reduced, the operation Mach number range of the air inlet channel is widened, and natural transition of an inner waverider part and an outer waverider part is achieved.

The invention designs an actuator fault detection and identification unit based on multiple observers and a decentralized fault-tolerant control framework for a near space aircraft with disturbances and parameter uncertainty by considering the problem of robust fault-tolerant control after the aircraft has actuator and control surface faults. A system can acquire the fault information of an actuator in real time. A self-adaptive sliding mode observer is designed for an attitude angular velocity loop. A designed observer has strong robustness, does not need to know the upper bound of the uncertainty or the disturbances. The information of a control surface damage fault and the disturbance information are fully implied. A fault-tolerant controller is designed based on an observer model. Through a designed fault-tolerant control system, robust fault-tolerance control is realized under the conditions of various different types of faults and multiple faults. The designed method is respectively applied to posture stabilization control and tracking control of the near space aircraft in the presence of actuator and control surface faults, and flight posture robust fault-tolerant control is realized.

The invention relates to a dynamic control reallocation based unmanned aerial vehicle attitude fault-tolerance control method. According to the method, dynamic inversion control is adopted, an attitude angle rate loop dynamic equation corresponding to a near-space vehicle is given firstly, further a control surface stuck and damage model is established, then an adaptive sliding mode observer is designed according to the control surface stuck and damage model of an attitude angle rate loop, the attitude angle rate loop and a designed observer dynamic equation are simultaneous, an angle loop controller and an angle rate loop controller are designed, a designed fault-tolerance control system does not need accurate damage information, interference information and the like which are all implied in the designed adaptive sliding mode observer, the implicit information is fed back to a pseudocontrol control law and an online reallocation algorithm LMI in real time so as to realize robust fault-tolerance control, the designed method is applied to attitude stabilization control and tracking control of the near-space vehicle under the condition of control surface faults, and flight attitude robust fault-tolerance control is realized.

The utility model relates to an active control test platform and a method for vibration of a near space aircraft model. The test platform leads an aircraft model distributed and stuck with a piezoelectric sensor and an actuator to be hung on an aluminum alloy outer framework, and is connected with a function signal generator of a power amplifier to drive an exciter fixed on the aluminum alloy outer framework and lead the aircraft model connected with the exciter to generate a vibration response; and a vibration response signal is collected by the piezoelectric sensor, and is outputted on the power amplifier and acted on the piezoelectric actuator through the operation of a computer inserted with a data acquisition and output card. The active control test platform has the characteristics of convenient composition, simple structure, and good expandability. The method can provide experimental verification realization means for the active control method for the vibration of the near space aircraft model, and provide technical realization supports for exploring the further practical applications of relevant control theories and methods.

The invention relates to the field of air vehicle guide and control system design, and provides a closed loop analytic real-time guide method; the control accuracy is high, and the speediness is good; the cruise-section guide problem of near space air vehicles can be solved under the conditions of uncertain parameters and disturbance influence; the robustness is strong. The invention adopts the technical scheme that the attitude sequence resolving-based air vehicle novel real-time guide method comprises the following steps: 1) according to the six-degrees-of-freedom equation of a near space air vehicle, providing a translational motion model, namely a translational motion equation set, and a rotational motion model, namely a rotational motion equation set; 2) simplifying the provided translational motion equation; 3) obtaining the closed loop feedback relation between the attitude angle instruction and the track, and providing an executable instruction for the control system. The invention is mainly applied to air vehicle guide and control system design.

The invention provides a robust adaptive trajectory tracking control strategy designed for a six-freedom-degree twelve-state model of an NSV (Near Space Vehicle) at a cruising flight stage. The method comprises the steps: firstly proposing a brand-new dynamic model approximation method which is applied to the design of a flight path controller; secondly designing an interference estimator independent of the controller through employing the adaptive technology; and thirdly employing dynamic inverse to be combined with a backstepping method, and respectively giving design methods for position, attitude angle and angular rate controllers, wherein an instruction filter is used for avoiding differential expansion in the backstepping design, and a tracking error caused by that a project approval instruction cannot be completely executed is corrected through a compensation item, so as to construct a robust item to inhibit the impact on the trajectory tracking from an interference estimation error. According to the embodiment of the invention, the above method is described in detail.

The invention discloses a near space vehicle (NSV) preset performance attitude tracking control method with input nonlinearity. The method comprises the following steps: (1) enabling an input nonlinearity link including a dead zone and input saturation be equivalent to an input saturation link by means of an added dead zone right inverse function, and adding the input nonlinearity to a NSV attitude model; (2) introducing preset performance limit of tracking errors, adjusting the transient performance and the steady-state error of a tracking error signal at the same time, and converting the tracking error signal constrained by preset performance into a unconstrained converted error signal by means of an error conversion function; and (3) constructing an auxiliary system and designing an adaptive neural network attitude tracking controller by means of a backstepping method. The method enables the NSV attitude model to be capable of tracking a desired attitude angle signal with parameter uncertainty, external disturbances, and input nonlinearity.

The invention discloses an MPPT control circuit based on an improved H-bridge DC-DC topology, and belongs to the technical field of control of a solar unmanned aerial vehicle power supply system DC-DC circuit. The MPPT control circuit comprises a solar array working voltage sampling circuit, a solar array working current sampling circuit, a bus voltage sampling module, a bus current sampling module, an overvoltage protection module, a logic module, an MPPT mode driving module and a direct-through mode driving module. According to the MPPT control circuit, the working mode of the DC-DC circuit is automatically judged and selected, and two adjusting modes of Buck and Boost under the MPPT mode are enabled to be smoothly switched effectively through transition of the direct-through mode so as to guarantee the stability of the system. Unified control of the MPPT mode and the direct-through mode can be realized, the maximum output power point of a solar cell array can be tracked in real time, the utilization rate of the solar cell array can be enhanced and the power supply output capacity of an unmanned aerial vehicle power supply system can be enhanced so as to provide reference and help for energy supply of near space aircrafts represented by the solar unmanned aerial vehicle.

The invention relates to a non-linear self-adaptive flight control method belonging to the technical field of near-space aircraft flight control. According to the invention, with regard to problems of indeterminate parameters in vertical control of near-space aircrafts, and with consideration given to advantages of sliding-mode control over non-linear system control and the feature that interval second-type fuzzy control with nondeterminacy fuzzy sets is especially suitable for solving problems of nondeterminacy, a second-type fuzzy control system is combined with sliding-mode control and self-adaptive control to design a flight controller to enable the flight controller to have better control performance. The stability and robustness of the near-space aircrafts are raised.

The invention discloses a convection heat-radiation system for temperature control of electronic equipment close to space vehicle, comprising an electronic equipment pressure maintaining cabin, a liquid storage tank, a pump, an in-cabin finned-pipe heat exchanger, a fan for an in-cabin radiator, an extravehicular finned-pipe heat exchanger, a fan for an extravehicular radiator, a pipeline and circulating fluid, the inlet of the pump is connected with the outlet of the liquid storage tank, the inlet of the liquid storage tank is connected with the outlet of the in-cabin finned-pipe heat exchanger, the inlet of the in-cabin finned-pipe heat exchanger is connected with the outlet of the extravehicular finned-pipe heat exchanger, the inlet of the extravehicular finned-pipe heat exchanger is connected with the outlet of the pump, and the circulating fluid working medium is filled in the pipeline; in the invention, a fluid loop coupling type convection radiator is arranged, waste heat in the electronic equipment is transmitted to the finned-pipe heat exchanger through the fluid loop and is discharged to the ambient atmosphere through the convection radiator driven by the motor; the convection radiator is used in special working environment at bottom part of a stratosphere to carry out heat radiation of high-power electronic equipment, and the structure is compact.

The invention relates to a proximity space multi-agreement label exchange network system. The system includes a consumer layer and a proximity space network layer. The proximity space network layer consists of access nodes and transmitting nodes. A router on any access node or transmitting node includes a network edge router module and a label exchange router module. Both the network edge router module and the label exchange router module include a control plane and a transmitting plane. The control plane includes an air control unit and a ground control unit. The air control unit and the ground control unit alternate the control information through an air-ground signaling interface. The system adopts the mode of air-ground integrated cooperative processing which can better solve the contradiction between limited resources and capabilities of a proximity space aircraft node and the requirements of high performances and a plurality of functions. Besides, the invention also relates to a processing method of the proximity space multi-agreement label exchange network system.

The invention relates to a low-density carbon-tile surface oxidation-resistant coating capable of resisting the temperature of 1700 DEG C and a preparation method thereof, and belongs to the technical field of heat protective materials. For the requirement of heat protective systems of novel near-space aerocrafts and reusable launch vehicles for the heat protective materials with integrated functions of high-temperature resistance, light weight and heat insulation, the invention provides a low-density carbon-tile surface oxidation-resistant coating system capable of resisting the temperature of 1700 DEG C and a preparation method thereof. The coating is a composite coating with a gradient transition structure. The composite coating comprises an oxidation-resistant layer and a surface high-temperature-resistant oxidation-resistant layer, wherein the oxidation-resistant layer is formed by doping oxidation-resistant components such as Si, B, C and O and the like in a porous carbon-fiber skeleton; and the surface high-temperature-resistant oxidation-resistant layer is composed of compact M(M is Hf, Zr)B2-MoSi2 ultrahigh-temperature ceramic components.

The invention relates to near space aerocrafts, in particular to a method for designing a double-air-inlet hypersonic-velocity wave-rider aircraft. Aerodynamic features are firstly designated, and then a design scheme meeting the features is reversely derived. A certain three-dimensional shock wave curved surface in a complex shape is designated, and the change rule of the transverse curvature center of the three-dimensional shock wave curved surface is obtained. A series of basic flow fields needed in wave rider design are reversely derived. Flow line tracking with different curvature centers and different longitudinal positions is carried out in the basic flow field at each circumferential position. Finally, a wave rider device capable of generating the designated complex three-dimensional shock wave curved surface is obtained, and the design scheme of the double-air-inlet hypersonic-velocity wave-rider aircraft is completed. Integration of the aircraft and air inlets is achieved, meanwhile, double-air-inlet design is achieved, and a solving method is provided for the design of a double-engine hypersonic-velocity aircraft.

The invention belongs to the technical field of radar target tracking, and particularly relates to a method for tracking trajectories of a hypersonic velocity near space aircraft. The method comprises the following specific steps of: according to earth universal gravitation, aerodynamic resistance and an aerodynamic lifting force which are applied to the hypersonic velocity near space aircraft in the cruise phase, establishing a motion model of the hypersonic velocity near space aircraft in the cruise phase; according to earth universal gravitation, aerodynamic resistance, an aerodynamic lifting force and an engine controlling force which are applied to the hypersonic velocity near space aircraft in the reentry phase, establishing a motion model of the hypersonic velocity near space aircraft in the reentry phase; according to echoed data of the hypersonic velocity near space aircraft in the cruise phase and in the reentry phase, obtaining the trajectories of the hypersonic velocity near space aircraft in the cruise phase and in the reentry phase; removing trace points which do not accord with the motion models of the hypersonic velocity near space aircraft from the trajectories of the hypersonic velocity near space aircraft in the cruise phase and in the reentry phase.

The invention discloses a back-stepping robust self-adaptive dynamic surface-based near-space aircraft control system. The system comprises the steps of: initially employing interpolation fitting to obtain an aerodynamic parameter of an aircraft cruising segment, and building a longitudinal uncertainty strict feedback block non-linear model of an aircraft; secondly, separately controlling speed and height according to state variable characteristic of the aircraft, sequentially solving a control signal by a back-stepping method, approaching unknown interference by employing a RBF neural network, and updating the parameter on line and in real time; and finally, adding a dynamic surface control method with regard to the problems of virtual control signal derivation difficulty and differentialexpansion. By the system, the problems of virtual control signal derivation difficulty and differential expansion are solved, and the system has favorable tracking property and robustness.