131 results about "Ground testing" patented technology

The invention discloses a lunar dust protection system for a low-energy particle detector. The lunar dust protection system comprises a protective cover, a nitrogen blowing opening, a pressing device, an unfolding device, a power supply subsystem and a control subsystem, wherein the protective cover and the nitrogen blowing opening are integrally designed; the protective cover covers the head of the sensor and is consistent with the head of the sensor in shape; the nitrogen blowing port blows nitrogen into the sensor in a ground test stage to prevent pollution; the blocking cap is used for blocking before the satellite is launched to prevent dust pollution; when the system is in a protection state, the pressing device presses the protective cover and the head of the sensor, and the protective cover is released through thermal cutting according to an instruction of the control subsystem; the unfolding device is matched with the pressing device to achieve unfolding of the protective cover, and the protective cover is locked after being released to a designated position. The power supply subsystem supplies power to the pressing device according to an instruction of the control subsystem; the control subsystem periodically collects state monitoring parameters and transmits the state monitoring parameters to the load management unit, controls power-on and power-off of the pressing device according to a ground remote control instruction, and executes unfolding operation of the system.

The invention discloses an aircraft high-altitude operation data calibration method based on a ground test, and the method comprises the following steps: S1, simulating all altitude environments, and selecting different working conditions; s2, calibrating operation data under different altitude environments and different working conditions; wherein in the step S2, the operation data comprises the fuel injection quantity, when the fuel injection quantity is calibrated, the working condition is selected, the optimal fuel injection quantity is set as an initial value under the 0-altitude environment, then adjustment is conducted on the basis of the initial value, and the fuel injection quantity under the same working condition at all the altitudes is determined according to the change condition of the aircraft power. The calibration process is relatively simple, the adjustment basis is reliable, the adjustment efficiency is high, the result is reliable, and the high-altitude flight reliability of the aircraft can be ensured.

The invention provides a drilling fluid flow monitoring device and method suitable for a mud runway, relates to the field of petroleum drilling ground testing instruments, and solves the problem that an existing monitoring method is low in overflow prediction precision and cannot find overflow in time. The device comprises a flow measuring unit (1), and the left side and the right side of the flow measuring unit (1) are connected with a flow guide unit (2) and a flow blocking unit (4) through flange plates respectively. The upper part of the diversion unit (2) is connected with the video monitoring unit (3); the flow measuring unit (1) is an electromagnetic flowmeter and is used for measuring the outlet displacement; the flow guide unit (2) is used for guiding drilling fluid to flow to the flow meter; and the flow choking unit (4) is used for controlling stable flow of the drilling fluid. According to the drilling fluid flow monitoring device suitable for the mud runway, real-time accurate monitoring of the flow of a drilling fluid outlet can be achieved, then monitoring of overflow and leakage of the drilling fluid is achieved, and well control safety is ensured.

The application discloses a hydrogen fuel cell power supply adaptability ground test method, system, equipment, medium and product of a UAV, relates to the UAV test technical field, obtains the ground test data of the to-be-tested UAV, carries out power demand analysis by using the ground test data, obtains a plurality of motor input total power, a plurality of motor output shaft power and weighted average propeller force efficiency, carries out power supply test on the preset hydrogen fuel cell by using each motor input total power, obtains a safety detection result and a plurality of variable load power response times, when the safety detection result is qualified, then carries out power supply adaptability detection according to each variable load power response time, each motor input total power, each motor output shaft power and weighted average propeller force efficiency, and obtains a power supply adaptability detection result. The application solves the problem that the existing hydrogen fuel cell lacks the safety buffer and power regulation capability of lithium batteries, which easily leads to the risk of stack overheating and reduces the reliability of the hydrogen fuel cell in the application of the UAV.

The application provides a kind of aircraft servo elastic dynamics model correction method considering operating surface gap, first aircraft structure is divided into substructure, and the aircraft dynamics model containing aileron gap nonlinearity is established according to substructure branch modal;And additional inertia force caused by aircraft operating aileron deflection is introduced in dynamics model, and the aircraft open-loop servo elastic motion equation containing aileron gap nonlinearity under independent generalized coordinate is obtained;According to sensor output equation, rudder transfer function and flight control system equation, the aircraft nonlinear servo elastic closed-loop system model containing aileron gap nonlinearity is established, and the model is corrected based on ground vibration test, rudder frequency response test and gap measurement test result, and the stability of the model is judged based on simulation analysis result and open-loop frequency response test result and is further corrected, which improves the accuracy of gap nonlinearity aircraft aerodynamic servo elastic characteristic evaluation, and is helpful to guide aircraft servo elastic ground test engineering practice.

This invention provides a satellite-launch separation signal system and a satellite-launch separation system, comprising: a first user unit that outputs two satellite-launch separation signal sources, each leading to two different contacts of a first analog separation relay; these signals then return to the first user unit via a first separation connector and a second separation connector; the input and output terminals of the first analog separation relay are each connected to a satellite display socket via cables; and a second user unit that outputs two satellite-launch separation signal sources, each leading to two different contacts of a second analog separation relay; these signals then return to the second user unit via a first separation connector and a second separation connector; the input and output terminals of the second analog separation relay are each connected to a satellite display socket via cables. This invention ensures that the signal is provided by the analog satellite-launch separation relay during ground testing, and avoids fatal malfunctions such as accidental separation commands issued before launch or mis-separation due to severe vibrations during the active phase, thus improving the reliability of the satellite-launch separation signal link.

The invention relates to the technical field of space robot ground testing, in particular to a space crawling robot surface and microgravity simulation system. Comprising a surface simulation assembly, an inclined plane adjusting assembly and a gravity balancing assembly, the surface simulation assembly is arranged on the inclined plane adjusting assembly, and the inclined plane adjusting assembly is used for adjusting the inclination angle of the surface simulation assembly; the gravity balancing assembly is arranged above the surface layer simulation assembly and used for being connected with the space crawling robot to be tested, and when the space crawling robot moves on the surface layer simulation assembly according to a set trajectory plan, the gravity balancing assembly can passively adapt to the motion trajectory of the space crawling robot; and the climbing motion of the simulated satellite surface in the space microgravity environment is completed. The space crawling robot simulates crawling on the surfaces of satellites and spacecrafts in the microgravity state, verification of the motion performance of the crawling robot under the space microgravity in a ground test is completed, and the problem that the ground test of the space crawling robot is difficult is solved.

The invention discloses a device for ground testing of a satellite two-dimensional radar mechanism. The device comprises a housing frame; a first driving motor and a second driving motor; a first speed reducer and a second speed reducer; the spherical parallel mechanism comprises a first transmission assembly, a second transmission assembly and a movable platform, the first transmission assembly is connected with the first speed reducer and the movable platform, and the second transmission assembly is connected with the second speed reducer and the movable platform; the two ends of the first pressure sensor and the second pressure sensor are connected with the movable platform and the output support respectively, and the first pressure sensor and the second pressure sensor are perpendicular to each other in the normal direction on the mounting surface of the movable platform; the information processor is connected with the first pressure sensor, the second pressure sensor, the first driving motor and the second driving motor; the radar mechanism driver is connected with the information processor and the radar mechanism; and the upper computer is connected with the radar mechanism driver and the information processor. The synchronous force increasing device has the advantage that synchronous force increasing of the radar mechanism can be achieved.

A multifunctional integrated ground testing system for spaceborne equipment is disclosed, covering the multifunctional testing field of spaceborne equipment. It provides reliable support for the functional performance testing of spaceborne equipment. The system includes an integrated management module, a telemetry and data processing unit, an interface forwarding simulator, a Ka-band converter, an RF link network, an interface forwarding network, and power supply equipment. Adopting a modular design, it features data interaction between the various components of the integrated ground testing system, human-machine interaction, remote control command parsing and distribution, telemetry information parsing, display and storage, health information monitoring, and dynamic signal simulation parameter calculation and distribution. This system solves major problems such as the limited number of connected spaceborne equipment and the single nature of functional performance test items, providing reliable and stable technical support for the ground testing of spaceborne equipment.

The application provides a novel satellite-rocket electrical interface system, which comprises an interface assembly arranged on the outer surface of a satellite cabin plate, the interface assembly comprising X1F connectors, X2F connectors and X10B connectors arranged side by side; a ground test cable assembly, one end of which is provided with a first connector for connecting a satellite ground test device, and the other end of which is provided with a second connector for being connected with the X1F connectors and the X2F connectors; wherein a working port is arranged on the rocket fairing at a position corresponding to the interface assembly, and the working port is configured to allow the ground test cable assembly to be connected with the satellite connectors of the interface assembly before launch or to be manually disconnected and withdrawn. Thus, the application simplifies the satellite-rocket interface relationship, improves the reliability of satellite-rocket separation, improves the launch capacity, and reduces the system development cost.

This invention relates to the field of spacecraft attitude control technology, and more particularly to an ultra-precise and ultra-stable on-orbit adaptive control method and system for pointing under additional stiffness disturbances. This on-orbit adaptive control method, by applying specific excitations to actuators and measuring their responses, combined with a recursive filtering algorithm, can accurately identify the total actuator stiffness, including the additional stiffness of the cable, in real-time on-orbit, overcoming problems caused by differences in ground testing environments and time-varying stiffness. Based on the real-time identified stiffness values, the parameters of the axial compensation controller and the actuator compensation controller are dynamically adjusted, enabling the control system to adapt to changes in the additional stiffness of the cable and maintain excellent control performance at all times. By employing dual compensation control at the axial load level and the actuator level, additional stiffness disturbances are synergistically suppressed at both the overall system and local actuator levels, improving the comprehensiveness and accuracy of control, and achieving ultra-precise and ultra-stable on-orbit adaptive control for pointing under cable additional stiffness disturbances.

The present application relates to the technical field of ground test of space liquid engine, and particularly relates to a large-thrust inclined test bench based on flexible leaf spring support and an in-situ calibration method. The test bench comprises a fixed frame platform arranged obliquely and a movable frame supported in the fixed frame gantry through multiple sets of leaf spring assemblies; the leaf spring assemblies constrain the movable frame to move only along the thrust axis direction; the front end of the movable frame is connected to a force measuring assembly through a transition frame, and the rear end is installed with a liquid engine through a adapter frame; the fixed frame is provided with an in-situ calibration assembly and a pre-tightening force assembly acting on the movable frame, and a safety limiting assembly. The in-situ calibration method applies a standard force to the movable frame through the calibration assembly, and simultaneously records the output of the force measuring assembly to establish a calibration relationship. The present application optimizes the super-large thrust bearing path through the oblique leaf spring support structure, has single-degree-of-freedom guiding and high lateral stiffness, and realizes high-precision, high-dynamic and high-reliability measurement of the thrust of the hundred-ton liquid rocket engine under real inclined working conditions in combination with in-situ hydraulic calibration.

The invention discloses an automatic test system and an automatic test sequence framework for a spacecraft. The system is characterized in that an automatic test sequence framework used for describing the automatic test of the spacecraft is defined, and the automatic test sequence framework is used for realizing the automatic test of the spacecraft through the following modules, the module comprises a sequence configuration sub-module, an automatic sequence execution sub-module, a result monitoring sub-module and a device-ground interface sub-module; and the sequence configuration sub-module is used for configuring and describing test steps of a test item by adopting the automatic test sequence framework to form an automatically executed test sequence. According to the invention, the ground test system for the automatic test of the spacecraft is realized by designing the automatic test sequence framework of the spacecraft, so that the test sequence is automatically executed, the data is automatically interpreted, the probability of errors introduced by human factors is reduced, and the overall test efficiency is greatly improved.

The invention discloses a high-precision gravity unloading device capable of achieving three-dimensional motion, and belongs to the field of gravity unloading. The device comprises an unloading truss assembly, a first longitudinal guide rail assembly, a second longitudinal guide rail assembly and two independent moving units, wherein the first longitudinal guide rail assembly and the second longitudinal guide rail assembly are different in installation height. Each set of motion unit comprises a longitudinal tackle assembly slidably connected to the longitudinal guide rail assembly, a transverse guide rail assembly bridged between the longitudinal guide rail assemblies, a transverse tackle assembly slidably mounted on the transverse guide rail assembly, a fixed pulley assembly fixedly connected to the transverse tackle assembly, and a lifting rope crossing the fixed pulley assembly to connect the tested rotating mechanism and the counterweight. Two sets of independent movement units are arranged, and the structural interference of a plurality of unloading points in the longitudinal opposite staggered movement process is avoided through the arrangement of longitudinal guide rails which are arranged in a high-low staggered mode. By optimizing the layout of the fixed pulley and the counterweight, three-dimensional high-precision dynamic follow-up unloading of a complex double-shaft rotating mechanism in a large rotating angle range is realized, and reliable microgravity environment simulation is provided for ground testing of aerospace equipment.

This utility model relates to a mobile testing system for ground testing of a turboshaft engine, belonging to the field of turboshaft engine technology, and solves the problem of the lack of mobility in existing whole-engine testing technologies. This utility model includes an experimental platform unit, a measurement and control unit, and a water circulation unit; the experimental platform unit, measurement and control unit, and water circulation unit are mounted on a mobile device; the experimental platform unit includes an engine testing platform and a dynamometer; the measurement and control unit includes a measurement and control device and a data acquisition device; the water circulation unit includes a circulating water cooling device, a cold water tank, and a hot water tank. This utility model enables ground testing and rapid transfer of a turboshaft engine.

The application provides an offshore wind turbine evaluation system and method based on ground testing and virtual testing, relates to the technical field of wind power generation, and comprises a virtual testing system and a ground testing platform.The virtual testing system is configured to simulate offshore environmental conditions and grid-connected operation conditions of a wind turbine to be tested, and to generate a control signal of the ground testing platform.The ground testing platform is configured to apply the control signal to the wind turbine to be tested, collect operation state data and grid interaction data in the process of application, and return the data to the virtual testing system.The virtual testing system is further configured to evaluate the operation performance and grid-connected performance of the wind turbine to be tested.The application combines the ground testing platform and virtual testing technology, realizes accurate monitoring and evaluation of the performance of the wind turbine, provides important data support for operation and maintenance management of a wind farm, and promotes the sustainable and healthy development of the offshore wind power industry.

The application discloses a bellows elastic load equivalent device, and relates to the technical field of aerospace. The bellows elastic load equivalent device is used for simulating the real state of a bellows of an aircraft during flight, and achieving the problem of consistent bending stiffness of a bellows simulation part during ground testing. The bellows elastic load equivalent device comprises a bellows simulation part, the bellows simulation part has two opposite end faces, one of the two end faces is used for being connected with a frame of an engine, and the other end face is used for being connected with a nozzle of the engine. The bellows elastic load equivalent device further comprises a swing mechanism and two fixed connecting pieces. The bellows simulation part has a spiral groove which spirally extends along the distribution direction of the two end faces. The bellows elastic load equivalent device further comprises two mounting structures. The bellows elastic load equivalent device is used in aerospace experiments.

The invention provides a novel satellite-rocket electrical interface system, which comprises an interface assembly arranged on the outer surface of a satellite deck, and is characterized in that the interface assembly comprises an X1F connector, an X2F connector and an X10B connector which are arranged side by side; one end of the ground test cable assembly is provided with a first connector used for being connected with satellite ground test equipment, and the other end of the ground test cable assembly is provided with a second connector used for being in butt joint with the X1F connector and the X2F connector; an operation opening is formed in the position, corresponding to the interface assembly, of the rocket fairing, and the operation opening is configured to allow the ground test cable assembly to penetrate through the operation opening to be connected with a satellite connector of the interface assembly or to be manually disconnected and withdrawn before launching. Therefore, the satellite-rocket interface relation is simplified, the reliability of satellite-rocket separation is improved, the carrying and launching capability is improved, and the system development cost is reduced.

The invention belongs to the field of remote sensing satellite system engineering, and relates to a ground large closed-loop test system and method for a space exploration load working mode. A new space exploration load on-orbit working mode is designed according to the characteristics of an observation target, and the interior of the load is designed into a local sky area mode and a window tracking mode; the whole satellite load observation mode is designed into a natural rendezvous monitoring mode, a fixed-point staring monitoring mode, a scanning imaging mode and a target tracking mode; and aiming at a detection load mode, a dynamic star simulator is adopted to simulate an imaging sky area, and a simulated motion fragment target is added into the dynamic star simulator, so that simulation of a real imaging scene is realized. Meanwhile, in order to ensure that the load can successfully match and verify the local star catalogue and the sky area, a normal window tracking mode is switched to, so that the pointing direction corresponding to the sky area simulated by the dynamic star simulator is consistent with the pointing direction of the control broadcast, the on-orbit real working condition and time sequence can be accurately simulated, and the sufficiency of the ground test is improved.

The invention belongs to the technical field of spacecraft ground test equipment, and particularly relates to a microgravity simulation device for spacecraft component ground test, which comprises a base, an air floatation support mechanism, an ultrasonic standing wave stabilization mechanism, a self-adaptive clamping mechanism, a servo drive error compensation mechanism, a control system and a test platform. The air floatation supporting mechanism provides vertical non-contact supporting force to offset gravity, the ultrasonic stabilizing mechanism maintains horizontal stability, the self-adaptive clamping mechanism is adaptive to parts of multiple sizes, and the servo compensation mechanism corrects deviation in real time. The composite suspension structure is combined with PID-fuzzy closed-loop control, the simulation error is controlled within 4.0%, the defects that existing equipment is low in precision, short in time, prone to pollution and poor in universality are overcome, long-time stable simulation can be achieved, and reliable support is provided for part ground verification.

The invention relates to the technical field of spacecraft ground testing, in particular to a multi-channel smooth ring testing system based on a thermal vacuum environment, which is characterized in that a tested multi-channel smooth ring is arranged in a thermal vacuum tank, and a rotor end of the tested multi-channel smooth ring is connected with a testing end of a thermal vacuum torque sensor through a shifting fork; an insertion loss tester and a thermal vacuum tank cabin-penetrating optical flange are connected with a stator end of a tested multi-channel smooth ring, the tested multi-channel smooth ring is driven to rotate under the driving of a thermal vacuum motor, and real-time rotating speed, rotating torque data and insertion loss of the tested multi-channel smooth ring are collected and sent to an upper computer. And the upper computer performs dynamic curve drawing and display on the received insertion loss, rotation torque and rotation speed data, so that synchronous monitoring of the detected multi-channel smooth ring is realized to facilitate correlation analysis. According to the invention, real-time and uninterrupted performance monitoring of the multi-channel smooth ring in the thermal vacuum test process is realized, and accurate data support is provided for failure analysis.

The application provides a recovery type load measuring system for a carrier rocket, which comprises a collecting unit, a collecting storage unit, a collecting storage unit antenna, a battery, a fiber optic demodulator, a wireless beacon device, a fiber optic strain measuring device, a pressure sensor and an overload sensor; through collecting and storing multiple load parameters in the flight process of the carrier rocket; in the process of returning of a sub-stage, the ground equipment receives short message data sent by the system to obtain the real-time flight position of the sub-stage; after the sub-stage lands, the wireless beacon device is started, the ground equipment identifies the radio signal sent by the wireless beacon device, so that the sub-stage of the rocket is searched and the collecting storage unit installed on the sub-stage is obtained; the data in the collecting storage unit is read and analyzed through the ground test equipment. Therefore, the load bearing condition of the rocket body structure under different attack angles can be directly obtained, and the system has the characteristics of strong independence, high reliability and wide adaptability.

The application discloses an outer vacuum obtaining and magnetic shielding device for ground testing of a hydrogen atomic clock, comprising an outer vacuum preparation mechanism, wherein the outer vacuum preparation mechanism comprises a two-stage vacuum pump set, the two-stage vacuum pump set comprises a rotary vane molecular pump and a diaphragm oil-free mechanical pump, the two-stage vacuum pump set further comprises a vacuum gauge, a magnetic shielding cover, a pipeline, a vacuum valve one, a vacuum valve two and a magnetometer; the molecular pump and the mechanical pump are connected in series, and the mechanical pump serves as a front-stage pump. The application integrates zero magnetic treatment and environment control functions on the basis of a traditional vacuum pump set, is suitable for spacecraft assembly testing environment, meets severe testing conditions of the hydrogen atomic clock, is convenient to maintain, is suitable for transportation, can realize real-time measurement and control of vacuum degree and magnetic intensity, can realize rapid pressure relief testing of the hydrogen atomic clock in the device, verifies performance indexes of the hydrogen atomic clock under extreme environment, and is a necessary device for ground testing of next-generation spaceborne precise time and frequency instruments.

The invention relates to the technical field of aircraft ground test equipment, in particular to an undercarriage fake part structure. The undercarriage false part structure mainly comprises a false part main body, an upper sleeve, an upper sleeve shaft rod, a fixing pin for fixing the upper sleeve shaft rod, a lower sleeve, a side supporting rod, a side supporting rod adapter, a connecting structure, a lower sleeve shaft rod, a vertical loading joint, a lateral loading joint, a fixing structure, a gasket and a course loading joint. The device can independently or compositely apply vertical, lateral and course forces, can simulate'static load ', 'dynamic load' and'composite load 'of a mechanical part, and is matched with fixed positioning of a sleeve shaft rod on the top and force input of a loading connector at the bottom to form a'reference-loading' closed loop, so that the loading force can accurately act on a target area of a false part main body, and the false part main body can accurately act on the target area of the false part main body. And test data errors caused by positioning deviation are reduced, so that the undercarriage false part structure is used for replacing a real part structure for test load application and transmission.

The invention provides a ground testing method and system for satellite-borne equipment, an upper computer and a lower computer, and can be applied to the technical field of equipment testing. The method comprises the steps that a script interpreter is used for converting a control script into a control instruction sequence, the control script is a script used for testing satellite-borne equipment on the ground, the control instruction sequence comprises a plurality of control instructions arranged in sequence, the control instructions comprise keywords, or the control instructions comprise keywords and parameters, and the keywords are used for indicating event types; acquiring a current control instruction from a plurality of control instructions included in the control instruction sequence by using an instruction executor, and executing the current control instruction according to an event type indicated by a keyword included in the current control instruction; and sending converted data which is obtained by executing the current control instruction and corresponds to the parameter to a lower computer by using an instruction executor in response to the event type of a write-in event or a timing event, so that the lower computer controls the satellite-borne equipment to execute an action indicated by the converted data.

The invention provides a rotary jet flow interference wind tunnel test system and a test method, and aims to solve the problem of unsteady aerodynamic characteristic research of aircrafts such as rotary missiles and the like under jet flow interference conditions. The system comprises a rotation driving and supporting subsystem, a high-precision force measurement and control subsystem and a rotation jet flow generation system, and active controllable rotation of a test model in a wind tunnel and real-time synchronous accurate measurement of six-component aerodynamic force / torque in a rotation state are realized. And high-pressure jet flow is stably supplied through the hollow main shaft and the rotary sealing joint. According to the test method, on the basis of a simulation criterion that the pressure ratio, the momentum ratio and the Slowhaal number are similar, the rotating speed of the model, the incoming flow Mach number and the jet flow pressure ratio are adjusted, and the system studies the rotation and jet flow coupled unsteady aerodynamic interference rule. According to the invention, a key ground test means is provided for aerodynamic characteristic research and optimization design of a rotary aircraft direct force control technology.

The embodiment of the invention provides a power distribution control method, system and equipment of a carrier rocket, the method, system and equipment are applied to a power distribution control system arranged on the carrier rocket, the power distribution control system controls power distribution of a ground power supply and a rocket battery to the carrier rocket, and the method comprises the steps of receiving a ground test starting instruction; controlling a ground power supply to supply power to a first preset load group of the carrier rocket according to the ground test starting instruction; receiving a power conversion instruction; stopping power supply of the ground power supply according to the power conversion instruction, and controlling the rocket battery to supply power to a second preset load group of the carrier rocket; receiving flight data of the carrier rocket and state data of a battery on the rocket; and according to the flight data of the carrier rocket and the state data of the on-rocket battery, controlling the on-rocket battery to distribute power to each load in a second preset load group of the carrier rocket. According to the embodiment of the invention, the power supply strategy of each load in the second preset load group can be dynamically adjusted according to the actual flight stage and the power supply health condition.

This application relates to the field of satellite cryogenic thermal insulation support technology, specifically to an active adjustable thermal resistance thermal insulation structure for gravity condition switching and its manufacturing method. The structure includes a hot-end interface, a hot-end support assembly, an isolation screw mechanism, a locking module, a heater, a fixing structure, a ground-side support pipe, an on-orbit support pipe, a cold-end support assembly, and a cold-end interface. The active on-orbit non-connected thermal insulation support technology employed in this application solves the problems of small separation gaps and poor repeatability associated with passive on-orbit non-connected thermal insulation support technologies. It achieves effective active separation and large displacement isolation between the locking module and the fixing structure, ensuring separation and isolation are unaffected by residual microgravity disturbances, improving safety margins, allowing for multiple ground tests and verifications, and enhancing mission confidence.

This application provides a method, apparatus, device, storage medium, and program product for measuring fuselage parameters, belonging to the field of aviation technology. The method includes: performing a flight test to obtain target flight test data for a target area of ​​the aircraft, the flight test data including the aircraft's vertical acceleration, fuselage angular acceleration about the pitch axis, longitudinal strain variation, and circumferential strain variation; performing a ground test to obtain target ground test data for the target area, the ground test data including the coefficients of pressure circumferential stress versus pressure, pressure longitudinal stress versus pressure, and bending longitudinal stress versus bending moment; and determining local fuselage physical parameters for the target area based on the target flight test data and the target ground test data, the local fuselage physical parameters including local torque and moment of inertia. This application can reduce the cost of obtaining local fuselage physical parameters.

This invention provides a simulation method for test equivalent signals based on the Euler discrete model. It primarily solves the problem of equivalent generation of gyroscope angle signals in test equivalent converters, enabling high-precision reproduction of the complex dynamic angle changes of an aircraft during flight. This overcomes the difficulty of accurately simulating gyroscope angle signals during ground testing. First, using predetermined aircraft pitch channel altitude trajectory data, a fast differentiator is used to solve for the differential data of the altitude and horizontal trajectories, thereby solving for the aircraft pitch angle and pitch velocity data. Then, the Euler method is used to establish a discrete measurement model of the gyroscope, solving for the measured value of the pitch velocity. Considering the influence of error drift, integration is performed to obtain the gyroscope's measurement signal of the pitch angle. This outputs a high-precision result of the gyroscope angle changes during high-altitude flight. This method has the advantages of high simulation accuracy and good agreement with reality.