A propellant slosh characteristics measurement flight test platform

Abstract

The application provides a propellant sloshing characteristic measurement flight test platform, which comprises a test aircraft and a test control system; the test control system comprises an aircraft control system, a data measurement system and a test measurement and control system; the aircraft control system and the data measurement system are installed on the test aircraft; the test measurement and control system is arranged on the ground; the test aircraft comprises a test tank, and the test tank is filled with a measured propellant liquid. The application uses the same upper tank configuration as a vertical take-off and landing liquid carrier rocket, simulates the propellant sloshing environment in the reentry and return flight process of the rocket, and measures the force, heat, optical and other parameters of the propellant.

Description

Technical Field

[0001] This application relates to the field of flight test platform technology, and in particular to a flight test platform for measuring propellant sloshing characteristics. Background Technology

[0002] Currently, liquid-fueled launch vehicles use different tanks to hold the required fuel (such as kerosene, methane, or liquid hydrogen) and oxidizer (such as liquid oxygen), both collectively referred to as liquid propellant. As the rocket flies, the propellant in the tanks is continuously consumed, and the cavity gradually increases. On the one hand, the violent shaking of the propellant will cause mixing, dissolution, and temperature exchange with the gas in the tank, resulting in changes in tank pressure. To ensure stable combustion of the rocket engine, the propellant needs to be kept at a certain pressure covering the propellant delivery port before engine ignition and during operation to prevent gas from mixing into the delivery pipeline. On the other hand, the shaking of the propellant will create a disturbing torque on the rocket's motion. If the shaking frequency is close to the rocket's flight and control frequency, it may cause resonance, further aggravating the shaking amplitude, and even affecting the rocket's flight attitude stability and engine operating stability.

[0003] In recent years, with SpaceX's Falcon 9 rocket adopting a reusable single-stage design and deploying low-Earth orbit internet satellites on a large scale, my country has also conducted key technology research on powered vertical recovery of rockets. In this flight mode, after completing the ascent phase, the return journey involves multiple flight phases, including attitude adjustment, recovery, high-altitude gliding, powered deceleration, atmospheric deceleration, and vertical landing. The wide flight airspace and velocity range, multiple engine start-ups and shutdowns, and complex and variable flight environment, along with the diverse flight profiles, all contribute to the complexity of propellant sloshing. Therefore, measuring propellant sloshing characteristics during the design of liquid-fueled launch vehicles for system analysis, design, and verification is essential for ensuring normal rocket flight.

[0004] Existing methods use servo motion platforms to measure propellant sloshing characteristics. While servo motion platforms have advantages such as fast operation, high measurement accuracy, and convenient observation, they also have limitations such as short motion stroke and limited test time. Drop tower microgravity tests have the advantages of long longitudinal stroke and traction control, and do not require complex power devices, but the test equipment is large in scale and the lateral motion space is limited.

[0005] Traditional sounding rockets can only provide a flight environment for the powered ascent phase and a limited freefall state. Fixed-wing aircraft need to generate lift during forward flight. Helicopters and rotorcraft generally need to hoist the mission payload under the fuselage, and the lateral control of the aircraft often requires large attitude adjustments. All of the above flight platforms are difficult to control to provide a low-G flight state and cannot simulate the sinking conditions required for propellant management in liquid launch vehicle tanks.

[0006] Therefore, the urgent technical problem to be solved is: how to provide a flight test platform for measuring propellant sloshing characteristics, using the same top-mounted tank configuration as vertical takeoff and landing liquid launch vehicles, to simulate the propellant sloshing environment during rocket reentry and return flight, and to measure the force, thermal, and optical parameters related to propellant management. Summary of the Invention

[0007] The purpose of this application is to provide a flight test platform for measuring propellant sloshing characteristics, which uses the same top-mounted tank configuration as vertical takeoff and landing liquid launch vehicles to simulate the propellant sloshing environment during rocket reentry and return flight, and to measure the force, thermal, optical and other parameters of the propellant.

[0008] To achieve the above objectives, this application provides a flight test platform for measuring propellant sloshing characteristics. The platform includes: a test aircraft and a test control system; the test control system includes a aircraft control system, a data measurement system, and a test measurement and control system; the aircraft control system and the data measurement system are installed on the test aircraft; the test measurement and control system is located on the ground; the test aircraft includes a test tank containing the propellant liquid to be tested.

[0009] The propellant sloshing characteristics measurement flight test platform described above further includes: a main structure, a landing device, a power unit, a power tank, a servo mechanism, and a control cabin. The landing device and the servo mechanism are fixedly connected to the tail end of the main structure. The power unit, the power tank, and the control cabin are all located inside the main structure. The test tank is also located inside the main structure.

[0010] The propellant sloshing characteristics measurement flight test platform described above includes a sensor and a transducer mounted on the test tank, the sensor and the transducer being used to measure the physical parameters of the test tank.

[0011] The propellant sloshing characteristics measurement flight test platform described above, wherein the aircraft control system and the data measurement system are located in the control cabin.

[0012] The propellant sloshing characteristics measurement flight test platform described above includes an aircraft control system comprising: a navigation device, a controller, a power unit, a servo control mechanism, a control battery, a power battery, and an aircraft data transmission radio; the navigation device is used for navigation and positioning; the control battery provides power to the controller; the power battery provides power to the power unit and the servo control mechanism; and the aircraft data transmission platform is used for wireless communication with a ground data transmission platform.

[0013] The propellant sloshing characteristics measurement flight test platform described above, wherein the power unit adopts a liquid rocket engine, a turbojet aero engine, or an electric ducted fan.

[0014] The propellant sloshing characteristics measurement flight test platform described above includes a data measurement system comprising: a sensor, a converter, a data acquisition and editing unit, a telemetry transmitter, a data storage unit, a beacon, and a measurement battery.

[0015] The propellant sloshing characteristics measurement flight test platform described above includes sensors such as a temperature sensor, a pressure sensor, a liquid level sensor, an image sensor, and an overload sensor.

[0016] The propellant sloshing characteristics measurement flight test platform described above includes a test control system comprising: a ground data radio, a telemetry receiver, and a main control computer. The ground data radio is used to establish two-way wireless communication between the test control system and the test aircraft. The telemetry receiver is used to receive telemetry wireless signals collected by the data measurement system. The main control computer is used to send remote control commands to the test aircraft and monitor the flight status of the test aircraft.

[0017] As described above, the propellant sloshing characteristics measurement flight test platform includes a task monitoring module and a data storage module within the main control computer. The task monitoring module provides a human-machine interface, sends remote control commands to the aircraft, and monitors the flight status. The data storage module stores and manages monitoring data and telemetry data.

[0018] The beneficial effects achieved by this application are as follows:

[0019] (1) Compared with traditional motion simulation methods, this application can overcome the limitations of short motion stroke, limited test time and limited motion space, and can simulate the bottoming conditions required for the management of propellant in liquid launch vehicle tanks under controlled low overload flight conditions.

[0020] (2) This application uses the same top-mounted tank configuration as vertical take-off and landing liquid launch vehicles to simulate the propellant sloshing environment during rocket reentry and return flight, and measures the force, heat, optical and other parameters of the propellant.

[0021] (3) The flight test platform of this application adopts different power configurations such as liquid rocket engine, turbojet aero engine and electric ducted fan. The flight test platform has multiple configurations of vertical take-off and landing aircraft to adapt to different levels of mission load, and can meet the test requirements of different scales and test parameters. Attached Figure Description

[0022] To more clearly illustrate the technical solutions in the embodiments of this application or the prior art, the drawings used in the description of the embodiments or the prior art will be briefly introduced below. Obviously, the drawings described below are only some embodiments recorded in this application. For those skilled in the art, other drawings can be obtained based on these drawings.

[0023] Figure 1 This is a schematic diagram of the structure of the experimental aircraft according to an embodiment of this application.

[0024] Figure 2 This is a schematic diagram of the structure of the test control system according to an embodiment of this application.

[0025] Reference numerals: 10-Main structure; 20-Test tank; 30-Control cabin; 40-Power tank; 50-Power plant; 60-Landing device; 70-Servo mechanism; 100-Experimental aircraft; 200-Experimental control system; 201-Aircraft control system; 202-Data measurement system; 203-Experimental measurement and control system. Detailed Implementation

[0026] The technical solutions of the embodiments of this application will be clearly and completely described below with reference to the accompanying drawings. Obviously, the described embodiments are only some, not all, of the embodiments of this application. All other embodiments obtained by those skilled in the art based on the embodiments of this application without creative effort are within the scope of protection of this application.

[0027] like Figure 1 and Figure 2As shown, this application provides a flight test platform for measuring propellant sloshing characteristics. The platform includes a test aircraft 100 and a test control system 200. The test aircraft 100 is a reusable vertical takeoff and landing (VTOL) aircraft. The test control system 200 includes a aircraft control system 201, a data measurement system 202, and a test measurement and control system 203. The aircraft control system 201 and data measurement system 202 are mounted on the test aircraft 100. The test measurement and control system 203 is located on the ground and operated by test personnel. The test aircraft 100 includes a test tank 20 containing the propellant liquid under test. The test control system 200 is used to perform sloshing characteristic measurement tests on the propellant liquid under test. Compared with traditional motion simulation methods, this invention overcomes the limitations of short motion stroke, limited test time, and limited motion space, and can achieve controlled low-overload flight conditions, simulating the sinking conditions required for propellant management in liquid launch vehicle tanks.

[0028] As a specific embodiment of the present invention, the data measurement system 202 and the aircraft control system 201 are made independent of each other to simplify the adaptive design for different test mission requirements and ensure the reliability of independent acquisition of test data.

[0029] like Figure 1 As shown, the test aircraft 100 also includes: a main structure 10, a landing device 60, a power plant 50, a power tank, a servo mechanism 70, and a control cabin 30. The main structure 10 is used for the connection and support of various components. The landing device 60 and the servo mechanism 70 are fixedly connected to the tail end of the main structure 10. The landing device 60 is used for takeoff and landing support and cushioning of the test aircraft 100. The power plant 50, the power tank, and the control cabin 30 are all located inside the main structure 10. A test tank 20 is also installed inside the main structure 10. When the test aircraft 100 is perpendicular to the ground, the internal structure of the main structure 10, from top to bottom, is as follows: test tank 20, control cabin 30, power tank 40, and power plant 50. The power plant 50 and the servo mechanism 70 are located below the center of mass of the test aircraft 100. The equipment of the aircraft control system 201 and the data measurement system 202 are installed in the control cabin 30 and close to the center of mass of the test aircraft 100.

[0030] As a specific embodiment of the present invention, the test vehicle 100 has an upper-mounted test tank 20 configuration and a thrust vectoring propulsion layout at the end. The test vehicle 100 of the present invention uses the same upper-mounted tank configuration and end-thrust vectoring propulsion layout as the vertical takeoff and landing liquid-fueled launch vehicle, and has the same inverted pendulum control method and flight attitude characteristics, which can simulate the flight mode of axial variable thrust and lateral stabilization control used by the vertical takeoff and landing liquid-fueled launch vehicle during reentry and return flight.

[0031] As a specific embodiment of the present invention, the power tank 40 stores an oxidant or a propellant to provide a power source.

[0032] like Figure 1 As shown, a sensor and transducer are installed on the test tank 20 to measure the physical parameters of the test tank 20. The sensor and transducer are installed inside or near the test tank 20 to facilitate the measurement of the parameters of the propellant liquid being tested inside the test tank 20.

[0033] As a specific embodiment of the present invention, the aircraft control system 201 and the data measurement system 202 are located in the control cabin 30 to facilitate the control of the test aircraft 100 and the monitoring of the parameters of the propellant liquid under test in the test tank 20.

[0034] like Figure 2 As shown, the aircraft control system 201 includes: navigation equipment, a controller, a power unit 50, a servo control mechanism, a control battery, a power battery, and an aircraft data transmission radio. The navigation equipment is used for navigation and positioning; the controller is used for data acquisition, navigation, control, refueling, pressurization calculation, power distribution, timing, and command output. The servo control mechanism is used to adjust the thrust direction. The control battery provides power to the controller; the power battery provides power to the power unit 50 and the servo control mechanism. Specifically, the control battery provides primary energy to the controller, the power distribution circuit inside the controller provides converted secondary energy to various peripheral devices, and the power battery provides power to high-power devices such as the power unit 50 and the servo mechanism 70. The aircraft data transmission platform is used for wireless communication with the ground data transmission platform.

[0035] like Figure 2 As shown, the aircraft control system 201 also includes: pressurization and delivery devices for each tank, such as auxiliary valves and gas cylinders for each tank.

[0036] As a specific embodiment of the present invention, the navigation device includes an inertial measurement unit, a satellite navigation receiver, and / or a relative navigation device, such as a ranging radar, an optical flow sensor, and an RTK positioning system.

[0037] As a specific embodiment of the present invention, the power unit 50 adopts a liquid rocket engine, a turbojet aero engine, or an electric ducted fan. The present invention employs different power configurations such as liquid rocket engines, turbojet aero engines, and electric ducted fans, and the flight test platform has multiple configurations of vertical takeoff and landing aircraft to adapt to different levels of mission payload (e.g., 100kg / 10kg / 1kg mission payload), which can meet the test requirements of different scales and test parameters.

[0038] like Figure 2As shown, the data measurement system 202 includes: sensors, converters, data acquisition and editing units, telemetry transmitters, data storage devices, beacon units, and measurement batteries. Each converter powers its corresponding sensor and converts its output into transmittable and computable digital quantities. The data acquisition and editing units power each converter and collect its readings, encoding the content into a centralized data frame format for transmission. The data acquisition and editing units are used to send telemetry data to the telemetry transmitter, which then transmits telemetry radio signals to the ground. The data acquisition and editing units also send stored data to the data storage device for storage and subsequent direct retrieval. The beacon unit continuously transmits radio signals for ground tracking and searching. The measurement batteries directly power the telemetry transmitter, data storage device, and beacon unit.

[0039] In a specific embodiment of the present invention, a sensor acquires analog signals and transmits them to a converter. The converter converts the analog signals transmitted by the sensor into digital signals and sends the digital signals to a data acquisition and editing unit. The data acquisition and editing unit transmits telemetry data to a telemetry transmitter and sends stored data to a data storage device. The telemetry transmitter sends telemetry wireless signals to the telemetry receiver of the test and control system 203, and the telemetry receiver sends telemetry data to the main control computer.

[0040] As a specific embodiment of the present invention, the sensors include temperature sensors, pressure sensors, liquid level sensors, image sensors, and overload sensors, etc. These sensors serve as sensitive devices for monitoring physical quantities such as temperature, pressure, liquid level, image, and overload. The present invention employs methods for measuring images, pressure, distributed multi-point temperature, and high-precision overload within the test tank 20, combined with a data acquisition method that utilizes memory recovery and real-time telemetry dual backup to ensure sufficient acquisition of propellant-related force, thermal, and optical parameters.

[0041] like Figure 2 As shown, the test measurement and control system 203 includes: a ground data radio, a telemetry receiver, and a main control computer. The ground data radio is used to establish two-way wireless communication between the test measurement and control system 203 and the test aircraft 100; the telemetry receiver is used to receive telemetry wireless signals collected by the data measurement system 202; and the main control computer is used to send remote control commands to the test aircraft 100 and monitor the flight status of the test aircraft 100.

[0042] In a specific embodiment of the present invention, the main control computer is configured with a task monitoring module and a data storage module. The task monitoring module provides a human-machine interface, sends remote control commands to the aircraft, and monitors the flight status. The data storage module saves and manages monitoring data and telemetry data. The task monitoring module contains task monitoring software; the data storage module contains data storage software. The task monitoring software is responsible for providing a human-machine interface, sending remote control commands to the aircraft, and monitoring the flight status. The data storage software saves and manages process data records received from the task monitoring and telemetry systems.

[0043] As a specific embodiment of the present invention, different power configurations and aircraft schemes can be adopted according to different mission requirements and test scales.

[0044] The first type is a suborbital flight test platform powered by a liquid rocket engine. This platform can provide a 100kg-class payload testing capability, acquiring payload data through space-to-ground telemetry and parachute recovery. Within this weight range, measurements such as internal tank imaging, pressure, distributed multi-point temperature, and high-precision overload can be used to test approximately 80.0kg of ambient / cryopropellant. The main engine is a liquid rocket engine, with bidirectional oscillation providing thrust and pitch / yaw control forces. The rocket's roll control force is provided by the final attitude control propulsion system. The final attitude control propulsion system is a small liquid rocket propulsion system with multiple thrust units that can be individually controlled by solenoid valves, capable of operating in a reaction control RCS mode. The flight overload range of the test vehicle 100 is axial: 0g to +2.0g, and the lateral flight overload range is ±4.0×10 -4 g~0.40g.

[0045] The second type is a vertical takeoff and landing (VTOL) flight test platform powered by turbojet engines. This platform can provide a 10kg-class payload testing capability, and both the test vehicle 100 and the payload are fully reusable. Within this weight range, external and internal image measurement, along with high-precision overload sensors, can be used to test approximately 7.0kg of room-temperature propellant simulation fluid. The test vehicle 100 uses adjustable thrust engines as its propulsion system, with two turbojet engines providing a maximum takeoff thrust of 1100N, each with a maximum thrust of 550N. Pitch, yaw, and roll attitude control is achieved through the oscillation of two vector nozzles, enabling kilometer-scale VTOL recovery demonstration and verification flight missions. The flight overload range of the test vehicle 100 is 0g to +1.5g axially and ±1.5×10⁻⁶ laterally. -2 g~0.25g.

[0046] The third type is a vertical takeoff and landing (VTOL) flight test platform powered by an electric ducted fan, capable of testing payloads up to 1 kg. Both the aircraft and payload are fully reusable and recoverable. Within this weight range, external image measurement and a high-precision overload sensor can be used to test approximately 0.5 kg of room-temperature propellant simulation fluid. The aircraft uses an adjustable-thrust ducted fan as its propulsion system, with a single 70mm / 6S electric ducted fan providing a maximum takeoff thrust of 2.7 kg. Pitch, yaw, and roll attitude control is achieved through four internal flow rudders, enabling a 3-minute VTOL demonstration flight mission. The flight overload range of the test aircraft 100 is 0g to +1.2g axially and ±1.0×10⁻⁶ laterally. -2 g~0.20g.

[0047] As a specific embodiment of the present invention, the propellant physical parameter measurement scheme is as follows: In order to obtain the actual overload affecting the propellant sloshing in the tank, in addition to the inertial navigation equipment configured on the test aircraft 100 itself, a three-channel high-precision overload sensor is equipped, with a measurement accuracy better than 2×10⁻⁶. -4 g; A fiber optic grating temperature measurement sensor is used to realize multi-point distributed temperature measurement inside the propellant, and a corresponding converter is configured to perform centralized sampling and send the data to the memory and telemetry link.

[0048] Traditional cryogenic monitoring technologies, such as resistance thermometers, are bulky, require point-to-point data acquisition, and are susceptible to strong electromagnetic interference. These limitations make them unsuitable for engineering structures operating in ultra-low temperature environments, such as real-time temperature gradient monitoring of liquid hydrogen fuel tanks in spacecraft. As a novel type of temperature sensor, fiber Bragg grating (FBG) sensors exhibit significant advantages over traditional temperature sensors, including smaller size, ease of embedding in structures, and immunity to electromagnetic interference.

[0049] As a specific embodiment of the present invention, a test method for a propellant sloshing characteristics measurement flight test platform is as follows:

[0050] The propellant liquid to be tested is introduced into the test tank;

[0051] Based on mission requirements and test scale, appropriate power configurations and aircraft schemes are set;

[0052] Simulate the propellant sloshing environment during rocket reentry and return flight;

[0053] The force, heat, and optical parameters of the propellant liquid being tested are measured using sensors.

[0054] As a specific embodiment of the present invention, physical quantities such as temperature, pressure, liquid level, image, and overload of the propellant liquid under test are monitored.

[0055] The beneficial effects achieved by this application are as follows:

[0056] (1) Compared with traditional motion simulation methods, this application can overcome the limitations of short motion stroke, limited test time and limited motion space, and can simulate the bottoming conditions required for the management of propellant in liquid launch vehicle tanks under controlled low overload flight conditions.

[0057] (2) This application uses the same top-mounted tank configuration as vertical take-off and landing liquid launch vehicles to simulate the propellant sloshing environment during rocket reentry and return flight, and measures the force, heat, optical and other parameters of the propellant.

[0058] (3) The flight test platform of this application adopts different power configurations such as liquid rocket engine, turbojet aero engine and electric ducted fan. The flight test platform has multiple configurations of vertical take-off and landing aircraft to adapt to different levels of mission load, and can meet the test requirements of different scales and test parameters.

[0059] In the description of this application, the terms "first" and "second" are used for descriptive purposes only and should not be construed as indicating or implying relative importance or implicitly specifying the number of indicated technical features. Thus, a feature defined as "first" or "second" may explicitly or implicitly include one or more of the stated features. In the description of this application, "multiple" means two or more, unless otherwise explicitly specified.

[0060] In the description of this application, the word "for example" is used to mean "used as an example, illustration, or description." Any embodiment described as "for example" in this application is not necessarily to be construed as being more preferred or advantageous than other embodiments. The following description is provided to enable any person skilled in the art to make and use the invention. Details are set forth in the following description for purposes of explanation. It should be understood that those skilled in the art will recognize that the invention can be made without using these specific details. In other instances, well-known structures and processes will not be described in detail to avoid obscuring the description of the invention with unnecessary detail. Therefore, the invention is not intended to be limited to the embodiments shown, but is consistent with the broadest scope of the principles and features disclosed in this application.

[0061] The above description is merely an embodiment of the present invention and is not intended to limit the invention. Various modifications and variations can be made to the present invention by those skilled in the art. Any modifications, equivalent substitutions, improvements, etc., made within the spirit and principle of the present invention should be included within the scope of the claims of the present invention.

Claims

1. A flight test platform for measuring propellant sloshing characteristics, characterized in that, The platform includes: an experimental flight vehicle and an experimental control system; The test control system includes an aircraft control system, a data measurement system, and a test measurement and control system; The aircraft control system and the data measurement system are installed on the test aircraft; The test and control system is located on the ground. The test aircraft includes a test tank containing the propellant liquid to be tested. The experimental aircraft also includes: a main structure, a landing device, a first power plant, a power tank, a servo mechanism, and a control cabin. The landing device and the servo mechanism are fixedly connected to the rear end of the main structure; The first power unit, the power storage tank, and the control cabin are all located inside the main structure; The test aircraft is positioned perpendicular to the ground. The internal structure of the main body, from top to bottom, consists of: a test tank, a control cabin, a power tank, and a power unit. The power unit and the servo mechanism are located below the center of mass of the test aircraft. The equipment of the aircraft control system and the data measurement system are installed in the control cabin and close to the center of mass of the test aircraft. The test vehicle has the same top-mounted test tank configuration as the vertical takeoff and landing liquid propellant launch vehicle, and adopts a thrust vectoring dynamic layout at the end; it has the same inverted pendulum control method and flight attitude characteristics; it simulates the axial variable thrust and lateral stability control flight mode used by the vertical takeoff and landing liquid propellant launch vehicle during reentry and return flight; it simulates the controlled low-overload flight state, the bottoming conditions required for propellant management in the liquid propellant launch vehicle tank, and the propellant sloshing environment during rocket reentry and return flight; and it measures the force, heat, and optical parameters of the propellant liquid under test through sensors. Depending on the mission requirements and the scale of the test, the following different power configurations and aircraft schemes are adopted: The first type is a suborbital flight test platform powered by a liquid rocket engine, which provides the capability to test 100kg-class mission payloads. The second type is a vertical takeoff and landing flight test platform powered by a turbojet aero-engine, which provides the capability to test mission payloads of up to 10kg. The third type is a vertical takeoff and landing flight test platform powered by an electric ducted fan, which provides the capability to test mission payloads of up to 1 kg.

2. The propellant sloshing characteristic measurement flight test platform according to claim 1, characterized in that, The test tank is equipped with sensors and transducers, which are used to measure the physical parameters of the test tank.

3. The propellant sloshing characteristic measurement flight test platform according to claim 1, characterized in that, The aircraft control system and the data measurement system are located in the control cabin.

4. The propellant sloshing characteristic measurement flight test platform according to claim 1, characterized in that, The aircraft control system includes: navigation equipment, controller, second power unit, servo control mechanism, control battery, power battery and aircraft data transmission radio; The navigation device is used for navigation and positioning; The control battery is used to provide power to the controller; The power battery is used to provide power to the second power device and the servo control mechanism; The aircraft data transmission radio is used for wireless communication on the ground data transmission platform.

5. The propellant sloshing characteristic measurement flight test platform according to claim 4, characterized in that, The second power unit uses a liquid rocket engine, a turbojet aircraft engine, or an electric ducted fan.

6. The propellant sloshing characteristic measurement flight test platform according to claim 1, characterized in that, The data measurement system includes: sensors, converters, data acquisition and editing equipment, telemetry transmitters, data storage devices, beacon units, and measurement batteries.

7. The propellant sloshing characteristic measurement flight test platform according to claim 6, characterized in that, The sensors include temperature sensors, pressure sensors, liquid level sensors, image sensors, and overload sensors.

8. The propellant sloshing characteristic measurement flight test platform according to claim 1, characterized in that, The test and control system includes: a ground data radio, a telemetry receiver, and a main control computer. The ground data transmission radio is used to establish two-way wireless communication between the test measurement and control system and the test aircraft. The telemetry receiver is used to receive telemetry wireless signals collected by the data measurement system; The main control computer is used to send remote control commands to the test aircraft and monitor the flight status of the test aircraft.

9. The propellant sloshing characteristic measurement flight test platform according to claim 8, characterized in that, The main control computer is equipped with a task monitoring module and a data storage module; The task monitoring module is used to provide a human-machine interface, send remote control commands for the aircraft, and monitor the flight status. The data storage module is used to save and manage monitoring data and telemetry data.

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