Thermal Vacuum Test Research Articles

On-orbit Experiment Plan of Loop Heat Pipe and the Test Results of Ground Test

It is becoming increasingly difficult to meet the challenging thermal control requirements of modern spacecraft missions with only existing thermal control devices such as conventional heat pipes. A loop heat pipe (LHP) is an effective method to overcome some of these thermal control constraints. The LHP is a passive two-phase heat transfer device that utilizes the evaporation and condensation of a working fluid to transfer heat and capillary force to circulate the fluid. The LHP can transport much heat for a long distance against gravity and has many other excellent characteristics, such as high controllability of operating temperature and a shutdown function. In this study, LHPs for space application have been developing. As a part of the study, a bread board model (BBM) of LHP was designed and fabricated. As a result of an on-ground test of the BBM, it was confirmed that the BBM fulfilled all requirement (e.g. maximum heat transport rate, minimum required heat load for start-up, operating temperature control and shutdown function). To adopt the LHP as a heat transfer device in practical spacecraft mission, the thermal characteristics under micro-gravity conditions should be examined in advance. An on-orbit experiment of a LHP radiator system is planned. This paper describes the test plan of on-orbit experiment of a LHP radiator system on the International Space Station (ISS) and the results of thermal vacuum test of flight model for on-orbit experiment on ground.

Geometric correction for TanSat atmospheric carbon dioxide grating spectrometer

Carbon Observing Satellite (TanSat) is the first of Chinese experimental satellite for monitoring the CO2 column densities in the atmospheric, while identifying sources and sinks on regions scales. Three grating spectrometers are equipped on TanSat to measure the reflected sunlight in the 0.76μm O2A-band, 1.61μm weak CO2 band and 2.06μm strong CO2 band respectively. Three bands have high spectral resolution and signal to noise ratio (SNR), while their spectral and radiometric calibration have been reported in many articles. This paper describes the geometric calibration of the essential payload (ACGS) on the TanSat. In the assembly process, the flatness only in the spatial dimension is strictly controlled owing to the physical obstruction. The geometric deflection error in one band should be tested and corrected by adjusting the starting rows of different spectral samplings. Meanwhile, the alignment error among the three bands also need processing in the same way to ensure that the slits of the three bands were matched the corresponding geographic target. Finally, the geometric correction results of the thermal-vacuum testing suggested that the alignment accuracy is well enough to match the requirements.

Hardware and GNC solutions for controlled spacecraft re-entry using aerodynamic drag

Traditionally, controlled spacecraft re-entries have been conducted using propulsive de-orbit burns which are risky, expensive, and may not be possible for all vehicles. Recently, the miniaturization of technology has ushered in a new class of small satellites (such as CubeSats) that are too small to host thrusters but may require a controlled de-orbit if they contain materials capable of surviving re-entry. For all space vehicles requiring a controlled re-entry, the ability to harness the naturally occurring aerodynamic drag force for orbit control provides a cheaper and more reliable alternative to chemical propulsion.This paper discusses a comprehensive method for drag-controlled re-entry that is applicable to any vehicle capable of modulating its ballistic coefficient. First, a novel guidance generation algorithm efficient enough to run onboard a CubeSat outputs a desired ballistic coefficient profile and corresponding numerically propagated trajectory that if followed, will lead the spacecraft to a desired de-orbit location. This guidance generation algorithm is based on an analytical solution that provides convergence guarantees, ensures rapid performance, and facilitates a controllability analysis. Next, the guidance tracking algorithm utilizes an extended Kalman filter and GPS measurements to estimate the position and velocity of the satellite relative to the guidance. A full state feedback linear-quadratic-regulator (LQR) control strategy is then used to drive the relative position and velocity to zero using solely aerodynamic drag. This paper also discusses a novel retractable drag de-orbit device (D3) that can be attached to existing CubeSat structures and can easily be scaled up for larger satellites. The D3 provides passive three-axis attitude stabilization using aerodynamic and gravity gradient forces and can be repeatedly modulated to perform aerodynamically-based orbital maneuvering and controlled re-entry. The design of the planned 2U CubeSat to test the D3 and control algorithms in flight is also discussed.The re-entry point targeting algorithms were validated through extensive Monte Carlo simulations which included realistic GPS measurement errors and drag force uncertainties. The algorithms were able to guide the satellite to a desired de-orbit location with an average error below 25 km and in all cases, the targeting error was low enough for debris mitigation purposes. The accuracy and reliability of these algorithms coupled with the D3 device that has successfully undergone thermal vacuum, vibration, and fatigue testing provide a cheap, reliable, and comprehensive attitude, orbit, and de-orbit control solution that can be used on large and small space vehicles, possibly replacing conventional propulsion and attitude control systems and making space more accessible to everyone.

Controlled Spacecraft Re-Entry of a Drag De-Orbit Device (D3)

The University of Florida Advanced Autonomous Multiple Spacecraft (ADAMUS) lab has developed a drag de-orbit device (D3) for CubeSats (small spacecraft intended for university and research group use). By modulating the D3 drag area, decaying orbital maneuvering and partial attitude control can be performed, and the host satellite can be made to de-orbit in a desired location. This paper details the design, manufacturing, and testing of the D3. Four retractable deployers are used to vary the drag area from 0.01 m2 to 0.5 m2. Each deployer is actuated independently using a brushed DC motor to drive the boom. An encoder affixed to the deployer measures the distance that the boom travels and the number of rotations it takes to reach that distance, and is used to stop the rotation after a preset distance. All manufacturing of the D3 device is performed in-house using a Computer Numerical Control (CNC) milling machine and manual lathe. Testing of the D3 consists of thermal testing, fatigue testing, and vacuum testing, which are also discussed in this paper.

Control System Design of Microwave Load Water-Cooling for Spacecraft Thermal Vacuum Test

The microwave energy released by satellite transponder in the vacuum heat test needs to be converted into heat energy and dissipated through the water-cooling system. To dissipate this heat energy, a microwave load water-cooling system is designed, which realizes the automatic operation and remote monitoring of the water pump and refrigeration unit through Siemens S7-300 and IFIX 5.0 configuration software.

Design of Hot-bulb Probe Low-Pressure Wind Speed Calibration System

In order to measure the wind speed in Mars rover thermal vacuum test correctly, a low-pressure wind speed calibration system need to be developed. The purpose of this paper is to establish a low pressure-wind speed calibration system, and give an error analysis of it. In this paper, the Hot-bulb Probe calibration scheme, calibration conditions are introduced. The errors caused by length measurement and speed measurement, disturbance of low field, measuring instrument are fully analyzed.

Creation of control algorithms thermal vacuum tests in the SimInTech environment

Paper is divote to results of the development of the thermal vacuum tests computer model of allowing to investigate impact of external factors on the devices which are in conditions of own external atmosphere of the spacecraft in the SimInTech environment of the technical systems dynamic modeling. For a solution an objective models of a control object of the thermal vacuum TVAC-1400 camera are created, and the regulators executed on the basis of the programmable industrial controller Siemens S7-300 setup of key parameters is made, control algorithms are implemented by the modes of tests. The developed model allows to exercise control of test programs of the different equipment, prevention of failures at an error of the operator or interruptions in engineering networks, to reveal the most adverse situations which can arise process of functioning of the thermal vacuum camera. Besides, the presented model can be used for a research of behavior of controlled parameters of thermal vacuum tests and working off of control algorithms. In article development process and settings of model of thermal vacuum tests, its submodels and a control bar, implementation of the main algorithms are step by step described. Also results of adequacy check to the developed model are presented.

Design of the mirror antenna of a spacecraft with the ultralight high precision size-stable reflector

The paper discusses the existing design of mirror antennas used on spacecraft. The requirements are formulated under which the reliable performance of antennas is ensured under operating conditions. Innovative construction of the space high frequency axisymmetric mirror antenna of an ultralight class is presented. Antenna’s body consists of several conjugate carbon fiber shells. Antenna’s design technology which gives an opportunity to create the shell of a reflector with minimal deflection from a full-paraboloid is developed. Using the finite element method, a parametric modal analysis of the antenna is performed. Based on this analysis, geometric parameters were determined that provide optimal mechanical and mass characteristics of the structure. The calculation results were used to create a prototype mirror antenna. It is shown that the created antenna possesses the parameters required for generation of the high-directional electromagnetic emission of Q and V frequency domain. Successful ground experimental perfection of the construction of the antenna is made. It included the stages of mechanical, thermal vacuum, and radio engineering tests. The proposed design can be used to create advanced spacecraft.

Preflight calibration of the Chinese Environmental Trace Gases Monitoring Instrument (EMI)

Abstract. An environmental trace gases monitoring instrument (EMI) is a nadir-viewing wide-field imaging spectrometer, which aims to quantify the global distribution of tropospheric and stratospheric trace gases, and is planned to be launched on 9 May 2018. The selected wavelength bands for EMI are ultraviolet channels: UV1 (240–315 nm), UV2 (311–403 nm) and visible channels: VIS1 (401–550 nm), and VIS2 (545–710 nm). The spectral resolution is 0.3–0.5 nm, and the swath is approximately 114∘ wide to achieve a one-day global coverage. The preflight calibration of the EMI is discussed in this paper. A tunable laser and rotating platform are adopted for an EMI wavelength calibration of the entire field of view. The accuracy of the wavelength calibration is less than 0.05 nm. In addition, the solar calibration mode shows the same results compared with Earth observation mode. A thermal vacuum test is performed to investigate the influence of in-orbit thermal vacuum conditions on the EMI, and EMI spectral response changes with pressure, optical bench temperature, and charge-coupled device (CCD) detector temperature are obtained. For a radiometric calibration of UV1, a diffuser plate with a 1000 W xenon lamp, which produces sufficient UV output, is selected. An integrating sphere system with tungsten halogen lamp is selected for the UV2, VIS1, and VIS2. The accuracies of radiance calibration are 4.53 % (UV1), 4.52 % (UV2), 4.31 % (VIS1), and 4.30 % (VIS2). The goniometry correction factor and irradiance response coefficient of the EMI are also calibrated on the ground for an in-orbit calibration of the solar. A signal-to-noise ratio (SNR) model of the EMI is introduced, and the EMI in-orbit SNR is estimated using the SNR and MODTRAN radiance models.

General Concept of Thermal Vacuum Tests of Space-Appliance Elements at Helium Temperatures

Three main types of cryothermal vacuum units for testing spacecraft products at liquid-helium temperature are considered. Methods for heat removal from test objects depending on their geometric and thermophysical properties are presented. The designs of the different thermal vacuum chambers are discussed considering the thermophysical processes occurring in them. The operating efficiency of the thermal vacuum chambers and their advantages and disadvantages are estimated based on an analysis of heat-transfer processes.

An easy-to-implement thermal test system for large deployable antennas

Large deployable antennas (LDAs), having a wide range of applications in aerospace engineering, encounter extreme thermal conditions when subjected to the space environment. Ground thermal test facilities are used to validate LDA deployability and accuracy under extreme thermal conditions. General antenna thermal test facility is thermal vacuum test facility, which is complex and costly as the vacuum test chamber has to be pumped down and large enough to accommodate LDAs. In this paper, an easy-to-implement thermal test system is presented, which simulates atmospheric thermal environment for LDAs using an air cycle refrigeration system and electric heaters. Additionally, a series of measures are used to ensure uniform temperature and limit airflow turbulence. Test results show that the test system can provide dry, uniform temperature and small disturbance thermal environment for LDAs.

Thermal Vacuum and Swivel Table Tests of a CMG and Fault Mechanism Analysis

AbstractThermal vacuum and swivel table tests of a control moment gyroscope (CMG) are designed and the start-up hesitation and rotation break of the CMG are analyzed. In order to study the fault me...

Evaluation of supercapacitors for space applications under thermal vacuum conditions

Commercially available supercapacitor cells from three separate vendors were evaluated for use in a space environment using thermal vacuum (Tvac) testing. Standard commercial cells are not hermetically sealed, but feature crimp or double seam seals between the header and the can, which may not maintain an adequate seal under vacuum. Cells were placed in a small vacuum chamber, and cycled between three separate temperature set points. Charging and discharging of cells was executed following each temperature soak, to confirm there was no significant impact on performance. A final electrical performance check, visual inspection and mass check following testing were also performed, to confirm the integrity of the cells had not been compromised during exposure to thermal cycling under vacuum. All cells tested were found to survive this testing protocol and exhibited no significant impact on electrical performance.

Brush-Wear Reduction in High-Altitude Direct-Current Motors

The factors of increased wear of current-carrying brushes in direct-current motors designed for high-altitude operation are considered. The results of testing six motors with different degrees of airtightness are presented. It is found out that increased airtightness makes the lacquer film less resistive (thinner) and increases the brush wear. A method of decreasing the wear of current-carrying brushes at standard and high altitudes is proposed that consists in installing a DMS-140 molybdenum disulfide lubricating brush. The lubricating brush is installed in an auxiliary brush holder or in a pocket structurally connected with the current-carrying brush holder on the header runway. The effect of the collector profile material (copper, cadmium bronze, chromium bronze) on the wear of current-carrying brushes is studied, as well as the efficiency of installing lubricating brushes for standard and high-altitude operations. The effect of lubricating brushes on the performance of a 5-W motor is considered as well as the results of testing seven 5-W motors equipped with lubricating brushes in conformity with the machine and warranty service specifications. Information is given on the results of long-term storage (for 1 and 1.9 years) of four motors with different capacities equipped with lubricating brushes, as well as in thermal vacuum tests of 10-W motors under p = 10–5–10–6 mm Hg.

Research of digital temperature measurement system in vacuum thermal test based on DS18B20

Temperature measurement is a very important measurement project in spacecraft vacuum thermal test, and thermocouple measurement system is generally used for temperature measurement. In order to reduce the number of wire measuring circuit and improving the measurement system of anti disturbance and high measurement accuracy in vacuum thermal test process, which can be used in digital temperature measurement system of vacuum thermal test design, realizes the digital temperature measuring equipment of spacecraft and ground equipment. The system is composed of digital temperature sensor DS18B20 and acquisition device. It can be connected with remote monitoring computer through LAN network to realize remote monitoring of temperature. The hardware structure, communication protocol and software design of the digital temperature measurement system are given in this paper. The vacuum thermal environment test and comparison with the platinum resistance temperature measurement system. The results show that the stability of the system is good and the difference between the measured value and the PT100 measurement is within 0.5 °C, and the linearity is about ±0.1% which satisfies the requirement of vacuum thermal test.

Research on Temperature Control Technology of Thermal Vacuum Test in Aerospace Field

The temperature control system is an indispensable key system for the thermal vacuum test. It is used to simulate the infrared heat flow environment of the spacecraft in orbit. In order to improve the stability and reliability of temperature control, this paper designs and develops a thermal vacuum test temperature control system based on the Modbus RTU serial communication protocol on the LabVIEW platform. The system combines the flexibility of software development on LabVIEW platform and combines the stability of serial communication with controller. This temperature control system has been used in thermal vacuum test, the control effect of overshooting within 1°C and temperature control accuracy within ±0.2°C. The system has high stability, which achieves the expected goal.

Proposal of Functional Thermal Control Systems for High-Power Micro-Satellite and Its Demonstration under Thermal Vacuum Condition

In previous years, several high-power micro-satellites below ~100 kg have been developed for high-functional spacecraft. This paper proposes a functional and high-power thermal control system with no power supply and a simple configuration for micro-satellite: 100 W, 3 U. The proposed system consists of a heat storage panel (HSP) with pitch type CFRP (Carbon Fiber Reinforced Polymer), a micro loop heat pipe (m-LHP) and a flexible re-deployable radiator (FRDR) as an active thermal control system. The aim of this research is to try not only to verify the thermal control devices, but also to perform a water phase change experiment as a payload using an electric power generation of 100 W in space environment. In this paper, the basic design of the satellite, the analysis of the feasibility by the thermal mathematical model, and the fabrication of thermal test model including water phase chamber are reported. The main results of thermal analysis as feasibility verification showed that the paddles could absorb the thermal energy up to 97 W at the solar input of 180 W, and the operating temperature of bus equipment became within the allowable temperature range (0°C - 40°C). At thermal vacuum test, the difference between the analysis and the experiment for the temperature history of water due to the discordance for the value of thermal conductance was discussed.

Investigation on the Thermal Control Algorithm of the Heat Sink Temperature Regulating System for Thermal Vacuum Tests

The heat sink temperature is often set in a semi-empirical way based on steady-state temperatures . As a result, the entrance parameters of the heat sink working fluid are often too conservative, and the rate of temperature rise is difficult to control. In this paper, transient thermal models for each component of the heat sink temperature regulation system are established. Then they are programmed into a dynamic simulation model by using Matlab/Simulink as the thermal control algorithm.It is shown that the model can accurately reflect the dynamic and steady state characteristics of the heat sink temperature regulation system, and can provide guidance for the selection of control strategies and working fluid parameters in the thermal vacuum test.

The Design of Large Arc-shape High-precision Walking Device in TV (thermal-vacuum) Conditions

The large-sized space environmental simulation test facility is mainly used for providing thermal radiation environment in high-vacuum, cold and dark space for satellites, spacecraft, lunar spacecraft to carry out whole satellite thermal vacuum test and for large equipment like antenna to carry out tests in TV conditions. Monitoring the spacecraft’s surface optical image or temperature is a main task in space environment simulation test. In previous tests, optical image test mainly took place in a fixed position, so the measuring location and angle are limited. This paper focuses on large spherical space environmental simulation test facility, and designs a large arc-shape high-precision walking device in a space environment simulation test device. It introduces key technology in detail like structure design, thermal design, processes and manufacturing, etc. The test results show that this device can work steadily and reliably under simulation space environments, and the precision exceeds 0.5°.

Characterization testing of Lockheed Martin high-power micro pulse tube cryocooler

This paper describes the thermal vacuum, microphonics, magnetics, and radiation testing and results of a Lockheed Martin high-power micro pulse tube cryocooler. The thermal performance of the microcooler was measured in vacuum for heat reject temperatures between 185 and 300 K. The cooler was driven with a Chroma 61602 AC power source for input powers ranging from 10 to 60 W and drive frequency between 115 and 140 Hz during thermal performance testing. The optimal drive frequency was dependent on both input power and heat reject temperature. In addition, the microphonics of the cooler were measured with the cooler driven by Iris Technologies LCCE-2 and HP-LCCE drive electronics for input powers ranging from 10 to 60 W and drive frequency between 135 and 145 Hz. The exported forces were strongly dependent on input power while only weakly dependent on the drive frequency. Moreover, the exported force in the compressor axis was minimized by closed loop control with the HP-LCCE. The cooler also survived a 500 krad radiation dose while being continuously operated with 30 W of input power at 220 K heat rejection temperature in vacuum. Finally, the DC and AC magnetic fields around the cooler were measured at various locations.