Thermal Vacuum Test Research Articles

Thermal Vacuum Testing of a Small Loop Heat Pipe with a PTFE Wick for Spacecraft Thermal Control

A loop heat pipe (LHP) is a two-phase heat transfer device that utilizes the evaporation and condensation of a working fluid to transfer heat, and the capillary forces developed in fine porous wicks to circulate the fluid. LHPs have been gaining increased acceptance for spacecraft missions, and recently, small LHPs on the order of a few hundred watts have been investigated for this purpose. In this study, a 100W class small LHP with a polytetrafluoroethylene wick as the primary wick was designed and fabricated for thermal vacuum testing. The LHP has a thermoelectric converter (TEC) to control the loop operating temperature. The thermal vacuum test was conducted to evaluate the LHP's thermal performance under a space-simulated environment such as ultra-high-vacuum, and black body radiation, except for a gravitational effect. The loop showed large thermal hysteresis before and after the large and small head loads. The TEC was able to control the loop operating temperature with a small amount of electrical power.

Qualifciation test series of the indium needle FEEP micro-propulsion system for LISA Pathfinder

The Laser Interferometer Space Antenna project (LISA) is a co-operative program between ESA and NASA to detect gravitational waves by measuring distortions in the space–time fabric. LISA Pathfinder is the precursor mission to LISA designed to validate the core technologies intended for LISA. One of the enabling technologies is the micro-propulsion system based on field emission thrusters necessary to achieve the uniquely stringent propulsion requirements. A consortium consisting of Astrium GmbH and the University of Applied Sciences Wiener Neustadt (formerly AIT) was commissioned by ESA to develop and qualify the micro-propulsion system based on the Indium Needle FEEP technology. Several successful tests have verified the proper Needle Field Emission Electric Propulsion (FEEP) operation and the thermal and mechanical design of subcomponents of the developed system. For all functional tests, the flight representative Power Control Unit developed by SELEX Galileo S.p.A (also responsible for the Micro-Propulsion Subsystem (MPS) development) was used. Measurements have shown the exceptional stability of the thruster. An acceptance test of one Thruster Cluster Assembly (TCA) over 3600 h has shown the stable long term operation of the developed system. During the acceptance test compliance to all the applicable requirements have been shown such as a thrust resolution of 0.1 μN, thrust range capability between 0 and 100 μN, thrust overshoot much lower than the required 0.3 μN+3% and many others. In particular important is the voltage stability of the thruster (±1% over the duration of the testing) and the confirmation of the very low thrust noise. Based on the acceptance test the lifetime of the thruster is expected to exceed 39,000 h generating a total impulse bit of 6300 Ns at an average thrust level of 50 μN. A flight representative qualification model of the Needle FEEP Cluster Assembly (DM1) equipped with one active TCA has performed a qualification program consisting of acceptance, vibration, shock, and thermal vacuum test. During the last test, the thermal vacuum test (TVT), a performance decrease was observed. According to a preliminary analysis, this performance decrease is not linked to the thermal conditions simulated in the TVT but might be rather linked to secondary effects of the TVT set-up.

Stable and self-adaptive performance of mechanically pumped CO2 two-phase loops for AMS-02 tracker thermal control in vacuum

Abstract A mechanically pumped CO 2 two-phase loop cooling system was developed for the temperature control of the silicon tracker of AMS-02, a cosmic particle detector to work in the International Space Station. The cooling system (called TTCS, or Tracker Thermal Control System), consists of two evaporators in parallel to collect heat from the tracker’s front-end electronics, two radiators in parallel to emit the heat into space, and a centrifugal pump that circulates the CO 2 fluid that carries the heat to the radiators, and an accumulator that controls the pressure, and thus the temperature of the evaporators. Thermal vacuum tests were performed to check and qualify the system operation in simulated space thermal environment. In this paper, we reported the test results which show that the TTCS exhibited excellent temperature control ability, including temperature homogeneity and stability, and self-adaptive ability to the various external heat flux to the radiators.

Development of a lightweight deployable/stowable radiator for interplanetary exploration

This paper experimentally and analytically evaluates the heat rejection/retention performance of a lightweight 100 W-class re-deployable radiator with environment-adaptive functions. This radiator, reversible thermal panel (RTP), which consists of flexible high thermal conductive graphite sheets and a single crystal shape memory alloy as a passive reversible actuator, changes its function from a radiator to a solar absorber by deploying/stowing the reversible fin upon changes in the heat dissipation and thermal environment. The RTP was considered as a candidate methodology for thermal control of Venus mission, PLANET-C, in order to save survival heater power. An RTP prototype was tested and evaluated. An analytical thermal model of the RTP was also developed, and basic performances of the RTP were evaluated. Thermal performance of the RTP when applied to the longwave infrared camera of the PLANET-C was evaluated with an analytical thermal model as functions of fin deployment directions and rear surface properties of the RTP’s fin. The analytical results showed that the RTP can save survival heater power in comparison to a conventional radiator. ► New thermal control device for spacecraft with high-thermal-conductivity graphite sheets. ► Heat rejection performance can be changed based on its temperature. ► Thermal vacuum test demonstrates autonomous thermal control performance. ► A case study was conducted, and effectiveness was analytically evaluated.

과학기술위성 3호 S-대역 송신기 비행모델 설계, 제작 및 시험

본 논문에서는 현재 KAIST 인공위성연구센터(SaTReC)에서 개발 진행 중인 과학기술위성3호(STSAT-3) 위성체의 위성 상태정보 전송을 위한 S-대역 송신기 비행모델 개발 및 시험에 관하여 서술하고 있다. STSAT-3의 통신시스템은 크게 상태정보, 명령 송수신을 위한 통신채널과 임무데이터 전송을 위한 통신채널로 구성되며 상태 정보 및 명령 송수신을 위하여 S-대역을 사용하며 임무데이터 전송용으로 X-대역 주파수를 사용하고 있다. S-대역 송신기(S-band Transmitter, STX)는 기능적으로 변조기, 주파수 합성기, 파워 앰프 및 전력공급기 로 구성되어 제작되었다. 전송데이터 변조 방법으로는 주파수 천이방식(Frequency Shift Keying, FSK)을 사용 하며 위성본체와의 데이터 통신은 표준방식인 RS-422통신 방식을 사용하였다. STSAT-3의 STX는 기능적 모듈화에 근거하여 설계 및 제작 되었으며 1.5W(31.7dBm) 송신 출력에 <TEX>$1{\times}10^{-5}$</TEX> 비트오율(BER) 성능을 만족한다. STX 비행모델은 성능시험, 환경시험(진동시험, 열진공시험)을 성공적으로 마치고 위성체 조립단계에 납품되었다. This paper describes the development and test result of S-band Transmitter flight model(FM) of STSAT-3 by satellite research center(SaTReC), KAIST. The communication sub-system of STSAT-3 is consist of two different frequency band channels, S-band for Telemetry & Command and X-band for mission data. S-band Transmitter(STX) functionally made of modulator, frequency synthesizer, power amp and DC/DC converter. The transmission data is modulated by FSK(Frequency Shift Keying) and the interface between spacecraft sub-module and STX is RS-422 standard method. The FM STX is based on modular design. The RF output power of STX is 1.5W(31.7dBm) and BER of STX is under <TEX>$1{\times}10^{-5}$</TEX> which meets the specification respectively. The FM STX is delivered Spacecraft Assembly, Integration and Test(AIT) level through the completion of functional Test and environmental(vibration, thermal vacuum) Test successfully.

Identification and Correction of Residual Image in the $\hbox{O}_{2}$ A-Band of the Orbiting Carbon Observatory

The detector used for the O2 A-band (0.76 μm) of the National Aeronautics and Space Administration's Orbiting Carbon Observatory (OCO) employed a HyViSI Hawaii-1RG sensor, operating at 180 K in a rolling read-out mode. During the thermal vacuum testing of the flight instrument, it was discovered that the detector exhibited residual images that lasted for many seconds and were of sufficient magnitude to compromise the mission objectives. Independent testing of flight-spare detectors revealed that the problem was common to all and was not simply a fault of the flight detector. The residual image was found to depend upon even-order derivatives of the spectrum, and its decay was a function of the number of frames rather than time. An empirical model was developed, which represented the measured spectrum in terms of the true spectrum and a history of all previous changes in the spectra. On the basis of the model, an algorithm was devised to correct spectra for the effects of residual image, using a time-marching analysis of a history of previous spectra. The algorithm was tested with spectra acquired during the second thermal vacuum test of OCO and was found to reduce the effect of residual image to almost the noise level of the detector. Numerical simulations indicate that residual image has a negligible impact on retrieved concentrations of O2 and CO2 once the spectra have been corrected.

Testing of the Superconducting Magnet and Cryogenics for the AMS-02 Experiment

The superconducting magnet, cryogenics, and detector systems of the AMS experiment was fully integrated and tested in test beam at CERN during 2009. In Spring 2010 the experiment underwent thermal vacuum tests at ESTEC, where it was operated in conditions simulating those that will pertain in orbit. All elements of the superconducting magnet and cryogenics performed as designed, and equilibrium operation was attained at several values of vacuum case temperature. Details of the tests are presented. A thermal model of the overall cryogenic system was calibrated from those measurements. The model was used to predict the cryogenic lifetime of the experiment, as it would be staged on ISS, to be (28 &amp;#x00B1; 6) months.

Operation characteristics of AMS-02 loop heat pipe with bypass valve

Loop heat pipes (LHPs) were designed for the alpha magnetic spectrometer (AMS-02) to dissipate heat from the cryocoolers. A bypass valve is applied to the LHP to keep the cryocooler temperature above its limit (−20°C) in cold environment. Extensive experiments were performed on operation characteristics of LHPs with the bypass valve for AMS-02 during thermal vacuum and thermal balance (TVTB) test. We found that the bypass valve can start up successfully in cold environment. With the bypass valve, the evaporator temperature is stable and can meet the requirement of the cryocooler. We analyzed three operating modes of the bypass valve. Set point temperature and regulation temperature shifts were observed and their relations with the bypass valve temperature were given.

저궤도위성 광학탑재체의 지상 열진공 시험을 위한 예비 열해석

인공위성의 열제어는 인공위성이 운용궤도상에서 겪는 고진공, 극한의 온도변화 환경에서 위성 구성품의 온도변화를 허용한계 온도 범위 내에서 유지하는데 목적이 있다. 본 연구에서는 저궤도 관측위성(LEO)의 광학탑재체에 대한 열해석 과정으로 열진공 시험 조건, 열진공 챔버의 형상, 위성 탑재체 내부의 열적 환경을 고려하여 열해석 모델을 구성하고 궤도 조건에 따른 열해석을 수행하였다. 또한 광학탑재체의 지상 열진공 시험 조건에 따른 열해석 수행하여 열진공 시험을 위한 시험조건을 정립하였다. The purpose of satellite thermal control design is to maintain all the elements of a spacecraft system within their temperature limits for all mission phases. The thermal analysis model for Low Earth Orbit satellite payload level simulation is established by considering thermal vacuum test environment condition, thermal vacuum chamber configuration, and satellite's payload inner thermal environment. The established thermal analysis model is used to determine thermal vacuum test conditions and test case requirements.

Large Vacuum Systems for Space Developments

Spacecrafts are subjected to severe environmental conditions. Loud sound and vibration are generated in their launch, and separation of spacecrafts from a rocket causes a great shock. Moreover, they are exposed to high-vacuum, very low temperature close to absolute zero, high temperature due to solar radiation, and other severe conditions not found on the earth. Once spacecrafts are launched, repair of them is almost impossible. So, it is important to make sure spacecrafts will function properly in such environment before lift-off. They are tested using test facilities which simulate the environmental conditions during launch and flight. In particular, this review shows thermal vacuum test and a space chamber, which simulates high-vacuum, very low temperature, and solar radiation.

STM 열진공 시험 결과를 이용한 EOS-C Ver.3.0 열제어계 설계 검증

(주)쎄트렉아이는 지구관측위성의 주 탑재체로 사용될 고해상도 전자광학카메라, EOS-C Ver.3.0을 개발 중이다. EOS-C Ver.3.0은 현재 운용중인 DubaiSat-1의 주 탑재체인 EOS-C Ver.2.0 개발 경험을 바탕으로 능동 열제어 방식과 수동 열제어 방식을 적절하게 혼용하여 보다 향상된 성능을 갖도록 설계되었다. 설계를 바탕으로 STM을 개발하여 인증(qualification) 수준의 열진공 시험을 수행하여 설계 여유(design margin)를 확인하였다. 또한 열평형 시험 결과를 이용하여 열제어계 설계에 사용한 열-수치 모델에 대한 검증 작업을 수행하였으며, 열-수치 모델이 실제 열적 특성을 잘 모사하고 있음을 확인하였다. A high-resolution electro-optical camera (EOS-C Ver.3.0), the mission payload of an Earth observation satellite, is under development in Satrec Initiative. We designed this system to give improved thermal performance compared with the EOS-C Ver.2.0 which is the main payload of DubaiSat-1 by optimizing the active and passive thermal control design. We developed the Structural-Thermal Model (STM) and verified the design margin by performing the qualification level thermal vacuum test. We also conducted the verification of its Thermal Mathematical Model (TMM) through the thermal balance test. As a result, it was confirmed that TMM faithfully represents the thermal characteristics of the EOS-C Ver.3.0.

소형 위성용 소형직류모터를 이용한 분리장치

본 논문에서는 위성 본체에 장착된 탑제체를 위성으로부터 분리하기 위한 비폭발식 분리 메커니즘을 제안하였다. 작동기로서 경제적이고 신뢰성 있는 유성기어가 통합된 소형 직류모터를 채택하였다. 제안된 분리장치는 별도의 지그나 구조물 없이도 재장전이 용이 하도록 하였으며, 단순한 구조로 제작이 수월한 장점이 있다. 완성된 분리장치는 분리반응속도 측정시험, 최대분리가능하중 측정 시험, 최대발생충격 측정 시험을 통하여 그 성능을 입증하였고, 진동시험, 열진공 시험을 통하여 우주 환경에서 사용 가능함을 보여주었다. 개발된 인공위성용 비폭발식 분리장치를 통하여 그동안 사용되었던 폭발식 분리장치 등의 해외위성용 분리장치를 대체 할 수 있을 것으로 기대된다. This paper describes development of a non-explosive separation device which can be equipped on small satellites. The spur geared micro DC motor, which has high reliability and advantage of price, is adopted as an actuator. The proposed separation device has resettability and it does not need extra jig to reload. In addition, the simple structure makes it easy to fabricate and assemble. To verify the performance of the proposed device, the response time tests, maximum preload tests and maximum shock level tests were performed. Also, through the vibration tests and thermal vacuum tests, feasibility of the proposed separation device was shown in launching and space environments. Therefore, we expect that the proposed separation device can replace the imported separation devices in near future.

인공위성 탑재품 수준 열진공 시험에 대한 열해석 모델의 개발과 환경시험 결과를 이용한 검증

Thermal analysis for the simulation of satellite component level thermal vacuum test processes was carried out by considering thermal vacuum test environment condition, thermal vacuum chamber configuration, and satellite`s inner thermal environment. The transient analysis results can be obtained for the temperatures of component and thermal vacuum chamber assemblies. The thermal analysis model was verified with the component thermal environmental test results by using enhanced thermal vacuum chamber.

月基对地观测跟踪转台设计与分析

Aiming at the mechanical condition during the moon landing,complex environment lunar surface features,and strict weight requirements,a tracking turntable for lunar-based earth observation was designed which consists of two main components: horizontal shaft elevation actuator modules,and vertical shaft azimuth actuator modules.In addition,Hour sensors were selected as the rotation position detector of elevation and azimuth shafts,and the pointing accuracy is better than 0.1°.An effective tracking turntable design configuration,optimized structure of supporter and selection of material were performed based on the finite element analysis method,and the stiffness of axis system and transmission accuracy of actuator module were analyzed to obtain the trade-off design with better capabilities of deep space environmental adaptability.To qualify the design for the specific mission requirements,the vibration and thermal vacuum tests were carried out with the result of 41 Hz eigenfrequencies and the required running temperature changing range(—40℃~+80℃).The experimental results show that the proposed design has advantages of compact volume,light weight and high reliability.

과학기술위성 3호 S-대역 송신기 인증모델 설계 및 제작

본 논문은 현재 KAIST 인공위성연구센터(SaTReC)에서 개발하고 있는 과학기술위성3호 (STSAT-3) 위성본체의 S대역 통신채널 중 송신기 개발에 관하여 서술하고 있다. STSAT-3의 통신시스템은 위성제어 및 상태정보 전송을 위하여 S대역을 사용하며 임무수행용 통신채널로는 X대역을 사용하고 있다. S대역 송신기(S band Transmitter, STX)는 변조기, 주파수 합성기, 파워 앰프, 전력공급기 로 구성되어 있다. 변조 방법으로는 주파수 천이방식(Frequency Shift Keying, FSK)을 사용 하며 위성체와는 RS-422통신 방식을 이용한다. STSAT-3의 STX는 모듈화에 근거하여 설계 및 제작 되었으며 1.5W(31.7dBm) 송신 출력에 1E-5 비트오율(BER) 성능을 만족한다. 현재 성능시험, 환경시험(진동시험, 열진공시험)을 성공적으로 마쳤다. This paper describes the development result of S-band Transmitter of STSAT-3 by satellite research center(SaTReC), KAIST. STSAT-3 has two kinds of communication channels, S- band for Telemetry & Command and X-band for mission payload. S-band Transmiiter(STX) consist of modulator, frequency synthesizer, power amp and DC/DC converter. The modulation scheme of STX is FSK(Frequency Shift Keying). The interface between spacecraft OBC and STX is RS-422. The STX is based on modular design. The RF output power of STX is 1.5W(31.7dBm) and BER of STX is under 1E-5. The Test of STX is completed successfully such as functional Test and environmental(vibration, thermal vacuum) Test.

과학기술위성 3호 대용량 메모리 유닛의 인증모델 설계 및 구현

본 논문에서는 개발된 과학기술위성3호 대용량 메모리 유닛의 검증모델의 성능 및 환경시험 결과를 나타내었다. 과학기술위성3호 대용량 메모리 유닛은 적외선 영상시스템(MIRIS)과 초소형 영상 분광기(COMIS)에서 최대 100Mbps로 수신한 데이터를 32Gb의 메모리에 저장한 후 지상으로 10Mbps의 속도로 전송하는 임무를 수행한다. 탑재체 데이터 수신 시험과 데이터 수신 시스템으로의 데이터 전송 시험을 통해서 성능을 검증 하였다. 또한 발사체 환경과 우주 환경에서의 성능 확인을 위해서 진동 시험과 열진공 시험을 수행 하였다. This paper describes the design and test results of engineering qualification model(EQM) of mass memory unit(MMU) for STSAT-3. The MMU for STSAT-3 having 32Gb mass memory capacity is capable of receiving and transmitting the mission data from MIRIS(Multi-purpose IR Imaging System) and COMIS(Compact Imaging Spectrometer) at 100Mbps and 10Mbps. The performance of EQM MMU was verified by the tests of data receiving from two payloads and data transmission to the data receiving system. Moreover, the vibration and thermal vacuum test was performed to verify the launch vehicle and space environments.

저궤도 인공위성 열-구조 모델 열진공시험 결과를 활용한 열모델 보정

Because thermal design of satellite carrying out mission in space is performed by thermal analysis result using thermal mathematical model, accuracy of thermal mathematical model is important and it can be improved by model correlation. Correlation steps of satellite thermal math model are composed of modeling of satellite configuration placed in thermal vacuum chamber, verification of correspondence between thermal math model and real satellite configuration, and adjustment of modeling parameters from major part to minor part etc. In this study, correlation success criteria was established and correlation for satellite thermal math model was performed using result of thermal vacuum test of satellite structure-thermal model to meet the success criteria. The overall results satisfied the criteria and this correlated thermal model was applied for detailed thermal design of satellite.

System for establishing best focus for the Orbiting Carbon Observatory instrument

A technology for establishing best focus of an optical system is described. This technology was recently used to establish best focus of the Orbiting Carbon Observatory spectrometers while the instrument was undergoing thermal-vacuum testing. Three piezo-actuated motors were used to adjust the tip, tilt, and piston of a focal plane assembly relative to the spectrometer's optical system. A set of optical displacement sensors measured tip-tilt-piston throughout the focusing process. With best focus established and confirmed using a pupil-slicing technique, the corresponding sensor measurements were used to specify the geometry and dimensions of a precision-ground shim ring.

Environmental Testing of NASA's Evolutionary Xenon Thruster Prototype Model 1 Reworked Ion Engine

NASA’s Evolutionary Xenon Thruster propulsion system is an advanced ion propulsion system that is oriented toward robotic exploration of the solar system using solar electric power. The subsystem includes an ion engine, a power processing unit, feed system components, and a thruster gimbal. The prototype model engine, PM1, was subjected to qualification-level environmental testing to demonstrate compatibility with environments representative of anticipated mission requirements. Although the testing was largely successful, several issues were identified, including the fragmentation of potting cement on the discharge and neutralizer cathode heater terminations during vibration, which led to abbreviated thermal testing, and the generation of particulate contamination from manufacturing processes and engine materials. The engine was reworked to address most of these findings and renamed PM1R, and the environmental test sequence was repeated. Thruster functional testing was performed before and after the vibration and thermal-vacuum tests. Random vibration testing, conducted with the thruster mated to the breadboard gimbal, was executed at 10:0 grms for 2min in each of the three axes. Thermalvacuum testing included three thermal cycles from 120 to 215 C with hot engine restarts. Thruster performance was nominal throughout the test program, with minor variations in a few engine operating parameters likely caused by facility effects. There were no significant changes in engine performance as characterized by the engine operating parameters, ion optics performance measurements, and beam current density measurements, indicating no significant changes to the hardware as a result of the environmental testing. The redesigned cathode heater terminations successfully survived the vibration environments. NASA’s EvolutionaryXenonThruster PM1Rengine and the breadboard gimbal were found to be well designed tomeet environmental requirements based on the results reported herein.

Tribological aspects of a strain wave gearing system with specific reference to its space application

Strain wave gearing (SWG), also known as harmonic drive gear, has been used in many space applications. In 2003, a development of grease-lubricated-type SWG for space flight applications was begun. The life tests of the SWG exposed to vibration test and thermal vacuum test were performed in vacuum. The life test results showed that the in-vacuum operating life of the SWG was significantly short compared with the in-atmosphere life, which is normally corresponding to a rolling fatigue life of wave generator bearing. This may be mainly caused by the severe wear at the wave generator—flexspline interface, where the lubricant is not supplied easily into the clearance in vacuum.