On the collaborative design and simulation of space camera: stop structural/thermal/optical) analysis

Numerical control machine tools are widely used in the manufacturing nowadays. The engineers often have difficulties discussing with another engineer about the problem in the CAM work, so a convenient tool for this is needed. In this paper, a framework of collaborative CAM simulation environment is presented for the goal of virtual enterprise. The main modular and key technology of them are illustrated, which include virtual part discussion environment, cutting parameter database management modular, G-code collaborative simulation modular and VRML based realistic cutting simulation modular. The prototype of system is implemented for small amount of users. In the end, the application of VRML based CAM simulation environment is discussed.

In order to verify the spacecraft performance in extreme temperature and vacuum, and to screen spacecraft early defect, generally spacecraft TV (Thermal Vacuum) test should be carried out before launch. Designed in small size and with low cost, nanosatellite is made from a large number of COTS (Commercial off the shelf) components; therefore, the test should be low-cost, simple and quick. With the intention of screen out early defects of the product in lower cost, nanosatellite developers usually use TC (Thermal Cycling) test to partially replace the TV test because TV test is more expensive. However, due to the air convection, TC test is different from TV test in heat transfer characteristics, which may be over-test or short-test in TC test. This paper aims to explore the applicability of different nanosatellites in TC/TV test. Using rule number analysis method, Heat Transfer model in vacuum and ambient environment has been built to analyse the characteristics of heat transfer under different temperature and characteristic length, and to deliver the recommended limits on using TC test instead of the TV test. The CFD and test methods are applied to verify the rule number analysis above.

The Cross-track Infrared Sounder (CrIS) on board the Suomi National Polar-Orbiting Partnership (S-NPP) Satellite, NOAA-20, and future Joint Polar Satellite System (JPSS) satellites J2 to J4 provide the hyperspectral infrared observations for profiling atmospheric temperature, moisture and greenhouse gases globally. In this study, the J2 CrIS prelaunch spectral calibration parameters including detector focal plane geometry (FOV size and offset angles, referred to as instrument line shape (ILS)) and Neon lamp effective wavelength are derived from the thermal vacuum (TVAC) gas cell test data. These derived parameters are evaluated using the spectral cross-correlation method by comparing with the simulated spectral lines from the line-by-line radiative transfer model (LBLRTM).

In recent years, reality has emerged as a key technology for improving and streamlining design, programming, manufacturing and training processes. Based on experiences in the fields of space robotics, industrial manufacturing, multi-physics and prototyping, virtual are currently being designed and implemented. Experience now shows that testbeds also have a high potential to serve as the foundation for e-Robotics as robotics developed to become a computationally intensive science that is best carried out in a collaborative environment and preferably in a highly distributed network environment. So far, the robotics community does not appear to already appreciate the advantages of principles of e-Science being applied to robotics, probably because most robotics applications so far are being built bottom up and are still mostly dedicated to specific problems, e.g. in the domain of controller development. But when it comes to building larger applications like e.g. a space robot mission, planetary exploration rovers, au-tonomous working machines, advanced vehicle assistant systems etc., the demand for highly realistic and collaborative simulation environment increases. The developments of virtual to be presented in this paper describe the symbiotic integration of advanced simulation techniques with distributed computing ca-pabilities and advanced man-machine interaction metaphors to pave the way for e-Robotics with applications not only in space but also in factories, on construction sites - and in the woods.

In this paper, specifications and capabilities of a newly built thermal vacuum chamber used for the system level thermal vacuum tests is presented. Thanks to the thermoregulation system, it is also possible to perform both subsystem and equipment level thermal vacuum tests. Vacuum condition of space is generated by primary and secondary vacuum pumping systems. Simulation of deep space temperature is achieved by black coated thermal shrouds fed by liquid or gaseous nitrogen. External fluxes which are solar, earth and albedo are simulated by radiative and conductive means in the chamber. All of the operations are managed and controlled by custom built user friendly software. The test data gathered from the sensors mounted on the test item accumulated by a data acquisition system. Two successful thermal vacuum test campaigns performed in the summer of 2012 has shown that it is a reliable test system for the thermal vacuum testing at system, subsystem and equipment level.

Multiple loop heat pipe is a high-functional thermal transport device. This work was conducted to confirm the working performance of Multiple loop heat pipe under thermal vacuum ambience with the working fluid ammonia. Asmall multiple loop heat pipe with two evaporators and two ra- diators was designed and fabricated. Then thermal vacuum test was conducted. The heaters were fasten on both evaporators, both radiators, both compensation chambers. In the case that both evaporators were heated, the multiple loop heat pipe can transport 120/120 W for 1.5 m, in the case that only one evaporator was heated, evaporator 1 can transport 80 W for 1.5 m, while eva- porator 2 can transport 120 W for 1.5 m. Two flow regulators were installed near the confluence of liquid line to prevent uncondensed vapor penetrating into returning liquid when the tempera- ture difference exists between two radiators. In the case that the heat load at both evaporators were 40/40 W and one radiator was heated, the flow regulator1 can tolerate the 160 W of heat load which was supplied to radiator1 while the flow regulator2 can tolerate the 100 W of heat load which was supplied to radiator2. To demonstrate the multiple loop heat pipe’s startup behavior at lowheat load, each of the compensation chamber was preheated to change the initial distribution of liquid and vapor in the evaporator and compensation chamber, in the result, each evaporator can start up at 5W through preheating.

A new deployable solar array system named RAPDAR™ (Roll-out and Passively Deployed Array) is being developed using elastic memory composite (EMC) booms and Thin-Film Photovoltaic (TFPV) cells. For stowage, the 2.5” diameter EMC booms are heated, flattened, and rolled (along with the TFPV array) around a 4” tip spool. The booms maintain their stowed geometry until heated by radiant solar energy, then deploy with the TFPV array in a controlled fashion by unrolling away from the spacecraft. Passive deployment of EMC booms was demonstrated in a thermal vacuum test, using infrared lamps to simulate the sun with the chamber walls cooled with liquid nitrogen. The booms were covered with MLI to increase their temperature during deployment and decrease their temperature during operation of the solar array. The 8’-long booms were cantilevered horizontally in the chamber with no offloading or secondary support and self-deployed at a rate of 3 inches per minute. A full-scale RAPDAR™ flight experiment is also under development that will use 20’-long booms to deploy an array 5’ wide, which produces roughly 800 watts of power. A full-scale engineering unit of this experimental array will be deployed in a solar-thermal vacuum chamber in April of 2007. This paper presents all testing to validate the solar-heated self-deployment of the booms, including in-air testing, thermal-vacuum testing, repeatability testing, and qualification deployments. The test plan for the full-scale solar simulator deployment is also presented.

The Nuclear Spectroscopic Telescope Array (NuSTAR) observatory was successfully launched on June 13 th , 2012 from a Pegasus XL launch vehicle into a low earth orbit (LEO). The NuSTAR mission is currently in the operational science phase, imaging the sky in the high energy X-ray spectrum. The spacecraft bus was designed, integrated, and tested by Orbital Sciences Corporation (Orbital). The Instrument was developed by teams from California Institute of Technology (Caltech), Jet Propulsion Laboratory (JPL), University of California – Berkeley (UCB), and other partners. This paper discusses the bus thermal control system (TCS) design, testing, and on-orbit performance. The NuSTAR TCS design was especially challenging due to the 10meter deployable instrument mast and the spacecraft requirement to point at any target in the sky. The full sky requirement leads to the requirement that the spacecraft withstand being exposed to the thermal environments in dramatically different ways. In some attitudes all of the bus radiators are shielded from solar flux, and in others the radiators can be exposed to full sun. A detailed thermal distortion analysis, which included a sensitivity analysis to many variables, was completed on the bus and instrument structure. The NuSTAR spacecraft utilizes four star trackers, each of which has its own unique thermal control challenges due to their different locations on the spacecraft. One of the star trackers required an innovative thermal control system due to its location inside the spacecraft bus cavity, which has a radiative environment temperature higher than the star tracker operational temperature limit. System level thermal vacuum (TVAC) and thermal balance testing presented a unique challenge because NuSTAR was in the instrument stowed configuration. During TVAC testing unique non-intrusive techniques were used to simulate some component dissipations, and special procedures were developed and utilized for difficult to verify heaters. The on-orbit thermal performance of the bus TCS has been excellent. The thermal model was correlated to on-orbit data to allow better prediction of thermal performance over the entire mission. The final analysis of the NuSTAR bus shows it will meet the TCS requirements for the 2-year mission as planned.

A distributed object is a reusable, self-contained piece of software that cooperates with other objects on the same machine or across a network in a plug-and-play fashion via a well-defined interface. The Numerical Propulsion System Simulation (NPSS) attempts to provide a collaborative aircraft engine design and simulation environment based on this concept. Many scientific applications in aerodynamics and solid mechanics are written in Fortran. Refitting these legacy Fortran codes with distributed objects can increase the code reusability. The remote variable scheme provided in NPSS helps programmers easily migrate the Fortran codes towards a client-server platform. This scheme gives the client the capability of accessing the variables at the server site. Through the operator overloading features in C++, remote variables can be used in much the same way as traditional variables. The remote variable scheme adopts the lazy update approach and the prefetch method. Preliminary performance evaluation shows that communication overhead can be greatly reduced.

A scalable virtual environment for large scale scientific data analysis

Environmental testing of spacecraft flight hardware is performed to detect design and workmanship defects and verify mission requirements prior to launch. At the space vehicle level of assembly, the thermal vacuum test simulates an environment particularly well-suited for verifying mission performance requirements. In this paper, the test objectives of the space vehicle thermal vacuum test are reviewed and an assessment is made of the effectiveness and value of the test. Recent thermal vacuum test data is used to determine how the uniqueness of the thermal vacuum test environment achieves test purposes and how the test ensures mission assurance for space vehicles.

A machine learning–based high volume sensor classification approach using clustering, statistical analysis, and screening on thermal vacuum test data in a medium-sized satellite

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.

Application of the generalized two-body radiant interaction characteristic for estimation ofthe thermal vacuum chamber temperature effect on a space vehicle during thermal vacuum testing

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.