JPL spacecraft sterilization technology program - A status report

Facility description and procedures for heat and ethylene oxide sterilization of spacecraft instrumentation, components, and material

Nonterrestrial utilization of materials: Automated space manufacturing facility

Four areas related to the nonterrestrial use of materials are included: (1) material resources needed for feedstock in an orbital manufacturing facility, (2) required initial components of a nonterrestrial manufacturing facility, (3) growth and productive capability of such a facility, and (4) automation and robotics requirements of the facility

Study of tooling concepts for manufacturing operations in space Final report

Mechanical linkage device for manufacturing operations with orbital workshop

Research and Technology

Langley Research Center is engaged in the basic an applied research necessary for the advancement of aeronautics and space flight, generating advanced concepts for the accomplishment of related national goals, and provding research advice, technological support, and assistance to other NASA installations, other government agencies, and industry. Highlights of major accomplishments and applications are presented

Exploration of the use of digital micromirror devices for highly multiplexed spectroscopy applications in astronomy

Highly multiplexed spectroscopic capabilities are critical to future astronomy space missions. Such capabilities enable large samples of spectral data to be collected in an efficient manner. The individual mirrors of a Digital Micromirror Device (DMD) can serve as slits in a multi-object spectrograph (MOS). This work explores several areas vital to the inclusion of DMDs in future astronomy missions: space qualification, optical performance, and the implementation of Hadamard Transform Spectral Imaging (HTSI). While DMDs were not designed for space, this work reports on testing that demonstrates that the devices can withstand the environmental conditions of a space mission. The optical properties of a DMD ultimately drive the wavelength range and quality of spectral data obtained from a DMD-based MOS. We have characterized the reflectance and contrast ratio of various DMDs from near ultra-violet through visible wavelengths and discuss the results. This work also discusses efforts in expanding the spectral sensitivity of DMDs. Maximizing spectral information over a spatial field of view (FoV) on the sky is highly desirable. In the multi-object spectroscopy mode, individual DMD micromirrors are selected to generate a sparse sample of spectra at individual locations. Additionally, a DMD can be used for integral field spectroscopy (IFS) by forming a long slit from a line of micromirrors, which is then altered to effectively scan across the FoV. In this work we evaluate an alternative technique, HTSI. HTSI has the advantage of a gain in signal-to-noise ratio (SNR) as compared to direct measurements with a long slit, when the observed signals are not photon-noise dominated. We have simulated the performance of HTSI with a DMD-based MOS to identify the limitations of the technique and scenarios where it is most advantageous. With both MOS and IFS capabilities, a DMD-based instrument is a versatile asset fit for a variety of astronomy missions

A Low-Power Optoelectronic Characterizer for CubeSat: LOCC and III-V Nitride Based LEDs

III-V semiconductor materials exhibit robustness and natural hardness when exposed to ionizing radiation and temperature swings. With these characteristics in mind, III-V Nitride Light Emitting Diodes (LEDs) are ideal devices for space-based applications and missions. The effects of ionizing radiation on optoelectronic devices comprised of III-V materials have been studied, but results have been obtained through experiments performed in terrestrial laboratories. While these laboratory tests may lend insight into device lifetimes, performance degradation, etc., they are no substitute for similar measurements and characterization performed in space.;Interest in small satellite applications have grown over the past decade. These solutions range from Earth imaging to communication networks. Small satellites provide a unique opportunity to gain an understanding of the reliability and operational characteristics of III-V based materials and other semiconductor devices while exposed to the environment of space. To meet the constraints of the small satellite, a Low-powered Optoelectronic Characterizer for CubeSat (LOCC) has been developed in PC/104 form, measuring 3.6 by 3.8 inches. LOCC performs current-voltage and electroluminescent measurements of LEDs while in space. The LOCC system is designed using low-power integrated circuits that can supply over 100 mA of current to LEDs while maintaining low power of 3.2W under operation.;This thesis presents the design, implementation, and control of the LOCC system. This includes system block diagrams, printed circuit board layouts, interfacing, firmware, and software. Additionally, the resulting current-voltage measurements, required wattage, and required data storage are presented to illustrate functionality. This instrumentation enables the study of optoelectronic devices in space, allowing future research to focus on producing radiation hard light emitting devices that can operate in environments with reduced shielding against ionizing radiation while maintaining device reliability

Imaging Payload System Design for Nano-Satellite Applications in Low-Earth Orbit

학위논문 (석사)-- 서울대학교 대학원 : 기계항공공학부, 2016. 8. 정인석.This study presents a complete hardware and software design of an imaging payload, InterFace Camera (IFCAM), for the CubeSat standard. The payload is intended to take earth images from 350km Low Earth Orbit and is a non-critical mission payload for Seoul National Universitys first CubeSat, SNUSAT-1. The camera has a MT9D111 2M pixel image sensor with a STM32F429ZI processor running at 180MHz, 4GB SD card for data storage, 8MB flash memory and 8MB expandable SDRAM. The camera is configured to take images in 160x120 pixels because of the limited downlink budget given to the imaging payload. Edmund Optics fixed focus M12x0.5 lens with 650mm IR-Cut off filter and the Ground Sampling Distance (GSD) is calculated to about 650km. The Ground Swath, with the optics, is 570x400km which cover South Korea. The payload underwent Random Vibration and Vacuum Testing and showed no issues in the design. Furthermore, the design proposed is modular and has been used for designing the On-Board Computer for SNUSAT-1.1. Introduction 1 1.1. Previous Work 1 1.2. Requirement 3 1.3. Imaging Payload Structure 3 1.4. Remote Sensing 4 1.5. Space Environment 5 2. Hardware 9 2.1. Design Constraint 9 2.2. Hardware Selection 10 2.3. Hardware Design 18 2.4. Prototyping. 19 2.5. Flight Model PCB Design 30 3. Software 31 3.1. Introduction 31 3.2. Software Development Tools/Resources 31 3.3. Software Description 35 4. Verification 40 4.1. Field Testing 40 4.2. Vibration Testing 42 4.3. Vacuum Testing 59 5. Improvements 63 5.1. Software 63 5.2. Hardware 64 6. Conclusion 66 6.1. Requirement Review 66 6.2. Overview 67 References 69 Abstract in Korean 71 Appendices 72Maste

Apollo experience report guidance and control systems: Primary guidance, navigation, and control system development

The primary guidance, navigation, and control systems for both the lunar module and the command module are described. Development of the Apollo primary guidance systems is traced from adaptation of the Polaris Mark II system through evolution from Block I to Block II configurations; the discussion includes design concepts used, test and qualification programs performed, and major problems encountered. The major subsystems (inertial, computer, and optical) are covered. Separate sections on the inertial components (gyroscopes and accelerometers) are presented because these components represent a major contribution to the success of the primary guidance, navigation, and control system

Thermal Modelling and Experiments for Small Satellites

There has been an increasing interest in CubeSat missions due to its small size, low cost and flexibility to accommodate different payloads. New missions with highly temperature sensitive payloads, increased power dissipation (by continuous miniaturization of electronic components and systems) and reduced radiating surfaces lead the thermal loads issues into a bigger challenge. One of the causes of failure in a satellite in space is the temperature peaks suffered during a full orbital cycle. Therefore, proper thermal control system design and test should be performed to guarantee the reliability of a spacecraft prior to launch. 3-AMADEUS is a unity CubeSat currently being developed in a partnership between CEiiA and UBI. The purpose of this mission is to demonstrate that a attitude determiner and control system exclusively magnetic is able to provide a three axis orbital attitude for the nanosatellites. The present work aims to perform thermal analysis to 3-AMADEUS CubeSat in order to ensure its survival as soon as it is placed in orbit. Therefore, it is required the understand the main heat transfer processes within a satellite, conduction and radiation, in order to validate the current methodologies used for thermal analysis. Hence, with the purpose of developing thermal models with higher reliability, two experiments were devised to be performed in a vacuum environment. The first experimental test consists in a study of heat exchange between two aluminum plates through radiation, using an infrared lamp as a heat source. Three distance configurations between plates and two lamp types were tested to comparison. This would emulate, for example, the heat transmission between different components within the satellite. Regarding the conduction experiment, most nano and micro satellites are composed of stacked PCBs, held together by spacers and rods and linked to the main structure. This is the primary mean to conduct the heat from the different components to the external radiating surfaces. A high thermal resistance is associated with the interface between the PCB and the spacers, which is an unknown parameter with a high impact on the thermal analysis. Therefore, a second experiment is carried out to study thermal contact resistance (or conductance) between them. In parallel, finite element software (MSC Nastran) is used to carry out a numerical study of the same experiments. The temperature distribution results of both numerical and experimental solutions were then compared, and the results were discussed. It was concluded that the results obtained in both experiments, in general, presented a good agreement. Finally, with the results obtained in the numerical simulations and using the validated methodology, a steady state thermal analysis was performed to 3-AMADEUSTem havido um crescente interesse nas missões e na obtenção de dados através da utilização de CubeSats. Estes, devido à sua dimensão e baixo custo têm uma grande flexibilidade em acomodar diferentes cargas úteis. No entanto, novas missões com cargas úteis e componentes altamente sensíveis à temperatura, o aumento da dissipação de energia (pela miniaturização de componentes e sistemas eletrónicos) e superfícies irradiadoras reduzidas levam a possíveis problemas térmicos. Uma das causas para a falha de um satélite em órbita são os picos de temperatura sofridos durante um ciclo orbital completo. Portanto, o projeto e o teste adequados do sistema de controlo térmico devem ser realizados, de forma a garantir a fiabilidade do satélite antes do seu lançamento de modo a reduzir a possibilidade de falha. O 3-AMADEUS é um CubeSat de uma unidade que está atualmente a ser desenvolvido numa parceria entre o CEiiA e a UBI. O propósito desta missão é demonstrar que um sistema de determinação e controlo de atitude exclusivamente magnético, pode ser capaz de fornecer atitude orbital de três eixos para os nanossatélites. O presente trabalho tem como objetivo efetuar análises térmicas ao 3-AMADEUS CubeSat para confirmar a sua sobrevivência assim que for colocado em órbita. Para isso, é necessário analisar os principais processos de transferência de calor num satélite, condução e radiação, de forma a validar as metodologias atualmente utilizadas para as análises térmicas. Assim, com o objetivo de desenvolver modelos térmicos com maior fiabilidade, foram realizadas duas experiências em vácuo. O primeiro teste experimental consiste num estudo da troca de calor entre duas placas de alumínio através de radiação, usando uma lâmpada de infravermelhos como fonte de calor. Foram testadas três configurações de distância entre as placas e dois tipos de lâmpadas para comparação. Este teste simularia, por exemplo, a transmissão de calor entre diferentes componentes dentro do satélite. Relativamente à condução, a maioria dos nano e microssatélites são compostos de PCBs empilhadas, mantidas juntas por espaçadores e varões roscados, conectados à estrutura principal. Esta é a principal forma de conduzir calor dos componentes para as superfícies irradiadoras. Associada à interface entre a PCB e os espaçadores, existe uma resistência térmica que é um parâmetro desconhecido com grande impacto nas análises térmicas. Desta forma, foi realizado uma segunda experiência para estudar a resistência térmica de contacto (ou condutância) entre uma PCB e espaçadores. Paralelamente, o software de elementos finitos (MSC Nastran) é usado para realizar um estudo numérico das mesmas experiências. Os resultados da distribuição de temperatura das soluções numéricas e experimentais foram então comparados e os resultados foram discutidos. Finalmente, com os resultados obtidos durante os testes foi realizada uma análise térmica em estado estacionário ao 3-AMADEUS CubeSat