The document summarizes a preliminary design review for a search and rescue satellite constellation. Key points:
1) The constellation would consist of 48 nano-satellites in 6 orbital planes to locate persons in distress within 15 minutes with 1 km accuracy.
2) Each satellite would have a mass of 3.11 kg, use passive thermal control, active 3-axis attitude control, a warm gas propulsion system, and communicate via two dipole antennas.
3) The design was updated since the preliminary design review, with the total satellite mass reduced to 2.3 kg and the delta-V budget reduced from 10.31 m/s to 7.94 m/s.
The LEISA Atmospheric Corrector (AC) on EO1 was selected in 1993 for a planned Pluto mission and later chosen for the Earth Observing-1 (EO-1) mission in 1996. It is a hyperspectral imager that collects moderate spatial and spectral resolution data to correct atmospheric effects in high spatial resolution multispectral images. It has contributed to EO-1 through atmospheric correction of ALI and Landsat-7 images and studies of spatial resolution degradation by comparing to Hyperion.
A fast-paced tutorial on satellite image geometry.
Mono & stereo collection geometry.
Effects of collection geometry on image quality, perspective and accuracy.
RPC & Physical Camera Models
Geometry of scan-oriented, map-oriented, orthorectified, and stereo image products
This tutorial for producers and users of satellite imagery provides a common vocabulary and understanding of collection and product geometry and effects.
The mission aims to map space debris in low Earth orbit between 1,000-3,000 km using 6 small satellites. Each satellite will use an infrared camera to image debris and calculate its orbit. The satellites will be placed in 3 evenly spaced orbital planes by a Delta II rocket and slowly lower their orbits over 30 days to map the entire region. Their design emphasizes modularity for low cost and mass, using commercial off-the-shelf components, with a focus on thermal control, power, communications and orbital maneuvering systems to complete the debris mapping mission.
This document describes the CubeSat High Impulse Propulsion System (CHIPS) being developed by CU Aerospace and VACCO Industries. CHIPS integrates primary propulsion, attitude control, and propellant storage into a single module compatible with CubeSats. It uses a micro-resistojet thruster and cold gas attitude control system. Testing of a prototype showed the resistojet can provide up to 563 N-s of total impulse using R134a propellant. The document discusses propulsion options for CubeSats and why electrothermal propulsion in the 70-400 second specific impulse range is best suited for enabling rapid orbital maneuvers within a day.
This document discusses the concept of an X-ray interferometer called MAXIM that could achieve micro-arcsecond resolution. It would consist of an optics spacecraft holding multiple flat mirrors in formation with a detector spacecraft to form interference patterns. The goal is to image phenomena like black hole accretion disks and supernovae with much higher resolution than current telescopes. A pathfinder mission is proposed with 100 microarcsecond resolution using two spacecraft separated by 1.4 meters as a technology demonstration.
The KMEC mission involves sending two spacecraft to Saturn over 6 years to study cosmic dust, ultraviolet imaging, and space recognition between the payloads. Each spacecraft is octagonal and 6m tall, made of aluminum. The 100kg payload includes dust, UV, and ranging instruments. A chemical propulsion system will perform orbital maneuvers. Power comes from an RTG and backup battery. Thermal control uses an RTG and radiator. The spacecraft structure is sized to withstand launch stresses and the environment at Saturn.
The document discusses the design of magnetic sail (magsail) systems for spacecraft propulsion. It describes a proposed demonstrator magsail with a 200m radius and 25.7kg mass, and an operational magsail with 20,000m radius and 7,060 metric tonne mass. The operational design could accelerate at 0.003185 m/s^2 and deliver over 100,000kg payloads to Mars or Saturn. Future advances in superconductors could enable magsails to deliver payloads of over 400,000kg to Jupiter and millions of kilograms to the outer planets.
The document discusses strategies for optimizing data collection for the proposed GEO-CAPE mission through intelligent observation studies conducted by NASA's Goddard Space Flight Center team. Key findings include developing an observation operations simulator to examine different instrument and scheduling options, identifying cloud detection algorithms including the value of shortwave infrared bands, and establishing the feasibility of approaches like onboard cloud detection to reduce data handling costs. The simulator could be used to test different "what if" scenarios incorporating actual cloud forecast data and characterize instruments based on scene footprints. Possible future work involves integrating a scheduler with the simulator for live simulations, updating the tool with user feedback, and further analyzing instruments and cloud detection capabilities.
The LEISA Atmospheric Corrector (AC) on EO1 was selected in 1993 for a planned Pluto mission and later chosen for the Earth Observing-1 (EO-1) mission in 1996. It is a hyperspectral imager that collects moderate spatial and spectral resolution data to correct atmospheric effects in high spatial resolution multispectral images. It has contributed to EO-1 through atmospheric correction of ALI and Landsat-7 images and studies of spatial resolution degradation by comparing to Hyperion.
A fast-paced tutorial on satellite image geometry.
Mono & stereo collection geometry.
Effects of collection geometry on image quality, perspective and accuracy.
RPC & Physical Camera Models
Geometry of scan-oriented, map-oriented, orthorectified, and stereo image products
This tutorial for producers and users of satellite imagery provides a common vocabulary and understanding of collection and product geometry and effects.
The mission aims to map space debris in low Earth orbit between 1,000-3,000 km using 6 small satellites. Each satellite will use an infrared camera to image debris and calculate its orbit. The satellites will be placed in 3 evenly spaced orbital planes by a Delta II rocket and slowly lower their orbits over 30 days to map the entire region. Their design emphasizes modularity for low cost and mass, using commercial off-the-shelf components, with a focus on thermal control, power, communications and orbital maneuvering systems to complete the debris mapping mission.
This document describes the CubeSat High Impulse Propulsion System (CHIPS) being developed by CU Aerospace and VACCO Industries. CHIPS integrates primary propulsion, attitude control, and propellant storage into a single module compatible with CubeSats. It uses a micro-resistojet thruster and cold gas attitude control system. Testing of a prototype showed the resistojet can provide up to 563 N-s of total impulse using R134a propellant. The document discusses propulsion options for CubeSats and why electrothermal propulsion in the 70-400 second specific impulse range is best suited for enabling rapid orbital maneuvers within a day.
This document discusses the concept of an X-ray interferometer called MAXIM that could achieve micro-arcsecond resolution. It would consist of an optics spacecraft holding multiple flat mirrors in formation with a detector spacecraft to form interference patterns. The goal is to image phenomena like black hole accretion disks and supernovae with much higher resolution than current telescopes. A pathfinder mission is proposed with 100 microarcsecond resolution using two spacecraft separated by 1.4 meters as a technology demonstration.
The KMEC mission involves sending two spacecraft to Saturn over 6 years to study cosmic dust, ultraviolet imaging, and space recognition between the payloads. Each spacecraft is octagonal and 6m tall, made of aluminum. The 100kg payload includes dust, UV, and ranging instruments. A chemical propulsion system will perform orbital maneuvers. Power comes from an RTG and backup battery. Thermal control uses an RTG and radiator. The spacecraft structure is sized to withstand launch stresses and the environment at Saturn.
The document discusses the design of magnetic sail (magsail) systems for spacecraft propulsion. It describes a proposed demonstrator magsail with a 200m radius and 25.7kg mass, and an operational magsail with 20,000m radius and 7,060 metric tonne mass. The operational design could accelerate at 0.003185 m/s^2 and deliver over 100,000kg payloads to Mars or Saturn. Future advances in superconductors could enable magsails to deliver payloads of over 400,000kg to Jupiter and millions of kilograms to the outer planets.
The document discusses strategies for optimizing data collection for the proposed GEO-CAPE mission through intelligent observation studies conducted by NASA's Goddard Space Flight Center team. Key findings include developing an observation operations simulator to examine different instrument and scheduling options, identifying cloud detection algorithms including the value of shortwave infrared bands, and establishing the feasibility of approaches like onboard cloud detection to reduce data handling costs. The simulator could be used to test different "what if" scenarios incorporating actual cloud forecast data and characterize instruments based on scene footprints. Possible future work involves integrating a scheduler with the simulator for live simulations, updating the tool with user feedback, and further analyzing instruments and cloud detection capabilities.
Performance Evaluation of SAR Image Reconstruction on CPUs and GPUsFisnik Kraja
This document summarizes the performance evaluation of synthetic aperture radar (SAR) image reconstruction on CPUs and GPUs. It describes porting the SAR application to NVIDIA CUDA GPUs and compares the performance results on CPUs with 8 and 16 threads and single and dual GPU configurations. Using GPUs provided better performance than CPUs, especially for large-scale images. A heterogeneous CPU+GPU approach improved performance over the CPU-only version by reducing data transfers between the processors. The best results were achieved with a pipelined dual-GPU implementation that reconstructed separate images in parallel to minimize data movement.
GPS measurements are affected by random and systematic errors that impact accuracy. Errors originate from satellites, signal propagation through the atmosphere, and receivers. Satellite errors include orbital inaccuracies and clock errors. Signals pass through the ionosphere, a dispersive layer, causing refraction dependent on frequency. Receivers have clock errors. Combining dual-frequency measurements eliminates ionospheric delay, the main source of error.
The minor project involved studying wind resource assessment methodology and conducting an assessment of a site in Tamil Nadu, India from May to July 2010. Key activities included visiting a wind turbine test station, analyzing wind speed and direction data, estimating energy output, and developing wind rose and frequency distribution charts. Exercises were performed to calculate Weibull parameters, estimate long-term wind speeds using reanalysis data, and develop a joint frequency distribution of wind speed and direction.
Satellite RF Communications and Onboard Processing Course SamplerJim Jenkins
Successful systems engineering requires a broad understanding of the important principles of modern satellite communications and onboard data processing. This course covers both theory and practice, with emphasis on the important system engineering principles, tradeoffs, and rules of thumb. The latest technologies are covered, including those needed for constellations of satellites.
This course is recommended for engineers and scientists interested in acquiring an understanding of satellite communications, command and telemetry, onboard computing, and tracking. Each participant will receive a complete set of notes.
This document proposes a mission to send a probe to 1000 AU within 50 years to explore the very local interstellar medium. It would use a gravity assist at Jupiter to eliminate angular momentum, fall into 4 solar radii from the sun for a high-speed propulsion burn, and reach speeds of 20 AU/year. Required technologies include high-Isp propulsion, thermal shields, long-life electronics, and autonomous operation. The proposed concept uses solar thermal propulsion and liquid hydrogen, carried on an Atlas V launch vehicle. The probe would perform in situ measurements of the interstellar medium and escape the heliosphere to study boundary regions.
The document discusses the generation of high resolution digital surface models (DSMs) from ALOS PRISM stereo imagery and efforts to create seamless DSM mosaics at the global scale. It provides an overview of the software used to generate DSMs from PRISM triplet images and details on archiving over 5,000 processed scene DSMs. It then describes the process used to create 1° x 1° tile mosaics, including correcting for biases between relative and absolute DSMs. Examples of mosaics created for Japan and Bhutan are shown and their contributions to studies on glacial lake outburst floods in Bhutan are discussed. The goal is to complete a global 10m resolution DSM mosaic.
This document summarizes the results of the first bistatic SAR experiment using JAXA's PALSAR satellite radar and NEC's Pi-SAR-L airborne radar. The bistatic image showed higher signal-to-noise ratio, lower azimuth resolution, and ability to detect different targets than the monostatic images. Analysis found the bistatic image had better visibility of small features due to its shorter receiver-target distance. Its lower resolution was likely due to Pi-SAR-L not having synchronization functions. Some targets appeared differently between images from their varying incidence angles and signal levels.
Space Radiation & It's Effects On Space Systems & Astronauts Technical Traini...Jim Jenkins
This course is designed for technical and management personnel who wish to gain an understanding of the fundamentals and the effects of space radiation on space systems and astronauts. The radiation environment imposes strict design requirements on many space systems and is the primary limitation to human exploration outside of the Earth's magnetosphere. The course specifically addresses issues of relevance and concern for participants who expect to plan, design, build, integrate, test, launch, operate or manage spacecraft and spacecraft subsystems for robotic or crewed missions. The primary goal is to assist attendees in attainment of their professional potential by providing them with a basic understanding of the interaction of radiation with non-biological and biological materials, the radiation environment, and the tools available to simulate and evaluate the effects of radiation on materials, circuits, and humans
The document discusses grid computing at CERN for the Large Hadron Collider experiment. CERN operates a worldwide computing grid with tiered levels to handle the massive computing and storage needs. Tier 0 is at CERN for data acquisition and distribution. Tier 1 centers have large storage and do data analysis. Tier 2 centers participate in simulation and analysis. The LHC generates 40 million collision events per second that are filtered and recorded, resulting in 15 petabytes of data per year. The computing grid is necessary to process and store this huge volume of data across the distributed centers.
The document summarizes the results of a demonstration test for the GEO Atmospheric Sounder (GAS) concept. The test met key requirements, including image resolution <10 mrad, beam efficiency >95%, relative temperature accuracy <2K, and polarization isolation >10 dB. It used a rotating interferometer with 21 dual-polarization receivers to image calibration sources and the sun. The results validated the interferometer concept and showed agreement with theoretical models.
MODELLING AND ATMOSPHERIC ERRORS IN GPS SIGNAL PROPAGATIONSHADABANSARI57
A BRIEF INTRODUCTION TO IONOSPHERIC AND TROPOSPHERIC ERRORS IN GPS SIGNAL PROPAGATION
DERIVATION OF FORMULA WITH THE BASIC LEVEL.
BASIC UNDERSTANDING OF FIGURE..
FOLLOW THE REFERENCE BOOK FOR MORE DETAIL.
The document summarizes the Level 1 processing pipeline (L1PP) used to process data from the Soil Moisture and Ocean Salinity (SMOS) mission. The L1PP takes raw instrument data and produces calibrated brightness temperature images at three levels - L1a, L1b, and L1c. Key algorithms described include calibration steps, removal of interference sources like sun and moon, image reconstruction, and geolocation. Validation results showed the L1PP algorithms could produce images with well-defined coastlines and improved radio frequency interference mitigation.
Advanced Satellite Communications Systems Technical Training Course SamplerJim Jenkins
This three-day ATI professional development course, Advanced Satellite Communications Systems, covers all the technology of
advanced satellite communications as well as the
principles behind current state-of-the-art satellite
communications equipment. System
design approaches discussed include Big
LEO, Little LEO, DTH, Direct Radio, VSAT,
commercial satellite imaging, multimedia satellite
services, and IP networking over satellite. There will
be course exercises and practice each day
20121205 open stack_accelerating_science_v3Tim Bell
CERN is a large particle physics laboratory located between Geneva and France that is seeking answers to fundamental questions about the universe. It operates the Large Hadron Collider (LHC) and collects massive amounts of data that are stored and analyzed using a worldwide computing grid. CERN is adopting open source tools like OpenStack for its computing infrastructure to better manage scaling its systems with limited staff increases. It is contributing to open source communities and training staff on these tools to gain valuable skills applicable outside of CERN.
Astrometric Data Processing in Amazon EC2
Proof of Concept to evaluate the Tecnical and financial Feasibility of using Amazon EC2 to run a scientific data processing applications for the European Space Agency
This document discusses a proposed mission concept to send a probe to 1000 AU within 50 years using currently feasible technologies. Key elements include:
1) Using a solar gravity assist at Jupiter to eliminate angular momentum, then falling to within 4 solar radii of the Sun to leverage the high speeds for an escape trajectory.
2) The probe would use a high-Isp propulsion system like solar thermal or nuclear thermal to accelerate during a 15-minute perihelion maneuver for solar system escape.
3) Enabling technologies discussed include high-temperature carbon-carbon shields, efficient radioisotope power, and laser optical communications. Follow-on studies are proposed to further develop the concepts.
The document summarizes early experiences using the Summit supercomputer at Oak Ridge National Laboratory. Summit is the world's fastest supercomputer and has been used by several early science projects. Two example applications, GTC and CoMet, have achieved good scaling and performance on Summit. Some initial issues were encountered but addressed. Overall, Summit is a very powerful system but continued software improvements are needed to optimize applications for its complex hardware architecture.
This document provides an overview of the Global Change Observation Mission (GCOM) and its first satellite, GCOM-W1. GCOM aims to continue long-term Earth observations to monitor climate change, the carbon cycle, and other variables. GCOM-W1 will carry the Advanced Microwave Scanning Radiometer 2 (AMSR2) instrument and join the A-Train constellation in 2012. AMSR2 improves on the previous AMSR-E instrument with an enhanced calibration system to provide more accurate measurements of Earth's water and energy cycles over 5 years. Future GCOM satellites will continue these essential climate observations through 2022.
Performance Evaluation of SAR Image Reconstruction on CPUs and GPUsFisnik Kraja
This document summarizes the performance evaluation of synthetic aperture radar (SAR) image reconstruction on CPUs and GPUs. It describes porting the SAR application to NVIDIA CUDA GPUs and compares the performance results on CPUs with 8 and 16 threads and single and dual GPU configurations. Using GPUs provided better performance than CPUs, especially for large-scale images. A heterogeneous CPU+GPU approach improved performance over the CPU-only version by reducing data transfers between the processors. The best results were achieved with a pipelined dual-GPU implementation that reconstructed separate images in parallel to minimize data movement.
GPS measurements are affected by random and systematic errors that impact accuracy. Errors originate from satellites, signal propagation through the atmosphere, and receivers. Satellite errors include orbital inaccuracies and clock errors. Signals pass through the ionosphere, a dispersive layer, causing refraction dependent on frequency. Receivers have clock errors. Combining dual-frequency measurements eliminates ionospheric delay, the main source of error.
The minor project involved studying wind resource assessment methodology and conducting an assessment of a site in Tamil Nadu, India from May to July 2010. Key activities included visiting a wind turbine test station, analyzing wind speed and direction data, estimating energy output, and developing wind rose and frequency distribution charts. Exercises were performed to calculate Weibull parameters, estimate long-term wind speeds using reanalysis data, and develop a joint frequency distribution of wind speed and direction.
Satellite RF Communications and Onboard Processing Course SamplerJim Jenkins
Successful systems engineering requires a broad understanding of the important principles of modern satellite communications and onboard data processing. This course covers both theory and practice, with emphasis on the important system engineering principles, tradeoffs, and rules of thumb. The latest technologies are covered, including those needed for constellations of satellites.
This course is recommended for engineers and scientists interested in acquiring an understanding of satellite communications, command and telemetry, onboard computing, and tracking. Each participant will receive a complete set of notes.
This document proposes a mission to send a probe to 1000 AU within 50 years to explore the very local interstellar medium. It would use a gravity assist at Jupiter to eliminate angular momentum, fall into 4 solar radii from the sun for a high-speed propulsion burn, and reach speeds of 20 AU/year. Required technologies include high-Isp propulsion, thermal shields, long-life electronics, and autonomous operation. The proposed concept uses solar thermal propulsion and liquid hydrogen, carried on an Atlas V launch vehicle. The probe would perform in situ measurements of the interstellar medium and escape the heliosphere to study boundary regions.
The document discusses the generation of high resolution digital surface models (DSMs) from ALOS PRISM stereo imagery and efforts to create seamless DSM mosaics at the global scale. It provides an overview of the software used to generate DSMs from PRISM triplet images and details on archiving over 5,000 processed scene DSMs. It then describes the process used to create 1° x 1° tile mosaics, including correcting for biases between relative and absolute DSMs. Examples of mosaics created for Japan and Bhutan are shown and their contributions to studies on glacial lake outburst floods in Bhutan are discussed. The goal is to complete a global 10m resolution DSM mosaic.
This document summarizes the results of the first bistatic SAR experiment using JAXA's PALSAR satellite radar and NEC's Pi-SAR-L airborne radar. The bistatic image showed higher signal-to-noise ratio, lower azimuth resolution, and ability to detect different targets than the monostatic images. Analysis found the bistatic image had better visibility of small features due to its shorter receiver-target distance. Its lower resolution was likely due to Pi-SAR-L not having synchronization functions. Some targets appeared differently between images from their varying incidence angles and signal levels.
Space Radiation & It's Effects On Space Systems & Astronauts Technical Traini...Jim Jenkins
This course is designed for technical and management personnel who wish to gain an understanding of the fundamentals and the effects of space radiation on space systems and astronauts. The radiation environment imposes strict design requirements on many space systems and is the primary limitation to human exploration outside of the Earth's magnetosphere. The course specifically addresses issues of relevance and concern for participants who expect to plan, design, build, integrate, test, launch, operate or manage spacecraft and spacecraft subsystems for robotic or crewed missions. The primary goal is to assist attendees in attainment of their professional potential by providing them with a basic understanding of the interaction of radiation with non-biological and biological materials, the radiation environment, and the tools available to simulate and evaluate the effects of radiation on materials, circuits, and humans
The document discusses grid computing at CERN for the Large Hadron Collider experiment. CERN operates a worldwide computing grid with tiered levels to handle the massive computing and storage needs. Tier 0 is at CERN for data acquisition and distribution. Tier 1 centers have large storage and do data analysis. Tier 2 centers participate in simulation and analysis. The LHC generates 40 million collision events per second that are filtered and recorded, resulting in 15 petabytes of data per year. The computing grid is necessary to process and store this huge volume of data across the distributed centers.
The document summarizes the results of a demonstration test for the GEO Atmospheric Sounder (GAS) concept. The test met key requirements, including image resolution <10 mrad, beam efficiency >95%, relative temperature accuracy <2K, and polarization isolation >10 dB. It used a rotating interferometer with 21 dual-polarization receivers to image calibration sources and the sun. The results validated the interferometer concept and showed agreement with theoretical models.
MODELLING AND ATMOSPHERIC ERRORS IN GPS SIGNAL PROPAGATIONSHADABANSARI57
A BRIEF INTRODUCTION TO IONOSPHERIC AND TROPOSPHERIC ERRORS IN GPS SIGNAL PROPAGATION
DERIVATION OF FORMULA WITH THE BASIC LEVEL.
BASIC UNDERSTANDING OF FIGURE..
FOLLOW THE REFERENCE BOOK FOR MORE DETAIL.
The document summarizes the Level 1 processing pipeline (L1PP) used to process data from the Soil Moisture and Ocean Salinity (SMOS) mission. The L1PP takes raw instrument data and produces calibrated brightness temperature images at three levels - L1a, L1b, and L1c. Key algorithms described include calibration steps, removal of interference sources like sun and moon, image reconstruction, and geolocation. Validation results showed the L1PP algorithms could produce images with well-defined coastlines and improved radio frequency interference mitigation.
Advanced Satellite Communications Systems Technical Training Course SamplerJim Jenkins
This three-day ATI professional development course, Advanced Satellite Communications Systems, covers all the technology of
advanced satellite communications as well as the
principles behind current state-of-the-art satellite
communications equipment. System
design approaches discussed include Big
LEO, Little LEO, DTH, Direct Radio, VSAT,
commercial satellite imaging, multimedia satellite
services, and IP networking over satellite. There will
be course exercises and practice each day
20121205 open stack_accelerating_science_v3Tim Bell
CERN is a large particle physics laboratory located between Geneva and France that is seeking answers to fundamental questions about the universe. It operates the Large Hadron Collider (LHC) and collects massive amounts of data that are stored and analyzed using a worldwide computing grid. CERN is adopting open source tools like OpenStack for its computing infrastructure to better manage scaling its systems with limited staff increases. It is contributing to open source communities and training staff on these tools to gain valuable skills applicable outside of CERN.
Astrometric Data Processing in Amazon EC2
Proof of Concept to evaluate the Tecnical and financial Feasibility of using Amazon EC2 to run a scientific data processing applications for the European Space Agency
This document discusses a proposed mission concept to send a probe to 1000 AU within 50 years using currently feasible technologies. Key elements include:
1) Using a solar gravity assist at Jupiter to eliminate angular momentum, then falling to within 4 solar radii of the Sun to leverage the high speeds for an escape trajectory.
2) The probe would use a high-Isp propulsion system like solar thermal or nuclear thermal to accelerate during a 15-minute perihelion maneuver for solar system escape.
3) Enabling technologies discussed include high-temperature carbon-carbon shields, efficient radioisotope power, and laser optical communications. Follow-on studies are proposed to further develop the concepts.
The document summarizes early experiences using the Summit supercomputer at Oak Ridge National Laboratory. Summit is the world's fastest supercomputer and has been used by several early science projects. Two example applications, GTC and CoMet, have achieved good scaling and performance on Summit. Some initial issues were encountered but addressed. Overall, Summit is a very powerful system but continued software improvements are needed to optimize applications for its complex hardware architecture.
This document provides an overview of the Global Change Observation Mission (GCOM) and its first satellite, GCOM-W1. GCOM aims to continue long-term Earth observations to monitor climate change, the carbon cycle, and other variables. GCOM-W1 will carry the Advanced Microwave Scanning Radiometer 2 (AMSR2) instrument and join the A-Train constellation in 2012. AMSR2 improves on the previous AMSR-E instrument with an enhanced calibration system to provide more accurate measurements of Earth's water and energy cycles over 5 years. Future GCOM satellites will continue these essential climate observations through 2022.
Mars CubeSat Telecom Relay Constellation_JPL FinalRohan Deshmukh
The document proposes a 4-CubeSat constellation in Mars orbit to augment telecommunications relay capabilities for surface assets. The CubeSats would be deployed from a 2022 Mars Telecom Orbiter into a 350 km circular orbit with 70 degree inclination. Each 6U CubeSat would carry an IRIS transponder, Ka-band reflectarray antenna, and UHF antennas to provide direct-to/from-Earth relay and cross-link relay with orbiters. Analysis showed the design could meet power and propulsion requirements to operate for two Mars years.
The document provides information on the Thailand Earth Observation System (THEOS) Earth observation mission. Some key details include:
- THEOS was developed by EADS Astrium for Thailand's GISTDA space agency to provide optical imagery for applications like cartography, agriculture, and disaster monitoring.
- The THEOS spacecraft was launched in 2008 into a sun-synchronous orbit and carries panchromatic and multispectral cameras.
- GISTDA operates the spacecraft and a ground station in Thailand to control THEOS and receive its imagery.
- Cooperation agreements between GISTDA and Taiwan's NSPO space agency allow them to share ground station resources and provide backup support for each other's missions
Importance of SSPS in SDG and ESG, and importance of antennas in SSPSAdvanced-Concepts-Team
SSPS has benefits for achieving SDGs and ESG goals by providing low emission, sustainable power. Key technologies include large antennas that can precisely direct microwave beams for power transmission. Antennas face challenges like developing arrays of thousands of precisely controlled elements. The presentation outlines a roadmap including technology verification experiments and test satellites to demonstrate SSPS technologies like deployable antennas in space before developing commercial-scale systems in geosynchronous orbit capable of generating several GW of power.
Design and First Results of an UAV-Borne L-Band Radiometer for Multiple Monit...Angelo State University
This document describes the design and initial results of an L-band radiometer mounted on an unmanned aerial vehicle (UAV) for soil moisture monitoring. The radiometer measures antenna temperature with 1.27K resolution. Software processes the raw data, applying calibration and georeferencing to produce soil moisture maps. Initial field tests show the system can distinguish between soil, water and sun glint reflections. The UAV system provides flexibility and high resolution for applications like precision agriculture compared to spaceborne radiometers.
The document summarizes options for re-flying the Orbiting Carbon Observatory (OCO) mission after the original spacecraft was lost during launch in February 2009. It discusses examining service platforms, access to space options, and conducting studies to evaluate rebuilding the OCO instrument and spacecraft ("Carbon Copy") as the lowest risk approach. While awaiting authorization to proceed, the OCO project team is working to reduce implementation risks and collaborate with other missions to be in the best position for an OCO re-flight.
The document summarizes options for re-flying the Orbiting Carbon Observatory (OCO) mission after the original spacecraft was lost during launch in February 2009. It discusses examining service platforms, access to space options, and conducting studies to evaluate rebuilding the OCO instrument and spacecraft ("Carbon Copy") as the lowest risk approach. While awaiting authorization to proceed, the OCO project team is working to reduce implementation risks and collaborate with other missions to be in the best position for an OCO re-flight.
The document summarizes the proto-flight test of the Dual-frequency Precipitation Radar (DPR) for NASA's Global Precipitation Measurement (GPM) mission. The DPR consists of Ku-band and Ka-band radars that will provide accurate 3D precipitation measurements from space. Electrical performance, vibration, and thermal tests have been conducted on each radar. While the proto-flight test was interrupted by an earthquake, it has resumed and will be completed to verify the radars can function as intended in space. The DPR will improve global precipitation observations when launched aboard the GPM core observatory.
The Cassini spacecraft has conducted a complex mission involving multiple orbits of Saturn with numerous targets and risks. It has experienced several anomalies over its journey that required workarounds or redesigning aspects of the mission, such as adjusting orbits to save the Huygens probe mission after a communications issue was discovered. Careful planning was required to develop the intricate tour involving 75 orbits and 150 maneuvers. Ongoing risk assessment and mitigation efforts have helped manage challenges over the long mission.
This document provides a preliminary study for the AROSAT satellite system. It discusses several key requirements including coverage area, resolution capabilities, duty cycle, onboard storage and download rates. It evaluates three potential spacecraft configurations and their impact on drag, solar array effectiveness and risk. Configuration #2 is preferred as it minimizes drag while having a simple solar array design. The document also examines how spacecraft altitude affects optical instrument parameters and the propulsion systems needed to compensate for atmospheric drag at different altitudes. Electric propulsion is recommended to enable lower orbits. Overall architectures are proposed for Configuration #2 that could meet requirements.
The document analyzes polarimetric decomposition of radar data from the NASA UAVSAR instrument to study the Deepwater Horizon oil slick. UAVSAR collected L-band radar data over a large area, including two flight lines passing directly over the main oil slick, providing returns from 26-65 degree incidence angles. Analysis of the polarimetric parameters like entropy, anisotropy, and intensity from decompositions like Cloude-Pottier and Shannon show the oil slick can be differentiated from surrounding clean water and variations within the slick detected. The low-noise UAVSAR data allow even relatively subtle changes in oil properties to be measured.
DUAL BAND GNSS ANTENNA PHASE CENTER CHARACTERIZATION FOR AUTOMOTIVE APPLICATIONSjantjournal
High-accuracy Global Navigation Satellite System (GNSS) positioning is a prospective technology that will be used in future automotive navigation systems. This system will be a composite of the United States' Global Positioning System (GPS), the Russian Federation's Global Orbiting Navigation Satellite System (GLONASS), China Beidou Navigation Satellite System (BDS) and the European Union’s Galileo. The major improvement in accuracy and precision is based on (1) multiband signal transmitting, (2) carrier phase correction, (3) Real Time Kinematic (RTK). Due to the size and high-cost of today’s survey-grade antenna solutions, this kind of technology is difficult to use widely in the automotive sector. In this paper, a low-cost small size dual-band ceramic GNSS patch antenna is presented from design to real sample. A further study of this patch antenna illustrates the absolute phase center variation measured in an indoor range to achieve a received signal phase error correction. In addition, this low-cost antenna solution is investigated when integrated into a standard multi-band automotive antenna product. This product is evaluated both on its own in an indoor range and on a typical vehicle roof at an outdoor range. By using this evaluation file to estimate the receiver position could achieve phase motion error-free result.
DUAL BAND GNSS ANTENNA PHASE CENTER CHARACTERIZATION FOR AUTOMOTIVE APPLICATIONSjantjournal
High-accuracy Global Navigation Satellite System (GNSS) positioning is a prospective technology that will be used in future automotive navigation systems. This system will be a composite of the United States' Global Positioning System (GPS), the Russian Federation's Global Orbiting Navigation Satellite System (GLONASS), China Beidou Navigation Satellite System (BDS) and the European Union’s Galileo. The major improvement in accuracy and precision is based on (1) multiband signal transmitting, (2) carrier phase correction, (3) Real Time Kinematic (RTK). Due to the size and high-cost of today’s survey-grade antenna solutions, this kind of technology is difficult to use widely in the automotive sector. In this paper, a low-cost small size dual-band ceramic GNSS patch antenna is presented from design to real sample. A further study of this patch antenna illustrates the absolute phase center variation measured in an indoor range to achieve a received signal phase error correction. In addition, this low-cost antenna solution is investigated when integrated into a standard multi-band automotive antenna product. This product is evaluated both on its own in an indoor range and on a typical vehicle roof at an outdoor range. By using this evaluation file to estimate the receiver position could achieve phase motion error-free result.
DUAL BAND GNSS ANTENNA PHASE CENTER CHARACTERIZATION FOR AUTOMOTIVE APPLICATIONSjantjournal
High-accuracy Global Navigation Satellite System (GNSS) positioning is a prospective technology that will be used in future automotive navigation systems. This system will be a composite of the United States' Global Positioning System (GPS), the Russian Federation's Global Orbiting Navigation Satellite System (GLONASS), China Beidou Navigation Satellite System (BDS) and the European Union’s Galileo. The major improvement in accuracy and precision is based on (1) multiband signal transmitting, (2) carrier phase correction, (3) Real Time Kinematic (RTK). Due to the size and high-cost of today’s survey-grade antenna solutions, this kind of technology is difficult to use widely in the automotive sector. In this paper, a low-cost small size dual-band ceramic GNSS patch antenna is presented from design to real sample. A further study of this patch antenna illustrates the absolute phase center variation measured in an indoor range to achieve a received signal phase error correction. In addition, this low-cost antenna solution is investigated when integrated into a standard multi-band automotive antenna product. This product is evaluated both on its own in an indoor range and on a typical vehicle roof at an outdoor range. By using this evaluation file to estimate the receiver position could achieve phase motion error-free result.
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This document discusses the design and testing of a dual-band ceramic patch antenna for use in automotive GPS applications. It aims to characterize the antenna's phase center variation and offset to enable high-precision positioning. The antenna was tested on a 250mm ground plane, integrated into a shark-fin antenna product, and mounted on a vehicle roof to measure its performance in different scenarios. The results will help estimate receiver position accurately by accounting for phase errors introduced by the antenna geometry.
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Webinar Recording: https://www.panagenda.com/webinars/hcl-notes-and-domino-license-cost-reduction-in-the-world-of-dlau/
The introduction of DLAU and the CCB & CCX licensing model caused quite a stir in the HCL community. As a Notes and Domino customer, you may have faced challenges with unexpected user counts and license costs. You probably have questions on how this new licensing approach works and how to benefit from it. Most importantly, you likely have budget constraints and want to save money where possible. Don’t worry, we can help with all of this!
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2. Contents:
1. Background 8. Structure
2. Requirements 9. ADCS sub-system
3. Completed PDR Design 10. Propulsion sub-system
and Summary 11. Cost Estimation
4. System engineering 12. Risk Management
5. Orbits and Constellation 13. Reliability
6. Geo-location 14. Work Breakdown Structure (WBS)
7. Formation Flying 15. Summary and Acknowledgments
6 November 2011
3. Background
• The oceanic surrounding is hazardous and present risks of
drowning , hypothermia, shark attacks and more…
• Due to the nature and size of the oceanic surrounding, the
process of receiving distress signals and locating people in
distress accurately is somewhat problematic.
• Between hundreds to thousands of sea-related accidents
occur every year.
6 November 2011
4. Customer Requirements
1. The system shall locate and a person in distress in any watery surrounding around
the world (oceans, seas, rivers…)
2. The user shall wear an emergency beacon that will transmit a distress signal when
activated.
3. The time interval from distress signal transmission to notification in one of the
ground stations shall not exceed 15 minutes.
4. The computed location of the person in distress shall be no more than 1 km of his
true location.
5. As an option, the system shall allow enhanced capability for future applications
such as search and rescue services for “land incidents”, given the appropriate
modifications.
6. The system shall be based on space and satellites technology.
7. The space segment should be implemented using Nano-satellites ("Cube-Sat").
8. Each satellite's mission life-time shall be at least 2 years.
6 November 2011
5. Top Level Mission Requirements
• A user in distress shall be detected in less than 15 minutes, from signal transmission to ground
station notification.
• A user in distress shall be geo-located with an accuracy < 1 km
• The distress signal shall be relayed to a ground station
• The system's services shall be affordable to the common end user.
• The system shall be capable to identify its users in distress, as valid subscribers.
• Earth coverage range shall be at least between latitudes +60⁰) and (-60⁰)
• International space-related standards and regulations should be met, as much as possible
Top-level System Requirements:
• "Cubesat" satellite platforms shall be considered.
• Each satellite's mass shall be less than 10 kg
• Each satellite life time shall be at least 2 years.
• Satellite bus shall be designed using space-proven COTS sub-systems and components, as much as
possible.
• Satellite's sub-systems shall withstand launch load and space environment.
• Geo-location shall be performed using DTOA technique, using 2 or 3 reception satellites.
6 November 2011
6. Completed PDR Design Work:
Electric Power System:
EPS + Matching Solar Array to Battery to Consumers Max
Mass
Battery Battery Efficiency Efficiency DoD
ClydeSpace 3U EPS +
Battery Pack
90% 90% 20% 170 g
Efficiency @ 5E14 e-
Solar Panel Qty. Efficiency (BOL) Cell Area Cell Weight
/cm2
Azure TJ 3G30C 26 29.1% 30.18 cm2 2.6 g 26.5%
Thermal Control:
•Passive Control
•Steady State mean temperature: -6⁰
6 November 2011
7. GPS
Communication: Other CubeSat
User Geolocation Satellite Telemetry Main CubeSat
Segment to Satellite & Control
Satellite Ground Station Satellite Ground
Station
Antenna Monopole Patch Parabolic 2 Dipole 2 Dipole Yagi
Dipole
Transmitter 2.4Ghz 2.4Ghz -- 450Mhz 400Mhz 400Mhz
Receiver -- 2.4Ghz 2.4Ghz 450Mhz 400Mhz 400Mhz
6 November 2011 User's Beacon
Ground Station Ground Station
for GeoLocation for Control and
Broadcasting: data receiving
and handling
Telemetry
6 November 2011
8. Launch Segment:
1. Poly-PicoSatellite Orbital Deployer “P-Pod MKIII”:
• mass 1.5 kg
• can carry 3 (1U) cubesats or 1 (3U) cubesats
• number of deployers can be mounted together on a L.V
2. Launch Vehicle: SpaceX - Falcon 1e
Payload
Inclination Mass capability Est.
Altitude space Accuracy Reliability
[deg] [kg] Cost
[m]
Any above D1.55 x i = 0.1 [deg]
LEO 800 to 700[km] Med $10.9M
9⁰ H1.7 Apogee = 15[km]
6 November 2011
9. PDR Summary - Mission:
Constellation : 700 km , i 45, e 0 , 48 satellites, 6 planes
Geolocation: TDOA algorithm, 97% location within 15 min,
3% location within 30 min
Formation: 2 satellites, In-plane formation, relative control,
distance = 200 ± 50 km
6 November 2011
10. PDR Summary - System:
Mass: 3.11 kg Communication: 2 dipole, receiver, transmitter
Thermal Ctrl: Passive Payload: Patch antenna, transceiver
Attitude Ctrl: Active, 3-axis Propulsion: Warm gas, Isp=100
Available Average Power: 6.78 W
6 November 2011
12. Mission Profile
ˆ
x
ˆ
y
ˆ
z
Launch and Initial Dispersion Mission De-orbiting
Deployment stabilization Operation
10 Mins 24 hours 14 days 2 years 1-2 years
6 November 2011
13. Budgets
ΔV Budget Mass Budget Power Budget
PDR CDR Power consumption [mW]
PDR CDR Sub System Consumers
Usage System Total System Total
ΔV[m/s] ΔV[m/s] Cruise Detection Maneuver
Mass [Kg] Mass [Kg]
Positioning OBDH 0.08 0.08 OBDH 200 600 600
Keeping ADCS 0.209 0.09 ADCS 430 630 630
Formation Propulsion 1.209 0.458 Propulsion 0 0 2000
Deorbiting 0
Thermal 0 0
Spare Thermal Control 0 0 0
0.94 Control
(20%) Communication 200 450 450
Communication 0.23 0.23
Total 10.31 Payload 0.105 0.105 Payload 200 450 450
GPS 0.003 0.003 GPS 200 200 200
Power 0.297 0.237 EPS 200 200 200
Structure 0.958 1.02 Structure 0 0 0
De-Orbit - 0.08 De-Orbit 0 0 0
Total 3.111 2.3 Total 1430 1880 4530
6 November 2011
14. Design Iteration:
Subsystem’s Mass: Satellite’s Mass:
Allocation for Mass [Kg] Comments
Sub-System
Sub-System Sub-System Dry Mass 2.3
Total Mass [Kg]
[Kg] 10% Margin 2.53 X+10%
Power 0.31 0.2376 Includes 10% Margin for
Fuel 0.031
ADCS 0.22 0.09 Fuel
Thermal Control 0 0 Includes:
Communication 0.265 0.23 Total 2.56 10% margin for Fuel and
Payload 0.11 0.105 10% margin for Dry Mass
GPS 0.004 0.003
OBDH 0.1 0.08
Propulsion 0.7 0.458
Structure 1.05 1.02
De-Orbit 0.2 0.08
Total 2.955 2.3
6 November 2011
19. Orbits and Constellation
PDR Summary:
• A Walker constellation 45:24/6/1
• Constellation altitude - 700 km
• Constellation inclination of 45⁰
• Total of 48 satellites.
• Total of 24 formations
• 2 satellites per formation with nominal distance of 200 km between satellites.
• 6 orbital planes, each orbital plane consisting of 8 satellites (4 formations)
• Satellite de-orbitization at EOL using propulsion to lower the satellite from 700
km to 650 km, requiring v 13.35 sec m
6 November 2011
20. Design Updates Since PDR:
• Altitude had been changed from 700 km to 710 km.
• At 710 km ionization dose is about 6 krad for 0.6 mm shielding thickness.
still well within the 10 krad restriction of the sensitive EPS system.
• In 2 years (mission life time) satellites decline approximately 10 km.
at EOL, altitude is around 700 - higher than the minimum of 697 km
• No altitude correction maneuvers are required throughout the entire mission.
Constellation Revisit Time Vs Altitude
Revisit Time [min]
Altitude [km]
6 November 2011
21. De-Orbiting
m
• PDR calculation for de-orbiting from 700 km to 650 km: v 13.35
sec
• From 710 km to 650 km – even higher: v 16.01 m
sec
• 2 Alternatives for De-Orbiting were considered:
Alternative #1: “Jack in the Box”
• De-Orbit mechanism designed and manufactured by NASA for the O/OREOS
mission.
6 November 2011
22. • NASA’s de-orbit mechanism increases satellite’s surface area, and
thus drag force, by 60%
• Device’s dimensions:
Material:
Aluminum plates
Germanium Film
28 cm
9.9 cm
9.9 cm Weight:
~200 gr (est.)
Device can be placed only on top or bottom panel
6 November 2011
24. Method of Operation:
• The conductive tape produces
current up the tape upon
interaction with ionspheric plasma
• Charged tape interacts back
with earth’s magnetic field
to produce Lorentz Force
that opposes orbital motion
and produces electrodynamic drag.
L
F I B dl
0
6 November 2011
25. Device’s Performance:
• The extended tape’s surface area is about 152 cm², increases
spacecraft surface area by 50 %
• Deorbit Time Prediction with mechanism:
CAESAR
Satellites
6 November 2011
26. De-Orbit Mechanism Selection
Criterion
Criterion
Weight
Propulsion Based "Jack-in-the-Box" nanoTerminator
Value Score Total Value Score Total Value Score Total
Mass 0.5 400 gr 1 0.5 200 gr 3 1.5 80 gr 5 2.5
Deorbit
0.2 24.4 yr 2 0.4 22.2 yr 3 0.6 <1 yr 5 1
Time
Compat-
0.3 1 0.3 3 0.9 5 1.5
ibility
Total 1.2 3 5
The nanoTerminator gives us the best deorbit time, for the lowest
additional mass, and is the easiest to integrate with the satellite.
6 November 2011
27. Geolocation
The TDOA location method
t21
1
s2 u
1
s1 u • The hyperbolic equation
c c can be transformed to a
2 2 2
si u Xi X Yi Y Zi Z quadratic form
c is the speed of light
M m u
uT Mu 2mT u m0 0 uT 1 0
mT m0 1
T
M 4 s1 s2 s1 s2 4d 2 I m 2
2 s2 s12 s1 s2 2d 2 s1 s2
2 2
m0 s2 s12 d2 2 s12 2
s2 d2 d c t21
6 November 2011
28. • If no measurement errors exist the target must lie on the hyperboloid defined
by this quadratic form where the 2 satellites in the formation are the focal of
the hyperboloid. In this case 3 TDOA measurements can define the 3 unknown
target coordinates.
Satellite Formation and Target on TDOA Hyperboloid Satellite Formation and Target on TDOA Hyperboloid
SAT1 SAT1
SAT2 SAT2
Target Target
TDOA Hyperboloid TDOA Hyperboloid
4
3
2 -4
1 -3
0
-2
Z
-1
-1
-2
0
-3
5 1
-4
2
4 5
0 3
2
4 Y
0 0
-2 -5 5
-4 -5 X
Y X
6 November 2011
29. • If the targets location is known to be constrained on the surface of the Earth only 2
more TDOA’s are needed to find the location.
• Based on the analytical solution shown by Ho and Chan for a 3 satellite formation and a
single TDOA measurement, we have derived an iterative algebraic method for a 2
satellite formation using 2 TDOA measurements.
Target in the Intersection of a Sphere and 2 TDOA Hyperboloids
SAT11
SAT21
Target
SAT12
SAT22
8
6
4
Z
2
0
-2
5
5
0
0
-5
-5 -10
Y X
6 November 2011
30. • In the presence of measurement errors the initial location can be far from the true
location of the target. In order to improve the initial location error an Extended Kalman
Filter starting with the initial location is used with all of the TDOA data. The estimated
target location then drifts from the initial location closer to the true location.
6 November 2011
31. Experiment Scale Down
• In order to improve the reliability of the geolocation algorithms and
examine them in a more realistic environment we have conducted an
experiment at the Distributed Space Systems Laboratory (DSSL) in the
Asher Space Research Institute.
Parameter Space Scale EchoLab Scale
Formation ~200 km ~500 mm
Target Range 700-3000 km 3000-4000 mm
V phase EM 300e3 km/sec Acoustic 340e3mm/sec
TDOA 0-300 microsecond 0-300 microsecond
SD time ~50 nanosecond ~50 microsecond
SD length ~0.015 km ~17 mm
6 November 2011
32. Acoustic TDOA Experiment:
The satellite formation is hovering on a 4 on 4 meters air table. The target is
mounted 3 meters above the table and transmits 40 KHz acoustic pulses.
Satellites
Ultrasound
Transmitter
6 November 2011
33. Satellite Formation Flight and Target Location on Table Plane
Nominal target range is 3.089[m]
Nominal distance in formation is 0.617[m]
0.2
0.1
0
-0.1
-0.2 Target
Y [m]
SAT1
-0.3 SAT2
-0.4
-0.5
-0.6
-0.7
-0.6 -0.4 -0.2 0 0.2 0.4 0.6
X [m]
6 November 2011
34. Evolution of Location Error
Time SD is 15[ sec] Sv's Location SD is 1[cm]
0.2
3 0
Z [m]
2.95
-0.2
0.1
0 -0.4
-0.1 -0.6
-0.2 X [m] Estimation
-0.8
Y [m] Initial
Target
6 November 2011
35. Final Estimation Error is 8.32[cm] with 3 = 20.9 [cm]
120
100
80
[cm]
60
40
20
0
2 4 6 8 10 12 14 16 18 20 22
[Estimation Steps]
6 November 2011
36. Formation Flying
PDR summary
• In the first semester we selected the following principals:
– 2 satellites per formation
– In-plane formation
– Relative control method
Nominal State Distance
Reaches Elliptic Verify Successful
+ Altitude
Boundary Maneuver Maneuver
Maintenance
• This means that in worst-case scenario, ∆V required to
m
maintain formation and altitude is V = 7.774
sec
6 November 2011
37. CDR Revisions:
1. No Altitude Maintenance
– Satellites are allowed to lose altitude
– Once correction is needed, the maneuvering satellite also changes its
altitude to that of its partner’s (Hohmann Transfer)
– In worst-case scenario, ∆V required to maintain formation is:
m
V = 2.52
sec
Elliptic
Distance Maneuver
Reaches Verify Successful
Nominal State
Boundary Hohmann Maneuver
Transfer
6 November 2011
38. 2. Statistical Analysis
• In an attempt to reduce ∆V required, we performed a statistical
analysis of the actual scenarios that may occur.
– Euler angles of satellite are normally distributed ( 0, 2.5 )
– 800 simulations performed
• Results:
– 69% of cases –
No correction required
– 31% of cases – 1 correction required
– In none of the cases
were 2 corrections needed
m
• Conclusion: ∆V needed to cover 99.99% of all cases is V = 0.18
sec
6 November 2011
40. Structure
PDR Summary
A 3U cubesat has been chosen for the satellite`s structure.
Inner Components Placement
Guidelines:
• Maximum distance between magnetometer and magnetic field generators
(magnetorquers, electrical components)
• Center of mass should be as close to geometric center as possible
• Thrust vectors should pass as close to the center of mass as possible
• Patch antenna facing Nadir direction
• GPS Antenna facing Zenith direction
6 November 2011
46. Analysis
A Finite Elements method is required – In order to reduce the
complexity of the geometric model a simplified model was suggested:
• The inner components are referred as “Point Mass”
• Three mass points simulate the three major parts
• Each point mass is connected through 8 points to the satellite’s
skeleton in order to simulate the real assembly
6 November 2011
47. Modal Analysis
The boundary conditions are fixed support on all eight legs of the
skeleton in order to simulate the satellite in the launch POD.
6 November 2011
48. Modal Analysis
1st mode 2nd mode 3rd mode
The first 6 modes are:
Mode Frequency [Hz]
1 695
2 708.11
3 755.18 4th mode 5th mode 6th mode
4 756.94
5 769.25
6 769.6
6 November 2011
49. Static Analysis
A 16g load was set in the longitudinal direction and a 2.75g was set in
the lateral direction.
The results show the satellite will endure the launch loads even with a
10 degree misalignment with its long axis.
16g
10 deg
6 November 2011
50. Static Analysis Results
Middle – Deformations Top – Stress Entire Satellite Deformation
6 November 2011
51. Attitude Determination & Control Subsystem
Requirements
1. Spacecraft shall be 3 axis stabilized
2. Spacecraft's long axis shall be Nadir Oriented
3. Maximum pointing error (per axis):
1. Cruise Mode: less than 5⁰
2. Engine Ignition: less than 10⁰
4. ADCS sub-system's mass shall be less than 190 grams
5. Maximum power consumption shall be less than 630 mWatt
6. Maximum time from deployment from launch pad until initial
stabilization shall be less than 24 hours
6 November 2011
52. PDR Review:
• Attitude control actuators: 3 magneto-torquers.
• Attitude Determination: Magnetometer + Analog Sun Sensors
(preliminary design)
• Hardware Selection
• Disturbance torque estimation
6 November 2011
53. Hardware Updates:
PDR CDR
Honeywell
Billingsley
HMC 5843
Magneto TFM65-VQS
(Integrated to OBC)
-meter
117 gr 50 milligram
3.51x3.23x8.26 [cm³] 4x4x1.3 mm
Satellite Services LTD Visio Torquer
Torquer rod (x3) PCB
Magneto
-Torquer m = 30 gr
m = 100 gr
L=7 cm
Size: 10 x 9 cm
D=0.9 cm
Dipole = 0.5 Am²
Dipole=0.2 Am²
6 November 2011
54. Analog Sun Sensor Design
Current to sun angle of attack relation: ˆ
y
I I max sin
ˆ
x
ˆ
z
I max is the current measured
when the sun shines directly
in the normal direction:
90
6 November 2011
55. The Current-Sun AOA Relations:
I1 I max cos 3 sin 1
I2 I max cos 3 sin 2 CAESAR
CAESAR
I3 I max sin 3
Top View Side View
I1
Finding the AOA angles using: sin 2 cos 1
and 1 arctan
I2
Sun’s vector in Body Frame is written as:
sin 1 cos 3 X 0 0 I1
VsB sin 2 cos 3 VsB 0 Y 0 I2 where, X , Y , Z 1
sin 3 0 0 Z I3
6 November 2011
56. Attitude Determination Algorithm
I
• Computing Sun Vector and Magnetic Vector in ECI - Vsun , Vmag
I
• Using sensor’s data to derive Sun Vector and Magnetic Vector in
B B
body frame: Vsun ,Vmag
• Finding a rotation matrix from Body Frame to ECI:
I I I I I B B B B 1
C B V sun V mag V sun V
mag V sun V mag V
sun V mag
• Finally, Finding rotation matrix from Body Frame to VVLH:
VVLH
CB CIVVLH CB
I
• From the rotation matrix it’s easy to derive Euler angles by:
C2,3 C1,3 C1,2
arctan , arctan , arctan
C3,3 2
C
2,3 C 2
3,3
C1,1
6 November 2011
57. Problem: During Eclipse sun’s Vector in body frame is unattainable.
Consequence: Attitude determination of the satellite during eclipse
is unattainable.
Solution: Rotational rate estimation from 3 attitude measurements,
using Lagrange interpolation formula:
t t3 t 2 t3 t1 2t3 t1 t2
3 t1 t2 t3
t1 t2 t1 t3 t2 t1 t2 t3 t3 t1 t3 t2
t t3 t 2 t3 t1 2t3 t1 t2
3 t1 t2 t3
t1 t2 t1 t3 t2 t1 t2 t3 t3 t1 t3 t2
t3 t 2 t3 t1 2t3 t1 t2
t3 t1 t2 t3
t1 t2 t1 t3 t2 t1 t2 t3 t3 t1 t3 t2
6 November 2011
59. Control Design
the control algorithm needs to deal with the following disturbances:
• Gravity • Engine Torque
• Solar Pressure
• Atmospheric Drag yˆ
• Magnetic Field ˆ
x
ˆ
z
g
6 November 2011
60. Control Design – State-Space
Our state-space equations will be:
P Q R
T
A
I m b ng nd 0
0 0 0 0 0 0
0 0 0 1 0 0 0 0 0 0 0 0
0 0 0 0 1 0 0 0 0 0 0 0
m1
0 0 0 0 0 1 b3 b2 1
0 m2 0 0 nd
2 Ix Ix Ix
P 4 0 1 0 0 0 0 0 1 1 P
b3 b1 m3 1
Q 0 3 2
0 0 0 0 Q 0 0 0
0 2 Iy Iy Iy
R 0 0 2 2
1 0 0 R
0 3 0 3
b2 b1 0 0 1
0 Iz
Iz Iz
While: ng – The gravity gradient disturbance moment
nd – The remain disturbances moments
m – The control dipole moment
b – The magnetic field
6 November 2011
61. Control Design – Control System Topography
euler
distubances
moments
nd
nd+ngg -K-
Y=eye*X
T
u=-Kx r2d angles
T_ctrl X
1 err u u
In1 -K-
StateSpace
PID+
Anti WindUp r2d1 rates
-K-
Anti WindUp
Ks
PID -K-
1
s
Integrator1
b
in x Ki Saturation1
P 1
err
-K- 1
u
Kp
Q 2 -K-
rates
R Kd
Scope3
6 November 2011
64. ΔV Budget
ΔV[m/s]
Usage
PDR CDR
Positioning
Keeping
Formation
Deorbiting
Spare (10%)
Total
6 November 2011
65. PDR Summary
• There were 3 missions for the Propulsion System:
1. Positioning.
2. Keeping formation.
3. Deorbiting.
• We selected a warm gas propulsion system of
“MicroSpace”.
• We designed an external high pressure gas tank for
the propulsion system.
• The total propulsion system mass was 436 g.
• The cost of the propulsion system without the
external gas tank was € 81,000.
6 November 2011
66. Design updates since PDR
• There are 2 missions for the Propulsion System:
1. Positioning.
2. Keeping formation.
• We noticed that “MicroSpace’s” propulsion
system is too heavy, complicated and expensive.
So, we designed a new Cold Gas Propulsion
System that meet our specific requirements.
• The total propulsion system mass is 429 g.
• The cost of the propulsion system is $ 7,321.
6 November 2011
67. The Propulsion System's Block Diagram
Block Diagram And Detailed Components
Pressure Straight
Main
Transducer Pipe
Connector
Pressure
Connector
Pressure Solenoid
Regulator Valve
Fill Pressure
Valve Regulator
Straight High
Latch
Connector Pressure
Valve
Tank
Curve Solenoid
Pipe Pressure Valve
Transducer Pressure
Solenoid Regulator
Valve
Thruster
Fill Valve House
Latch Thruster
Valve
Gas Tank
6 November 2011
68. Strength Analysis And Optimization
• In order to design the most optimal
components, we made analysis with
“SimulationXpress”.
• At iterative work, we fit the wall thickness to
the applied pressure(at extreme conditions of
50°C), so we get the optimal weight.
6 November 2011
69. Design Parameters Optimization
In order to choose the most suitable design parameters we made graphs and at
iterative way we gathered to the best solution.
Thrust Vs. Pc Thrust Vs. Area Ratio
0.8 tpulse Vs. Area Ratio 0.08
Thrust Vs. Pc
Thrust Vs. Throat Diameter
Thrust Vs. Area Ratio
0.6
0.08 0.8
420
1.5 0.075
F [N]
F [N]
0.4
0.6 0.07
0.2
t pulse [sec]
0.0751
400
F [N]
F [N]
0 0.065
0.4
F [N]
5 10 15 20 25 30 35 40 45 50 55 60 20 40 60 80 100 120 140 160 180
Pc [atm] Area Ratio Ae/At
t pulse Vs. Pc Isp Vs. Area Ratio
0.5
380 80
3000
0.07 0.2
75
tpulse [sec]
Isp [sec]
2000
0 0
1000 360
0.065 0.1 5
0.2 10
0.3 15
0.4 20
0.5 25 0.6 30 0.7 350.8 40 0.9 45 1 50 1.155
70 60
20
20 40
40 60
60 80 100[atm]120 140 160 180
Throat Diameter [mm] 140
80 100
Pc 120 160 180
Area Ratio Ae/At
tpulse Vs.tRatio Vs. Pc
60 Area Ae/At
0 65
pulse Diameter
Throat
5 10 15 20 25 30 35 40 45 50 55 20 40 60 80 100 120 140 160 180
Pc [atm]
mIsp Vs. Area Ratio
prop
Vs. Area Ratio Area Ratio Ae/At
3000
3000
80
44 Thrust Vs. Throat Diameter tpulse Vs. Area Ratio
1.5
420
tpulse [sec]
2000
[sec]
1 2000
42
75 t pulse [sec]
mtprop [gr]
400
Isp [sec]
F [N]
pulse
0.5
40 1000
380
1000
0
0.1 0.2 0.3
70 0.5
0.4 0.6 0.7 0.8 0.9 1 1.1 360
20 40 60 80 100 120 140 160 180
38 Throat Diameter [mm]
Area Ratio Ae/At
0 0
tpulse Vs. Throat Diameter
mprop Vs. Area Ratio
3000 5 10 15 20
0.5 25 0.6 30 0.7 350.8 40 0.9 45 1 50 1.155 60
36 65 0.1 0.2 0.3 0.4 44
20
20 40
40 60
60 80 100[atm]
Pc 120
Throat Diameter 120
80 100 140
[mm] 140 160
160 180
180
tpulse [sec]
2000 42
mprop [gr]
Area Ratio Ae/At
Area Ratio Ae/At
40
1000
38
6 November 2011 0
0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8
Throat Diameter [mm]
0.9 1 1.1 36
20 40 60 80 100 120 140 160 180
Area Ratio Ae/At
70. Cold Gas Thruster - Specifications
The Cold Gas Thruster parameters:
(T≅278K)
Parameter Value Parameter Value
Throat diameter 0.3 mm 10.31 m/sec
Exit diameter 3 mm Pulse time 370 sec
Propellant mass (N2) 37.6 gr
Thrust 75.2 mN
Tank Pressure 137 atm = 2,015 Psi
Isp 74.8 sec
Pc 6 atm
Pe ~0 atm
6 November 2011
72. Cost Estimation - Propulsion example
Gas Pressure Pressure Latch Solenoid Control
Part Name Thruster Fill Valve Fasteners
Tank Regulator Transducer Valve valve Board
Part's Cost $ 303 66 1,500 885 900 233 400 452 724
Time [Min] 120
Rate
Assembly Work 100
[$/Hour]
Cost $ 200
Opacity Test - Time [Min] 120
Helium mass Rate
Work 70
spectrometer [$/Hour]
with bell jar Cost $ 140
Quantity for the constellation 48 96 96 48 48 96 48 48 48
Total Cost per Satellite $ 7,268
Total Cost for Entire Constellation $ 348,856
Pressing pattern $ 10,000
Non- Opacity Equipment $ 5,000
recurrent Environmental
30,000
testing $
Total non-recurrent cost $ 45,000
6 November 2011
73. Cost Estimation
Recurrent cost $
Components Cost for Entire Non-recurrent cost $
Cost per satellite
Constellation
Satellite Structure 6,585 316,113
Propulsion System 7,268 348,856 45,000
ADCS 19,604 941,008
Payload System 13,355 641,072
Communication System 23,149 1,111,192
Power 17,368 833,664
Formation 16,600
Geolocation 35,480
Total cost 87,329 4,191,905 97,080
Constellation cost 4,288,985
Launching the entire constellation (6 launches) costs : ~ $ 67,714,464
6 November 2011
74. Risk management
Every project has risks –uncertainties that weren't anticipated earlier
Risk Management -identifying, analyzing and responding to project risk.
project risks are uncertainties that may result in schedule delays, cost overruns,
performance problems, adverse environmental impacts or other undesired
impacts.
Pf - The likelihood of the event
C - The potential consequence to the project
R - Risk factor R P C
f
6 November 2011
75. Risks analysis-The 7 major risks in the project
Likelihood
Not Likely Likely Very Likely Consequence
Low Risk Low Risk Low Risk Benign
Low Risk Medium Risk Medium Risk Medium
Low Risk Medium Risk High Risk Harsh
Risk Pf C R-Risk Factor
Propulsion system: Safety risk-the system contains 0.8 0.8
high pressure, chance of explosion.
0.64
Risk Mitigation: Performing experiments and tests on the system and particularly on the tank.
Propulsion system: Schedule risk- the launch 0.8 0.7 0.56
company will not agree to launch the satellite.
Risk mitigation: Experiments and higher safety factors.
Propulsion system: Technical risk- The amount of 0.7 0.7 0.49
gas might not be enough -sun storms increase drag.
Risk mitigation: Increasing the percentage of spare gas in the tank. This spare gas will be used in
unexpected weather in space.
6 November 2011
76. Risk Pf C R-Risk Factor
Propulsion system: Technical risk-Center of mass 0.7 0.6
wouldn't coincide with the engine`s nozzle. 0.42
Risk mitigation: Designing a new propulsion system or moving components in the satellite.
Launch: Finding time windows suitable for the launch of 0.5 0.9
24 pairs of satellites in 6 different launch dates.
0.45
Risk mitigation: Communicating with launch provider in advance as possible in order to decrease the
probability of such failure.
Orbits and constellation: Technical risk-Satellite collision 0.4 0.9
with space debris
0.36
Risk mitigation: Running Debris assessment simulations using NASA's Debris Assessment Tool, and
STK. Launching redundant (extra) satellites to account for damaged satellites.
Attitude control: Stabilization of the satellite by the 0.5 0.8
attitude control system.
0.4
Risk Mitigation: Testing the satellite in a laboratory and performing simulations.
6 November 2011
77. Summary: Number of risks
system
The propulsion system has the found
most risks in the project and Propulsion 8
the consequence of its risks is Attitude control 2
the most severe. This is Geolocation 1
understandable since the Structure 2
launch 2
propulsion system is new and
Formation 2
we don’t have previous Keeping
experience with such systems. Orbits and 2
This means we would have to constellation
Electrical Power 4
perform more experiments
Communication 2
and tests on the system and of /payload
the system with the satellite Thermo control 1
in order to mitigate the risks. Total
6 November 2011
78. System Reliability
Reliability: “the probability that a device will work without failure
over a specific time periods or amount of usage” *IEEE, 1984].
R e t
R - Success Probability, - Failure Rate , t -Time Period
Series Reliability:
A B C RS RA RB RC
Parallel/Redundant Reliability :
A
B RP 1 1 R A
1 RB 1 RC
C
6 November 2011
79. For our mission t=2 years and is taken as constant,
2 years
so reliability is computed as: R e dt t
0
Example: Propulsion Subsystem Reliability
R5=0.9801 R6=0.992 R7=0.99
Pressure Solenoid
Thruster
Regulator Valve
Pressure Propellant Latch
Fill Valve
Transducer Tank Valve
R1=0.988 R2=0.9999 R3=0.996 R4=0.9994
Pressure Solenoid
Thruster
Regulator Valve
R5=0.9801 R6=0.992 R7=0.99
2
Rpropulsion R1 R2 R3 R4 1 1 R5 R6 R7 0.9638546
6 November 2011
80. Mission Reliability
In order to calculate the mission reliability we calculated the
reliability of each phase of the mission:
Satellite
Initial Mission
RLLaunch
0.97 R 0.94
Stabilization
S
RPositioning
Ph 0.91 ROperation
M
0.902 RDDeorbit
0.97
(Phasing)
De-Orbit
ADCS ADCS Computer
ADCS Computer Computer EPS
EPS EPS Communication Communication
Communication Communication EPS
Propulsion Propulsion
Payload
Mechanism
RMission RL RS RPh RM RD 0.735
6 November 2011
81. Summary - Compliance to Requirements:
Mission
Requirement Result Compliance
Constellation Revisit Time < 15 min 14.74 min
Geo-Location Location Radius < 1 km 97% < 1 km
De-Orbiting within 25 years 1-2 years
Orbits Global Coverage between Available Coverage:
latitudes +60⁰) and (-60⁰) +60⁰) and (-60⁰)
~$87K per Sat
Cost Cost-efficient
~$4.2M Total
System
Satellite’s Mass Each Satellite’s mass < 10 kg 2.56 kg
6 November 2011
82. Acknowledgments
We’d like to express our appreciation and gratitude to all
Those who have helped us:
Prof. Pini Gurfil, Dr. David Mishne, Dr. Zvi Hominer,
Dr. Avi Vershavski, Ofer Slama.
And special thanks to our supervisor
Jacob Herscovitz
6 November 2011
על מנת לפשט את המשדר ככל האפשר, אין תקשורת בין המטרה והלוויינים למעט פולסים בודדים שמשדרת המטרה החל מרגע מסוים. ללא מידע על זמן או מיקום השידור שיטת האיתור שנבחרה היא מדידת הפרש הזמנים בקליטת הפולסים. כל הפרש זמנים מתאר היפרבולואיד במרחב שזוג הלוויינים הם המוקדים שלו.
בהיעדר שגיאות מדידת זמנים ומיקומי לוויינים מיקום המטרה מאולץ על כל אחד מההיפרבולואידים המתאימים למדידות הפרשי הזמנים.באמצעות 3 מדידות ניתן להגדיר 3 היפרבולואידים, אשר בחיתוך שלהם ממוקמת המטרה.
אם ידוע כי המטרה ממוקמת על פני כדוה"א, ניתן להשתמש באילוץ נוסף זה יחד עם שתי מדידות הפרשי זמנים על מנת לאתר את המטרה.כאשר מבנה של 3 לוויינים קולט שידור אחד ומפיק ממנו שתי מדידות הפרש זמנים, קיים פתרון אנליטי למציאת מיקום המטרה המשדרת על פני כדוה"א.על מנת לצמצם את מספר הלוויינים במבנה למינימום הכרחי של 2 לוויינים, הרחבנו את השיטה האנליטית לשיטה איטרטיבית המאפשרת באמצעות קליטה של 2 פולסים לאתר את המטרה המשדרת על פני כדוה"א.
בנוכחות רעש מדידה, המיקום הראשוני המתקבל משתי מדידות באמצעות השיטה האיטרטיבית עשוי להיות לא מדויק מספיק על מנת לעמוד בדרישות המשימה.באמצעות מספר מדידות גדול ניתן לשפר את המיקום הראשוני באמצעות משערך קלמן מורחב. בדוגמא שלפנינו ניתן לראות תוצאות של סימולציה בה המיקום הראשוני התקבל מחוץ למעגל 1 ק"מ, ולכן חורג מדרישות המשימה.באמצעות מספר עשרות מדידות נוספות שגיאת המיקום מצטמצמת למספר עשרות מטרים בלבד.
על מנת לבחון את אלגוריתם האיתור בתנאים אמיתיים יותר, ולא רק בסימולציה ממוחשבת, ערכנו ניסוי במעבדה למערכות חלל מבוזרות, שבמכון אשר לחקר החלל.האתגר בניסוי זה הוא לבחון מערכת המתוכננת לטווחים של מאות ואלפי ק"מ במעבדה על פני כדוה"א. מכיוון שבמערכת המתוכננת לחלל הפולסים נעים במהירות האור,הפרש הזמנים שהיה נמדד במעבדה על פני מטרים בודדים בין שני מקלטים היה מספר ננו שניות. בהיעדר יכולת למדוד הפרשי זמנים כאלו, פולס השידור האלקטרומגנטי הוחלף בשידור אקוסטי כך שהפרשי הזמנים הנמדדים הם באותו סדר גודל של הזמנים המתוכננים למערכת בחלל.
לתאר את המעבדה ולהראות עוד תמונות.
ניתן לראות את מסלול שני הלוויינים על שולחן האוויר, ואת היטל המטרה על השולחן (המטרה ממוקמת כ- 3 מטרים מעל השולחן).
תוצאות שערוך מיקום המטרה באמצעות אותו האלגוריתם בו בוצעו הסימולציות למעט שינוי פרמטר מהירות הפולס ממהירות האור למהירות הקול.במקום אתחול לפי אילוץ מטרה לפני כדוה"א, אתחלנו את המיקום הראשוני לפי גובה המטרה מעל השולחן. סטיות התקן של מדידות הזמנים התקבלו כתוצר נוסף של תוצאות הניסוי.
כאן מוצג תקציב הדלתא וי למשימה. ניתן לראות, שעיקר תפקידה של מערכת ההנעה הוא למקם את הלוויין במסלולו. בנוסף לכך, משתמשים בהנעה גם לשמירת מבנה.