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Past Conference Papers:
Responsive Spacecraft
| Paper Number RS1-2003-4002: The Trailblazer Class of Low Cost Space Vehicle | |
| W. Paul Blasé (TransOrbit), Charles F. Radley (TransOrbital) | |
| View/Download:Presentation | Paper | |
| Abstract: The concept of responsive space has many overlapping aspects. Military requirements to obtain tactical data rapidly or to reconstitute a decimated on-orbit constellation represent one class of drivers for responsiveness. Another class is represented by commercial news broadcast requirements to obtain satellite imagery of areas struck by a major natural disaster or in a state of war. Finally, there are many scientific phenomena that are either short term or that occur with little advance warning which could benefit from a rapid response capability. Two components at the core of providing the rapid response capability are launch vehicles that are launch-on-demand and spacecraft that are rapidly reconfigurable to match mission requirements. This paper addresses the latter: a spacecraft bus originally designed for the first commercial mission to the Moon that is reconfigurable and able to host various payloads that can support a wide range of missions. The TrailBlazerTM spacecraft is a 90-cm diameter, 85-cm tall three-axis stabilized 8- sided prism with photovoltaic cells on each face to provide power independent of orientation. The symmetric design was chosen to reduce the risk of mechanical failure associated with gimbaled sensors, antennas or solar panels, and to reduce spacecraft mass. The tradeoff is the necessity to reorient the spacecraft for sensor pointing and to orient the antenna for data transmission. A number of factors can be adjusted to handle a wide range of missions. Although the surface area available for the solar cells is fixed, the type of solar cells can be varied from standard space-qualified silicon cells to high performance gallium arsenide cells, providing additional power depending upon payload power requirements. Similarly, as a function of mission duration and the associated radiation environment, a range of available off-the-shelf spacecraft control units, communications systems, and sensors have been identified that can be installed rapidly as required. Thus, it will not be necessary to have completely built up spacecraft in inventory. Rather, basic buses can be kept in inventory, along with various “plug-and-play” modules that can be installed rapidly, keeping costs down. In addition, the bus can be mated with a range of kick stages to give it a broad range of orbit capabilities: LEO, MEO, GEO, or lunar. Various combinations of optical sensors are possible, including dual high or low-resolution visible sensors, one high and one low-resolution visible sensor, or combinations of visible, infrared and ultraviolet sensors. Because of its small size, the TrailBlazerTM spacecraft is also easily adaptable to a wide range of launch vehicles, so that if one is not available to accommodate a rapid response, the spacecraft could be mated rapidly to another launch vehicle. Trailblazer’s small size also allows it to be easily transportable from one candidate launch site to another. The TrailBlazerTM thus represents a leap forward in defining the equation for combining rapid response, versatility, and low cost into a single entity to fulfill the needs of both commercial and military customers | |
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| Paper Number RS1-2003-4003: Implementing Standard Mircosatellites for Responsive Space | |
| Jeffrey L. Janicik (SpaceDev) | |
| View/Download:Presentation | Paper | |
| Abstract: Today satellites tend to be large, expensive, power-hungry, slow to assemble test and integrate, and are generally unique to each payload. In addition they often take as much as 90 days to “commission” on-orbit. This traditional way of doing things can slow the testing, demonstration and fielding of newer, smaller, higher performance payloads. The approach, articulated here, is to use the “microcomputer” way of thinking and apply it to the space industry, which has tended to be bogged down for decades in the bigger is better “mainframe” way of thinking. Because of these problems and variables, the goal is to develop high performance modular microsatellites and a corresponding microsat operating system quickly and efficiently, not custom for each mission or payload. This can be simplified if functionality becomes a goal, where capabilities like radiation hardness, shock, vibrations, etc., are simply built into the microsatellite, with the system able to perform to a certain acceptable level, relieving the payload developer from such considerations. An appropriately modular system can utilize current and future subsystem and software technology, while naturally maintaining stateof- the-art performance. In turn this can dramatically reduce the size, mass and power requirements of subsystems and will result in on-demand launch and responsive, quick onorbit commissioning. Consequently, the most performance can be put into the smallest, lightest package for payloads, and ensure the shortest time to productivity on orbit. | |
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| Paper Number RS1-2003-5002: Lessons Learned From Past Reusable Launch System Designs | |
| Gregory Peralta (Lockheed Martin) | |
| View/Download:Presentation | Paper | |
| Abstract: The X-33 single-stage-to-orbit (SSTO) technology demonstrator funded by the United States Government and industry provided significant data to support SSTO design. The system focused on developing a highly responsive, operational efficient SSTO that could validate the basic technology and proof-ofconcept. Recent endeavors, such as the NASA Space Launch Initiative Two-Stage-to-Orbit and SOV programs have utilized the methodology taken on the X-33. The X-33 enjoyed many design successes. Although, several set backs resulted in the premature termination of the X-33 program. In hopes of benefiting future designers of responsive, reusable launch systems, this paper will discuss the design philosophy and testing approach of the X-33 by emphasizing the lessons learned This paper will also review past programs and illustrate why future operational programs need distinctly separate developmental and operational programs. This separation must be accomplished by completely qualifying components and procedures during the development phase of a program. Flight and mission performance as well as ground system operations efficiencies can be achieved through a “design to operations” systems approach with correct implementation of updated technologies where appropriate. Government investment must be directed toward operations certification of new system technologies as much as it is on traditional flight performance qualifications. Tremendous improvements can be made in system operations, which will translate into program efficiencies and lower overall system cost from the first launch and continuing through the life of the program. | |
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| Paper Number RS1-2003-6002: Fast Responsiveness Experiment Flight Opportunities Using SSPC | |
| Gerald Murphy (Design_Net Engineering), Kirk Stewart (Design_Net Engineering) | |
| View/Download:Presentation | Paper | |
| Abstract: The need to rapidly and affordably fly small space missions for science investigations, technology demonstrations, and risk reduction efforts is a frequently cited critical deficiency in our national space efforts. For small payloads whose missions can be accomplished from a low altitude, mid-inclination orbit, this problem can now be solved. Using a NASA mission flown in CY2000 as a proof-of-concept demonstration, Design_Net Engineering, LLC has performed IR&D and developed concepts and designs for a flexible system of Small Satellite Payload Carriers (SSPC™). These systems are small enough to be accommodated late in the manifesting cycles for launch on Space Transportation System (STS) missions to the International Space Station (ISS). Launch in Orbiter locker or soft stowage areas enables payload missions with less than 6 month lead time and very low launch costs. An SSPC™ is deployed by shuttle astronauts during a short EVA, using standard EVA proven hardware. SSPC™s are externally attached to the space station at one of many attachment locations used by the astronauts during station assembly. An SSPC™ does not require any station resources, which is a key in obtaining accommodation. They operate independently of the space station by providing their own electrical power, telemetry, command, thermal control, avionics, and experiment interfaces. They communicate and operate directly to Earth through the payload user’s selected ground support system. This may be the DoD’s AFSCN, a NASA or commercial network, or a university ground station, and an SSPC™ operates as would an independent small satellite, although it is attached to ISS. These SSPC™ systems require no costly attitude control or propulsion systems, yet numerous attitudes, orientations, and mission durations are possible because of the large number of attach locations. Built affordably, as two separable modules, the SSPC™s use a main base module to house the typical satellite payload support or satellite bus systems and a second module to accommodate added payloads. The main base module stays on orbit after an initial deployment mission to provide a resident support capability for subsequent missions which use interchangeable payload modules. Payloads and missions may be of varied durations and benefit from the cost effectiveness of the reuse of the base module. Ultimate responsiveness is achieved by this system and space flight can be available as quickly as a payload can be built, integrated into a payload support module, and delivered for the next STS mission to the ISS. No other current concept offers this responsiveness or potential for the lowest cost flights on a regular basis. Typical payload support capabilities are in the neighborhood of 30 lbs, 30 to 50 watts, and 10 to 20kbps average data rate. Final capabilities will be determined by the sponsors’ requirement trades and applications. This concept is currently being evaluated for development by the government. Commercial opportunities are also under consideration. This paper will describe the concept, system, alternatives, and payload support capabilities. | |
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| Paper Number RS1-2003-6003: On-Demand Wavelength Tuning of Detector Responsivity for Multi-Mission Scenarios | |
| D.A. Cardimona (Kirtland Air Force Base), D. Huang (Kirtland Air Force Base), C. Morath (Kirtland Air Force Base), D. Le (Kirtland Air Force Base), B. Klemme (Kirtland Air Force Base) | |
| View/Download:Presentation | Paper | |
| Abstract: The ability to reconfigure a sensor in order to enhance performance and/or perform multiple missions is a very desirable attribute for future sensor systems. If a sensor system could reconfigure itself to exploit signals in the wavelength ranges from the UV through the IR and into the millimeter-wave regimes, that system could support missions such as cold-body detection, target discrimination and identification, plume-to-hard-body handover, surveillance through clouds, chemical/biological weapons detection, etc. and would be assured of operation 24 hours/day, 7 days/week, in all weather conditions and at very long distances. If this reconfiguration could be done with a single detector, the savings in cost, weight, and power consumption would be substantial. In this paper we present some of the work we have been doing in the area of wavelength tunability of single detector structures. An applied bias field can tune the response of a semiconductor quantum well infrared photodetector via the quantum Stark effect, and an applied magnetic field can tune the response via the Landau levels that appear in such a structure. The magnetic field can also enhance the quantum efficiency of the quantum well detector. In addition, if the quantum well structure is composed of two or more coupled wells, the tuning effect is enhanced. Alternatively, a lateral transport scheme, in which photocurrent travels across a multiquantum-well pixel rather than top-to-bottom within the pixel, has been developed that shows great promise for it’s tuning characteristics. With the addition of quantum dots within each quantum well layer, the quantum efficiency of the device can be improved. . | |
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| Paper Number RS1-2003-6005: Near Space Maneuvering Vehicle | |
| Robert E. Blackington (Schriever Air Force Base) | |
| View/Download:Presentation | Paper | |
| Abstract: US military and government agencies rely on space assets as intelligence, surveillance, reconnaissance (ISR), and communications platforms. Due to orbital constraints, these assets are not always in the optimal location in times of need and currently there is no "launch on demand" capability. Although aircraft have been proven as effective ISR and communications platforms, they are expensive to deploy to forward locations, limited in altitude and range, and are exposed to enemy threats. The Air Force Space Battlelab (SB) is exploring the concept of using a lighter-than-air craft, known as the Near Space Maneuvering Vehicle (NSMV), as an inexpensive alternative high altitude platform for a variety of critical payloads | |
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| Paper Number RS1-2003-6005: A Low Cost Flight Computer Using GPS | |
| Michael Castle (SiRF Technology), Luke Robinson (Cambridge University) | |
| View/Download:Presentation | Paper | |
| Abstract: A novel and recently proven implementation of a flight computer is presented. Traditionally a large proportion of the total launch cost is the avionics, in particular low drift gyros and fast powerful servos. This paper describes the implementation of a flight computer using a commercial GPS receiver and lightweight servos, which has the purpose of providing attitude corrections to maintain vertical flight. A GPS chipset from SiRF has up to 40 MIPS spare processor capacity for user tasks. In addition the chipset has a low interrupt rate so real time control loops can be run. A rocket steered by a GPS flight computer can have low velocity while maintaining vertical attitude. With low velocity, low acceleration forces, and small corrective forces, the vehicle can have an optimized lightweight airframe and thin-walled tanks which significantly improves the mass fraction. The GPS flight computer moves small control surfaces using lightweight servos and logs the complete flight. The logged data is used to tune the control loops and improve the system response. The performance of a twin staged sounding rocket with a GPS flight computer for a commercial imaging application was simulated, and found to have potential to replace existing satellite imaging. The GPS flight computer also sequences vehicle recovery, and controls a parafoil to steer back towards the launch point, reducing the cost of vehicle recovery. Another advantage of reduced initial speeds is that the vehicle can be recovered without resorting to a ‘self-destruct’ capability. In addition the flight computer can shutdown the engine, dump the propellants and deploy the recovery system in the event of a fault. Sub-scale vehicles can be flown many times for little cost, rather than a few expensive full-scale test launches. This enables an improvement in overall reliability, and a cost reduction since fewer full-scale launches are required. Good results were obtained from the GPS flight computer in several sub-scale launches, the tests showed the potential of the GPS flight computer to lower the cost barriers to space. The purpose of the paper was to design and implement a flight computer to enable a new generation of sounding rockets with new applications. | |
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| Paper Number RS1-2003-8002: I-Cone® for Rapid Response and Low cost Access to Space | |
| Michael J. Cully (Swales Aerospace), Peter Alea (Swales Aerospace), Nils Gustafsson (Saab Ericsson Space AB) | |
| View/Download:Presentation | Paper | |
| Abstract: I-Cone® is an innovative approach to providing payload launch opportunities while at the same time taking advantage of the excess launch vehicle performance available with the Evolved Expendable Launch Vehicle (EELV). The genesis of the I-Cone® concept is the integration of a standard set of space vehicle subsystems into a standard conical launch vehicle adapter, in effect creating an “intelligent cone: or I-Cone. The I-Cone® is capable of providing payloads and small satellites a Fast, Frequent, Flexible and Affordable (F3A) access to space. The I-Cone® concept is designed for use with the Delta IV and Atlas V (EELV) and is compatible with Delta II and Sea Launch Vehicles. The main I-Cone® structural components are derived from flight heritage payload adapters and separation systems, developed by Saab Ericsson (SE) Space, which minimizes the development risks and production costs. I-Cone® space vehicles can be essentially transparent to the Primary payload of a typical EELV manifest. The launch site processing flow for an I-Cone® has a “no impact” approach on the standard EELV Primary payload processing flow. The I-Cone® space vehicle concept is suited for a wide variety of technology demonstration and short term operational missions. The baseline concept features typical payload resources of a 100 kg of mass, with 150 Watts of orbit average power, and a standard downlink data rate of 2.0 Mbps. The baseline I-Cone® Space Vehicle is capable of providing a pointing accuracy of 10-50 arc·sec, a propulsion system with 90 kg of mono-propellant Hydrazine, and a mission life exceeding one year. The use of I-Cone® for Low Earth Orbit (LEO) missions is emphasized in this paper, although Geosynchronous Transfer Orbit (GTO) launch can be accommodated by the I-Cone® also. The modular approach to the I-Cone® space vehicle structure permits an extraordinary level of flexibility for meeting emerging specialized launch requirements. Micro-and nano-satellites can also be accommodated in an I-Cone® variation that incorporates a dispenser. Variations on the I-Cone® dispenser theme include a passive dispenser that provides additional propulsion and attitude control after separation from the launch vehicle. The I-Cone® concept can argument the potential return on investment for any EELV launch as it provides a cost effective and flexible solution particularly for Technology demonstration missions. This paper will first present what needs the I-Cone® design addresses for access to space. This paper will also provide the generic mission requirements for the I-Cone® design, describe baseline I-Cone® implementation architecture, discuss payload accommodations and provide baseline implementation. Finally this paper will discuss potential mission designs for which I-Cone® can be applied to. This paper is derived, in part, from a study performed in Reference 1. | |
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| Paper Number RS2-2004-5000: Development and Operations of Flight Systems for Responsive Missions | |
| Michael J. Mahoney (Universal Space Lines), Layne Cook (Universal Space Lines) | |
| View/Download:Paper | |
| Abstract: Vehicle flight operations (and more particularly, the operation of a vehicle’s flight through its onboard software) must be built from the “ground up” to support responsive, safe flight at a significantly lower cost than that experienced with today’s space flight vehicles. To accomplish this, we must deviate from the current flight system paradigm. We must design and develop systems from the beginning considering the entire life cycle (how to fly it, flight software maintenance and upgrades, etc.) This includes developing flight test prototypes and test articles early in the system life cycle to gain “real” experience and determine the “real” requirements such that the ultimate design leads to the “correct” development product. Flight operations (and more specifically, the flight systems and guidance navigation and control (GN&C) operations) can be developed early in the program to support rapid test articles, rapid mission planning and rapid flight software development, without an army of people doing each function. This development “system” can then also support the operations of the flight systems with flight software maintenance and turn-around and pre, during and post flight analysis. One tool that supports this design philosophy has already been partially developed under NASA’s NRA 8-30 Space Launch Initiative (SLI) program. It is called the Integrated Development and Operations System (IDOS). IDOS is an integrating environment designed to support the flight software life cycle needs of reusable launch vehicles (per SLI goals) and by extension other aerospace vehicles. This integrated environment provides for design, development, implementation, test, validation, operation (mission planning and execution) and maintenance of advanced GN&C algorithms in a flight operations environment. Using IDOS, we have demonstrated an order of magnitude reduction in the required effort to transform an advanced GN&C “idea” (algorithm) into working flight software operating on a flight like processor in real time. | |
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| Paper Number RS2-2004-2001: Transforming National Security Space Payloads | |
| T. Ryan Space (Directorate of Development and Transformation), Vincent Deno (Directorate of Development and Transformation), Edward Jones (Directorate of Development and Transformation) | |
| View/Download:Presentation | Paper | |
| Abstract: This paper describes the benefits of and rationale for the current transformational trend in the Department of Defense (DoD) to provide direct tactical support to warfighters through development, acquisition, and operation of responsive space payloads. Of special near-term significance are efforts focusing on smalland micro-satellites and their close collaboration with responsive launch development programs, such as the Responsive Access, Small Cargo, Affordable Launch (RASCAL) and Force Application and Launch from CONUS (FALCON) programs. The Air Force has formalized a Tactical Satellite (TacSat) program, originally initiated by Vice Admiral (retired) Arthur Cebrowski of the Office of Force Transformation (OFT), to invigorate concept exploration of and experimentation with responsive, tactically focused systems. | |
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| Paper Number RS2-2004-2005: Responsive Space Requires Responsive Manufacturing | |
| Todd Mosher (USU), Brent Stucker (USU) | |
| View/Download:Paper | |
| Abstract: In order to meet the responsive space needs of the future, improvements need to be made not only to the product of satellites themselves, but also the process by which they are created. While many continue to focus primarily on product innovation, with the exception of large commercial satellite projects like communication satellites, the Boeing 601/702 platform, and the Iridium program, few are successfully addressing how do we change how we build satellites rather than changing the satellites we build. Small satellites especially offer opportunities for process innovation and the application of advanced manufacturing techniques to this process. Within the mechanical and aerospace engineering department at Utah State University, professors in space engineering and advanced manufacturing have teamed to focus on making satellite manufacture more responsive. Through a series of studies for the U.S. government, concepts to realize responsive manufacturing to enable responsive space will be discussed and reported. | |
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| Paper Number RS2-2004-3003: A Modular Design for Rapid-Response Telecons and Navigation Missions | |
| Phillip Davies (SSTL), Doug Liddle (SSTL), John Paffett (SSTL), Sir Martin Sweeting (SSTL), Alex da Silva Curiel (SSTL), Stuart Eves (SSTL) | |
| View/Download:Presentation | Paper | |
| Abstract: In order to achieve an ‘economy of scale’ with respect to payload capacity the major trend in telecommunications satellites is for larger and larger platforms. With these large platforms the level of integration between platform and payload is increasing leading to longer delivery schedules. The typical lifecycle for procurement of these large telecommunications satellites is now 3-6 years depending on the level of non-recurring engineering needed. Surrey Satellite Technology Ltd (SSTL) has designed a low-cost platform aimed at telecommunications and navigation applications. SSTL’s Geostationary Minisatellite Platform (GMP) is a new entrant addressing the lower end of the market with payloads up to 250kg requiring less than 1.5 kW power. The development of GMP was supported by the British National Space Centre through the MOSAIC Small Satellite Initiative. The main design goals for GMP are low-cost for the complete mission including launch and operations and a platform allowing flexible payload accommodation. GMP is specifically designed to allow rapid development and deployment with schedules typically between 1 and 2 years from contract signature to flight readiness. GMP achieves these aims by a modular design where the level of integration between the platform and payload is low. The modular design decomposes the satellite into three major components - the propulsion bay, the avionics bay and the payload module. Both the propulsion and avionics bays are reusable, largely unchanged, independent of the payload configuration. Such a design means that SSTL or a 3rd party manufacturer can manufacture the payload in parallel to the platform with integration taking place quite late in the schedule. In July 2003 SSTL signed a contract for ESA’s first Galileo navigation satellite known as GSTBV2/A. The satellite is based on GMP and ESA plan to launch it into a MEO orbit late in 2005. The second flight of GMP is likely to be in 2006 carrying a geostationary payload consisting of six Ku band transparent transponders. Once the platform is flight proven, SSTL will be able to offer it to commercial and institutional operators when there is an urgent need for capacity for example to introduce new services, for gap fillers, for frequency filing missions and for technology demonstration missions. | |
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| Paper Number RS2-2004-4004: Remote Anywhere: Web-Based Spacecraft Integration and Checkout | |
| Michael Chaffin (MicroSat Systems), Alan Bibbero (MicroSat Systems), Greg Hegemann (MicroSat Systems) | |
| View/Download:Presentation | Paper | |
| Abstract: MicroSat Systems Inc, (MSI) is proposing a web-based “Remote Anywhere” approach to spacecraft integration and checkout to enable the tactical, mobile launch scenarios of the next generation of micro-sat systems. In this paper we will discuss hardware and software solutions to a web-based methodology for compressing the time from identification of a “Launch on Demand” (LOD) need to “On- Orbit Operation” for a spacecraft. In our discussion we will describe the architecture and show examples from our integration experience. Prior to call-up, an inventory infrastructure will have been developed around mission sets encompassing an LOD launch vehicle, spacecraft and payload kits with qualified processes. Spacecraft environmental testing will have been previously completed on an initial qualification unit with inventory units acceptance tested so the remaining assembly tasks focus on electrical and software checkout. To support this, a modular test system approach is required. As the common bus becomes standardized so must the test system. MSI is creating a test infrastructure based on a common software re-use library. A set of software scripts developed to test a bus can be used for multiple configurations for different mission scenarios, taking full advantage of the commonality of the core spacecraft bus design. Core functionality of design in the primary spacecraft system, command & data, power and attitude determination & control is maintained from bus to bus. This “Plug & Play” approach accommodates the concept of a standardized set of ground tests that can be developed, automated, and reused with minimal, if any, modifications to blocks of code that have been written. As libraries of test scripts are developed, the time and NRE costs for future missions are dramatically reduced. A second component necessary to the LOD concept is the ability to quickly and efficiently access spacecraft command and telemetry ground systems with minimal “human in the loop” requirements. MSI is operating at the leading edge of technology by embracing a “remote anywhere” command and telemetry system. This web-enabled technology allows access to spacecraft subsystems and payloads from anywhere in the world using secure web based protocols, encryption techniques and a minimal set of ground equipment. Implementing this rapid response capability minimizes costs by leveraging the existing Internet infrastructure and common desktop tools. This gives spacecraft integration teams the ability to have a virtual test support team at hand keeping staffing, travel and personnel overhead costs to a minimum. The cost savings implied here are significant with additional capacity to handle multiple missions simultaneously and run integration operations 24/7. As testing progresses from sub-system verification to bus level and payload equipped spacecraft testing telemetry can be made available to anyone, anywhere in the world. | |
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| Paper Number RS2-2004-5002: Plug-and-Play - An Enabling Capability for Responsive Space Missions | |
| Thomas Morphopoulos (Microcosm), L. Jane Hansen (HRP Systems), Jon Pollack (HRP Systems), Jim Lyke (AFRL), Scott Cannon (USU) | |
| View/Download:Presentation | Paper | |
| Abstract: A self-organizing network concept, leveraging commercial approaches is under development to support responsive space avionics networks. The work is being done in support of an Air Force contract, including the following elements: a network manager (hardware and network medium specific component), mission manager (mission objective specific), and GN&C algorithms for a four state activity (power-on, initialize, nominal GN&C, safe). The current work emphasizes the resource manager, which is responsible for discovering resources as they come on-line. It also manages real-time data descriptions and health/status information for potential consumers of each produced element within the overall network. These mechanisms form a basic system for plug-and-play, in which the components of a system can be rapidly assembled with minimal need to write detailed, low-level code pertaining to the interface of each element. The resulting automation allows system designers to focus on design of higher-level software in an object-oriented fashion, a process that itself might be automated under this concept. | |
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| Paper Number RS2-2004-5004: Leveraging COTS Hardware for Rapid Design and Development of Small Satellites at the USAF Academy | |
| Cristin Anne Smith (USAFA) | |
| View/Download:Presentation | Paper | |
| Abstract: The purpose of the United States Air Force Academy (USAFA) Space Systems Research Program is to give cadets the opportunity to “learn space by doing space” while also providing an orbiting platform for Air Force and Department of Defense (DoD) science experiments as FalconSAT-3 is designed to do. This paper describes small satellite programs at the U.S. Air Force Academy’s Space Systems Research Center. FalconSAT- 2 and FalconSAT-3 are student-built small satellites that provide low-cost access to space for DoD space research & development payloads as well as platforms for student experiments. Rapid, low-cost design is achieved by leveraging Commercial Off-the- Shelf (COTS) hardware to the greatest extent possible. FalconsSAT-2, still searching for an alternate launch opportunity, was the first to demonstrate the use of COTS modules for this use. FalconSAT-3, scheduled to be launched in 2006, recently completed critical design and built upon the successful FalconSAT-2 experience to develop an even more capable spacecraft bus. By using the off-the-shelf equipment, student involvement in satellite and mission design has been accelerated and provided the capability to challenge students This paper is declared a work of the U.S. Government and is not subject to copyright protection in the United States. through more intense participation over the years. Realizing the rapid turnover and extended commitments of students in a senior undergraduate program, there is a delicate balance to be found; one between comprehensive mission and satellite design requirements, and adequate experience in a multi-million dollar, real world space program. The development of the FalconSAT program will first be described in the context of the progress made, followed by a more detailed discussion of the COTS hardware for more efficient development of small student satellites as simple payload platforms for educational and technological purposes. | |
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| Paper Number RS2-2004-5005: Optimizing For Responsive Space Design | |
| Terrance Yee (MicroSat Systems) | |
| View/Download:Presentation | Paper | |
| Abstract: MicroSat Systems, Inc. (MSI) has recently been challenged by new programs to radically streamline the development process for spacecraft, reducing a traditional 2-3 year small satellite development to just 12 months. This type of accelerated schedule can only be accomplished with a drastic reduction in design time to just a few months. In the following paper we present the key methods used to achieve this reduction in design time and contrast them with traditional design methods and milestones. In our discussion we will focus on requirements generation, interface design, and design documentation, especially change control and tracking. Requirements generation takes on a different emphasis when schedule forces a program to use almost exclusively existing components. In these cases existing component capabilities dictate to a large extent what can be accomplished in the mission. Thus, requirements are not so much dictated by the customer as negotiated based on what can be accomplished with the hardware at hand. Interface design is also predicated upon the use of mostly existing hardware and mature payloads. This means that the principle challenge is making hardware choices that require a minimum of modifications to work together and deciding the most efficient place to make those modifications. Often the trade space is skewed toward the fastest solution, not necessarily the most elegant, lowest mass/power or the least expensive. To minimize the documentation burden and allow the engineers to work as fast as possible on finishing the design, a number of compromises are made. Rather than run most changes through a formal change control board, the line engineers are empowered to make changes to their own areas of expertise without needing outside pre-approval as long as their changes don’t overstep boundaries into another subsystem. If changes do cross boundaries the systems lead and the leads of the affected areas have to be involved in approving the change. Change documentation is also shortened to a simple spreadsheet log of changes for most items, although formal change paperwork is still used for the more involved, systemic issues. Throughout the design process, the MSI programs employ small, focused teams which cross several organizational boundaries to make truly integrated product development teams. These teams work without regard to organizational origin in their day-to-day activities and are empowered to make decisions in real time. To do this, the team needs to communicate effectively with each member directly rather than work through intermediaries and representatives. Communication is kept informal via e-mail and impromptu teleconferences rather than organized into meeting after meeting. | |
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| Paper Number RS2-2004-6001: Guided Self-Assembly for ProtoSat Combination | |
| Gregory Brault (AFRL) | |
| View/Download:Presentation | Paper | |
| Abstract: To support the objectives of responsive space through rapid payload development, the Air Force Research Lab (AFRL) is exploring the idea of a softwarereconfigurable network of “ProtoSats” that can be arranged in three-dimensional configurations using software tools, and then constructed by a user with limited feedback given by the modules. Each ProtoSat is a standard replicable module with similar size and interconnect hardware, although unique in its payload. These ProtoSats combine to form “MacroSats”, which exploit the synergy of the ProtoSats to address mission objectives. To examine some of the infrastructure challenges associated with rapid assembly/configuration, recent AFRL work explored concepts in “Guided Self- Assembly”, modeling ProtoSats as 3”x 3” printed circuit boards with directional peripheral interconnects. This work has led to insights useful in developing detailed theories for ProtoSat design, communication, and coordination. The insights of the associated experimentation and ideas related for follow-on research are discussed in this paper | |
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| Paper Number RS2-2004-6002: Responsive Space Through Adaptive Avionics | |
| Denise Lanza (Scientific Applications International Corp.), Jim Lyke (AFRL), Paul Zetocha (AFRL), Don Fronterhouse (Scientific Simulation), Dave Melanson (Mission Research Corp.) | |
| View/Download:Presentation | Paper | |
| Abstract: This paper will review the need for an improved strategy in avionics to address responsive space objectives. General strategies will be described for achieving responsive space through reconfigurable electronics and computer-aided design. The Adaptive Avionics Experiment (AAE) is introduced as a specific embodiment of these principles, and its key elements are described. Status and future plans are discussed. | |
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| Paper Number RS3-2005-1006: A TACSAT Update and the ORS/JWS Standard Bus | |
| Jay Raymond (Office of Force Transformation), Greg Glaros (Office of Force Transformation), Patrick Stadter (APL), Cheryl Reed (APL), Eric Finnegan (APL), Michael Hurley (NRL), Charlie Merk (NRL), NRL (NRL), NRL (NRL), NRL (NRL) | |
| View/Download:Presentation | Paper | |
| Abstract: In May of 2003, the Office of the Secretary of Defense’s Office of Force Transformation (OFT) undertook an initiative to perform Operationally Responsive Space (ORS) experimentation. Two years later the first experiment, TacSat-1, is launch ready, TacSat-2 is in the integration and test phase, TacSat-3 is underway, and TacSat-4 is in the planning phase. The TacSat-3 experiment took the important step of creating a joint process for mission selection. Each experiment tests key elements needed for a truly operational system, emerging as the Joint Warfighting Space (JWS) system. A necessary element of this system is a spacecraft bus with accepted standards for interfacing with each segment of this ORS/JWS system. The OFT and Space and Missile systems Command (SMC) have therefore undertaken a four phase initiative to develop and test bus standards and then transition them for acquisition. This effort involves multiple government laboratories, industry, and academia participants. The four phases of this initiative provide steady, tangible steps to spiral warfighting capability and receive operational feedback while moving toward an acquisition. This paper discusses this standard bus initiative with emphasis on Phase 3, which is led by the Naval Research Laboratory (NRL) and Applied Physics Laboratory (APL) team. For context, this paper includes portions of the 2003 and 2004 papers and discusses the status and current challenges of ORS/JWS. | |
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| Paper Number RS3-2005-3003: AeroAstro's SMARTBus™: A Low-Cost Modular Approach Enabling Responsive Space Missions | |
| Scott A McDermott (AeroAstro), Luis G. Jordan (AeroAstro) | |
| View/Download:Presentation | Paper | |
| Abstract: The long lead and cycle times currently associated with development and launch of satellite systems has established a prohibitive environment for responsive deployment of technology and tactical capability to orbit. To address these critical deficiencies in lead time and ease of space access, AeroAstro has developed and built a modular spacecraft architecture known as SMARTBus™. SMARTBus defines systemic, mechanical, electrical, and logical (software) interfaces that allow spacecraft modules to interact with each other based on their functions rather than their implementation. One attitude determination module may be implemented based on sun sensors, another based on a star tracker, another based on a GPS; each offers different attitude determination capabilities; but from an interface standpoint, they behave the same. In this way, a mission requiring a given set of capabilities may be built up from pre-existing and pre-qualified modules offering those capabilities, and all of the modules can interact with each other because that interaction is based on providing functionality rather than controlling a specific implementation. SMARTBus challenges the traditional spacecraft systems approach by incorporating a modular bus design with “smart” software architectures. Intrinsic to the design is the “Plug-and-Sense” capability that enables the SMARTBus module stack to not only detect the presence and orientation of integrated subsystem modules, but also ascertain their function and key performance parameters. Additionally, the system utilizes a heuristic, self-interrogation approach to provide a robust means of performing configuration and diagnostics activities. This capability transcends nominal housekeeping routines to include an enhanced degree of system autonomy for both initial station acquisition and checkout, as well as mission-specific operations. This flexible functionality will enable scalable multi-mission compatibility, long shelf-life, rapid call-up and field integration for launch, intelligent built-in test capability for rapid initialization on-orbit, and variable batch manufacturability. | |
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| Paper Number RS3-2005-3005: The Little Probe that Could: Four Months from ATP to Launch | |
| Gerald Murphy (Design_Net Engineering), Ken Center (Design_Net Engineering) | |
| View/Download:Presentation | Paper | |
| Abstract: Responsive Space requires Responsive spacecraft! Last year Design_Net discussed just such a small spacecraft and examined the programmatic lessons that were learned from that development. This was the Floating Potential Probe (FPP). Flight 4A for the ISS was to deploy a new set of high voltage solar arrays which, due to large amount of electron current collected from the LEO plasma, caused the vehicle ground to shift significantly. The potential shift could lead to destructive and dangerous (for EVA) arcing. A set of plasma contactor units (PCUs) were being deployed to provide for emission of electrons collected by the arrays thus bringing the potential closer to zero and mitigating the arcing danger, however no means was in place to verify that the PCUs were working. In late July 2000, the ISS program office at JSC issued an engineering change notice that directed the development of some means to independently assess the performance of the PCU’s, and to have hardware available for launch on STS-97 (ISS Flight 4A) the very mission scheduled to deliver and install the first set of large Station solar arrays on November 30th. Such a schedule allowed only 4.5 months to design, build, test, manifest, complete EVA training, and deliver for launch. NASA Glenn Research Center (GRC), NASA Johnson Space Center (JSC), and Design_Net Engineering (DNet) formed a unique team to try to accomplish the directive. Previous papers (1,2) have discussed in detail the design of the FPP and some of the programmatic lessons that were learned which are applicable to responsive space. In addition to these lessons, a number of engineering and technological innovations are required to support responsiveness. Design_Net has expanded the lessons from the FPP responsive bus to more generic application. In this paper will focus our attention on three key aspects from a technological point of view that support rapid response. We focus on: 1) the systems engineering process and the technologies and tools that are enablers for rapid design/configuration in the early stages of a program; 2) the interaction of configurability with requirements and; 3) the role of standards and Plug and Play (PnP) capability in achieving rapid development and integration. | |
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| Paper Number RS3-2005-3006: HexPak2 - A Flexible, Scalable Architecture For Responsive Spacecraft | |
| Michael Hicks (Lockheed Martin), Michael Enoch (Lockheed Martin), Larry Capots (Lockheed Martin) | |
| View/Download:Paper | |
| Abstract: HexPak2 is a deployable space structure that provides the characteristics essential for successful responsive space missions, including ease of scalability, a geometry naturally adapted for plug-and-play architectures and multiple mission-specific component layouts, and a large deployed aperture from an optimal stowed volume. It consists of hexagonal bays that stack when stowed to efficiently use payload fairing volume, but deploy to a planar structure with deck area many times the fairing cross-section. The large deployed area to fairing size ratio supports large aperture payloads, multiple payloads, heat rejection significantly beyond traditional designs, multiple manifest with minimal wasted support mass, and easy access on orbit for expansion and flexibility for reconfiguration. Since each bay is fabricated and tested individually, and easily accessible from all sides, the time per unit mass to manufacture a complete spacecraft is greatly improved over more traditional structures. For missions that require a large number of platforms, the modular structure offers easy interchangeability of HexPak bays which makes it possible to maintain a consistent production flow even during periods of parts shortages. Standard physical interfaces also allows for commonality in tooling, fixturing, testing and ease of satellite integration. The hexagonal geometry is near optimum for taking advantage of available fairing envelopes and the folded structure is self-supporting which minimizes the need for additional structure to support launch. HexPak has proved to be ideally suited for fairing cross sections as little as 1 meter to as large as 5 meters. | |
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| Paper Number RS3-2005-4001: Key Elements of Rapid Integration and Test | |
| Terrance Yee (MicroSat Systems) | |
| View/Download:Presentation | Paper | |
| Abstract: MicroSat Systems, Inc. (MSI) is currently supporting AFRL in the Roadrunner/TacSat-2 program to demonstrate the development of a tactically useful small satellite in just 14 months. This rapid development requires a flight integration schedule that is less than 4 months between flight hardware arrival and Launch Readiness Review and includes the integration of fourteen experiments and system environmental testing. This paper will review the lessons learned so far in integration and test and the key elements for success in the rapid development process. Management of I&T activities is a critical component in rapid development. Flexible scheduling in complex missions with very short time frames is the key to efficiently using test time. To adapt to the often fluid schedule requirements of I&T, it is necessary to have several options for activities the test team can perform at any given moment. Therefore when Component A runs into problems, Components B, C, or D can be tested during the originally scheduled time with no net impact to schedule. To achieve this objective multiple items must be ready ahead of schedule and several different teams need to work in parallel preparing future tests, debugging troublesome equipment offline, and conducting tests in the test environment. Automated or computer script-driven testing adds tremendous flexibility to the process. By capturing the specific technical expertise for a particular test in the script, any reasonably familiar operator can execute the test,allowing for different experts to work on multiple items. This approach has provided the Roadrunner program enough flexibility to have as many as six different test teams each in various stages of test preparation, execution or documentation. This large number of teams allows for testing in multiple shifts and across an extended workweek, as some teams can have down time while the other teams are working. The use of scripts decouples specific people from specific tests to allow the critical debugging work to happen offline while other testing continues uninterrupted. Script-driven testing also makes handoff to flight operators easier by providing a knowledge bridge between groups. A key element to rapid testing is a streamlined documentation system that efficiently captures requirements, test knowledge, problems, standard procedures, and verification status without unduly burdening the test teams. To this end, the Roadrunner team uses a minimal set of documentation including a daily test log, a simple Problem/Failure database, an Excel verification matrix, simplified test procedures and “Test Flows”, which are brief documents that tie requirements to scripts and set the framework for the test. | |
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| Paper Number RS3-2005-4003: Software as a Tall Poll in Achieving Rapid Configuration and Integration | |
| Kenneth Center (Design_Net Engineering), Gerald Murphy (Design_Net Engineering), Robert Strunce (Star Technology Corporation) | |
| View/Download:Presentation | Paper | |
| Abstract: The concept of “configurable” spacecraft that can be quickly assembled to meet the needs of operationally responsive space has many elements that make its realization challenging. Having certain standard components “on the shelf” reduces lead times, standardizing interfaces improves the ease of integration and can lower harness complexity and cost, but how do we actually deal with the configuration process and how can we assemble the software rapidly and reliably? Software is a major cause of mission failure and without attention to this aspect of the problem, reliable responsive configuration will never be possible. Design_Net Engineering, under an SBIR contract from MDA, has teamed with Star Technologies to develop a process, an architecture, and a set of tools to be used in enabling this rapid configuration. The tools are built upon Star Tech’s Spacecraft Dynamics Tool (SDT) and the architecture of a unique modular software design developed by Design_Net on previous flight programs. The tool is a system design and configuration utility which enables precise simulation of the on-orbit behavior of the spacecraft and its payload. The architecture allows scripted operations to be authored and modeled before freezing the configuration of the spacecraft. After complete mission investigation in the virtual environment with the selected configuration, the flight code can be assembled “at the push of a button.” Portions of this program are being integrated into the capability of the flight Testbed at the Air Force Research Laboratories (AFRL) in Albuquerque, New Mexico. | |
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| Paper Number RS3-2005-4005: Standardization to Optimize Integration and Testing | |
| Norman C. Anderson (USAF), Guy G. Robinson (Jackson & Tull), David R. Newman (MicroSat Systems) | |
| View/Download:Presentation | Paper | |
| Abstract: The Air Force Research Laboratory is involved in the evolution of space systems to a model that is more responsive to operational needs. A key tenant of the concepts for future responsive space support includes the ability to rapidly integrate missionspecific payloads into flexible spacecraft busses on a very short (a few days) timeline. Traditional spacecraft integration and test (I&T) processes require several months for even the most simplistic spacecraft. An analysis of a number of spacecraft I&T processes reveals that a large percentage of the I&T schedule is expended in verification and validation of the interfaces between subsystems. Further, there is much time and money spent in the constant reconfiguration of the test support equipment to accommodate the multitude of unique interfaces used on a typical spacecraft. While these customized interfaces may allow for the development of a spacecraft optimized for minimum mass or peak payload performance, these metrics are of secondary interest in a system intended to be rapidly integrated and deployed in response to user needs. There are design approaches for subsystem interfaces used in many other industries that have the need to minimize the time required for integration and test. The automotive industry and the developers of personal computer systems are two such environments. This paper will discuss the results of analyzing the integration and test flow of typical programs, identify candidate interfaces and the projected savings to be gained by adopting them from other industries, and propose a process flow for a responsive spacecraft based on these new interfaces. | |
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| Paper Number RS3-2005-4007: Responsive Space Center of Excellence | |
| John E. Hicks (National Nuclear Security Administration) | |
| View/Download:Presentation | Paper | |
| Abstract: The objective of a Responsive Space Center of Excellence is to provide a conduit to rapidly design, build, test and field an operationally relevant microsatellite system; reduce timeliness for development, test, launch, checkout; and provide a responsive capability to the Joint Force Commander within 7-days. The National Nuclear Security Administration (NNSA) is well postured to develop and manage a Responsive Space Center of Excellence to deliver low cost satellite components and systems for Responsive Space through the application of lean manufacturing, Six Sigma tools and other business management process tools. The objective of this paper is to illustrate that the above mentioned business processes are a must for the success of Responsive Space to provide the Joint Force Commander microsatellite support in 7-days. | |
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| Paper Number RS3-2005-5001: Space Plug-and-Play Avionics | |
| Jim Lyke (AFRL), Don Fronterhouse (Scientific Simulation), Scott Cannon (USU), Denise Lanza (Space Applications International Corporation) | |
| View/Download:Presentation | Paper | |
| Abstract: The Air Force Research Laboratory is developing a system for rapidly building spacecraft based on adapting “plug-and-play” (PnP) approaches for use in space. This space plug-and-play avionics (SPA) system is based on an interface-driven set of standards intended to promote the rapid development of spacecraft busses (platforms) and payloads. As such, SPA is an open systems framework, combining commercial standards (such as USB) with carefully chosen hardware and software extensions necessary for modern real-time embedded systems (e.g. fault tolerance, higher power delivery, self-description). This paper will review the status of SPA and the efforts being made to standardize SPA through the AIAA. | |
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| Paper Number RS5-2007-6005: Net-Centric Operations and Responsive Spacecraft - A Guide to Implementation | |
| Jeffrey L. Janicik (Innoflight) | |
| View/Download:Presentation | Paper | |
| Abstract: Net-Centric Operations is clearly a significant enabler for Responsive Space. One of the key enablers that to Net-Centric Operations is the Global Information Grid (GIG). The GIG represents a globally interconnected, end-to-end set of information capabilities and processes for collecting, processing, and managing information on demand to warfighters, policymakers, and support personnel. The GIG fulfills a fundamental principle of Net-Centric Operations by securely connecting people and systems regardless of time or place, providing vastly superior situational awareness and better access to information for accelerated decision-making. At the core of the GIG is a high throughput TCP/IP network. The “I/O” to the core is data-generators and data-users. The users include secure IP network users, command centers, and field units such as infantry and armored divisions. The generators are an increasing number of sources that both front line and rear command posts desire to utilize. These sources can be spacecraft assets (imaging), UAV (imaging & offensive weapons), remote controlled platforms (Remotely Operated Vehicles, deployable surveillance) and Beyond Line of Site (BLOS) near space platforms such as balloons and high-altitude airships. Most of these sources, especially those in space, are not networked and those that are typically rely on point-to-point stovepipe communications systems. In extending the GIG to the scores of DoD users, it is important to do the same on the source side all the way to individual payloads and sensors. Once both the users (shooters) and the sources (sensors) have full and transparent connectivity, the true benefits of the GIG concept will be fully utilized and expanded beyond the terrestrial domain. This paper will describe how the aerospace industry should implement IP in the spacecraft bus design to realize this sensor-to-shooter benefit of net-centric operations while keeping cost and schedule to a minimum. The key facets to this implementation include the following: a network-enabled bus that establishes IP addresses at each desired sensor, secure IP-compatible communications security (COMSEC), unaltered use of RFC’d Internet standards, and smart adaptation of the physical layer to maximize bandwidth efficiency. | |
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| Paper Number RS3-2005-5003: Small Cell Lithium-Ion Batteries: The Responsive Solution for Space Energy Storage | |
| Chris Pearson (AEA Technology), Carl Thwaite (AEA Technology), Nick Russel (AEA Technology) | |
| View/Download:Presentation | Paper | |
| Abstract: During a solar eclipse, spacecraft rely on batteries to power all on-board electrical systems. Advances in battery technology have led to lighter products that, in turn, allow spacecraft to carry heavier and more capable payloads. AEA Technology has pioneered the current state of the art in the space community: ‘small-cell’ Lithium-ion battery technology. This paper focuses on the direct applicability, and benefits, of this approach to Responsive Space. Traditionally, space batteries consisted of a single series connected string of ‘large cells’. Large cells are sized (in terms of capacity) according to mission requirements, meaning that cell qualification programmes for individual missions are common. The small cell approach involves taking Commercially available Off The Shelf (COTS) Lithium-ion cells, qualifying them for space, and using a strict batch test and screening process to ensure the continued quality of cell batches for space flight. This obviates the need for cell qualification for each programme. The technology has proved to be ideal for small satellite missions, due to the low-cost of small cell battery designs compared to rival large cell energy storage solutions. The maturity of the design concept, and therefore low risk of utilisation, allows Protoflight programmes to be adopted for all but the most specialised of applications. A protoflight programme reduces cost due to the lack of need for a dedicated qualification battery unit and test programme. | |
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| Paper Number RS4-2006-3001: Analysis of Modular Spacecraft Bus Design for Rapid Response Missions | |
| Lucy E. Cohan (Massachusetts Institute of Technology), Richard-Duane Chambers (Massachusetts Institute of Technology), Rachel K. Lee (Massachusetts Institute of Technology), Col. John Keesee (Massachusetts Institute of Technology) | |
| View/Download:Presentation | Paper | |
| Abstract: Rapid Response spacecraft are becoming more essential due to current affairs. The long development and testing times of typical satellites necessitate a change of paradigm to accommodate responsive space timeline requirements. One necessary component of this paradigm shift is a standardized bus. A standardized bus allows for a minimal amount of bus redesign and testing for each mission. Instead of forcing the bus to conform to the payload, the payload must conform to a set of predetermined requirements imposed by the bus. By reducing the need for satellite redesign and test, standardized buses allow for mission readiness in a matter of weeks rather than years. However, using standardized buses reduces payload flexibility and leads to buses that are oversized, overdesigned, or otherwise inappropriate for a particular payload. This study proposes that a modular bus might provide standardized interfaces for responsiveness, yet still provide some flexibility to match the needs of the payload. However, modularity comes at a price, introducing inefficiencies and testing cost. This paper presents a quantitative analysis of the cost and efficiency of two competing standardized bus options: a traditional monolithic design and an emerging modular architecture. The study further attempts to quantify this tradeoff and determine the optimal degree of modularity for a responsive satellite bus. The degree of modularity is determined by specifying which, if any, subsystems should be considered as separate modules that can be upgraded or replaced, and which subsystems should be a part of an integrated bus common to every mission. The study has been undertaken using MATLAB® simulations. The individual simulation components represent various satellite subsystems, as well as satellite demand, cost, testing time, and inventory size. The codes are run to determine the efficiency, cost, reliability, response time, and inventory of each configuration of modular and integrated subsystems across a range of payloads. Specifically, this study explores payloads of the following three types: a communications payload in low earth orbit, a communications payload in highly elliptical orbit, and an optics payload in low earth orbit. The efficiency is defined as being the excess mass, power, and volume capacity created by utilizing the standardized bus that is designed to work with many payloads as opposed to a monolithic bus designed specifically for the given payload. The response time is defined as the time from the mission call to the time that the satellite is ready for launch. Additionally, there is an efficiency associated with the amount of inventory required to maintain mission readiness. The study establishes the optimal combination of modular and integrated subsystems, as well as testing strategy and inventory for responsive space missions of these types. | |
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| Paper Number RS4-2006-3003: Low-Cost Repsonsive Exploitation of Space by HAUSAT-2 Nano Satellites | |
| Young-Keun Chang (Hankuk Aviation University, South Korea), Suk-Jin Kang (Hankuk Aviation University, South Korea), Byoung-Young Moon (Hankuk Aviation University, South Korea), Byung-Hun Lee (Hankuk Aviation University, South Korea) | |
| View/Download:Presentation | Paper | |
| Abstract: This paper addresses the development and design of the HAUSAT-2 (Hankuk Aviation University SATellite-2), being developed by SSRL (Space System Research Lab.) of Hankuk Aviation University. This is the second satellite system development program executed at the university level in Korea. The HAUSAT-1, a next generation 1kg class picosatellite, has already been developed by SSRL as the first ultra-small satellite and is planned to be launched in the first quarter of 2006 by a Russian “Dnepr” launch vehicle. The development of ultra-small satellite such as HAUSAT-1 and HAUSAT-2 project offers graduate and undergraduate students great opportunities to understand the satellite design process, analysis, manufacturing, assembly, integration, test, launch and operation, as well as providing practical experience working as a team member. In addition, these ultra-small satellites can also be utilized as a space technology test bed. Main mission objectives of the HAUSAT-2 are to study the scope of activities and ecology of animals using Animal Tracking System (ATS) and collect space environment data of mission orbit from Electric Plasma Probe (EPP) as a space science payload. The secondary mission objectives are to provide the following space technology verifications: performance verification of a star tracker manufactured by SaTReC-i and a spaceborne GPS receiver manufactured by NAVICOM. The HAUSAT-2 is a nano-satellite, having a mass of 25kg with 30cm x 30cm x 39cm hexahedron configuration. It is being designed to operate in LEO with 650 ~ 800km altitude sun synchronous orbit. The three-axis stabilization is being implemented with pitch bias momentum method. The electrical power subsystem includes 8 cell Li-Ion batteries, 5, 12, and 28 volt regulators, and 5 gallium arsenide (GaAs) solar panels capable of generating more than 21.7 watts average solar power at end-of-life (EOL). Link budget analysis results allow the HAUSAT-2 communication subsystem to implement the amateur bands for uplink (VHF) & downlink (UHF) communications and 2 watt radio radiation power. The command & data handling subsystem(C&DH) includes an OBC (On-Board Computer) consisting of a MPC860T microprocessor operated by VxWorks O/S and a TCA (Telemetry & Command Assembly) with 89C50 microcontroller. The design mission life of the HAUSAT-2 satellite is expected to be 2 years. The HAUSAT-2 incorporates five types of operation modes; Initial, Normal, Science (Mission Mode), Communication, and Safe. The power requirements at individual modes are different and calculated by considering average and maximum power consumption. The critical design of the HAUSAT-2 has been completed. The STM (Structural-Thermal Model) was developed as the first system model used for verifying structural and thermal design margin. The qualification level vibration and thermal tests have been conducted on the STM. Detailed circuit design and parts selections were carried out at the module level and EM (Engineering Model) units and payloads have been manufactured in which Integrated system performance and flight software algorithm were verified through the ETB (Electrical Test Bed) tests. Box-level qualification tests were achieved to ensure required performance in launch and space environments. | |
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| Paper Number RS4-2006-3004: Bandit: A Platform for Responsive Educational and Research Activies | |
| Michael Swartwout (Washington University in St. Louis) | |
| View/Download:Presentation | Paper | |
| Abstract: There are many potential paths to improving the responsiveness of space systems. At Washington University, we are investigating three: drastically-reduced spacecraft size, drastically-reduced mission lifetime, and pre-placement of assets on-orbit. Extremely small spacecraft (under 10kg) are believed to be more responsive due to their low part count (reducing cost / time of fabrication and assembly), ease of handling/integration and increased ability to fit in the unused corners of payload fairings (i.e., as lastminute additions to already-manifested launches). Missions that last days or hours have significantly less risk of environmental degradation and need less power margin, allowing the use of less-expensive parts and/or eliminating redundant systems. The combination of small size and short mission enables such vehicles to be pre-positioned on larger host vehicles, allowing them to be activated as needed for their specific mission. From an education standpoint, very small, short-duration spacecraft are within the capabilities of an undergraduate team to design, build and operate within their “lifetime” as students. What missions – if any – can be met by such small, short-duration systems? We believe that one such mission is on-orbit servicing. On-orbit servicing (inspection, repair, refueling) is a key enabling technology for future missions, and it has “responsive” needs of its own. In 2005, both NASA and the Air Force flew demonstration servicing missions, with several more planned for the near future. Servicing missions have both ‘long-period’ functions (power generation, long-range communications, momentum management) and mission-specific ‘short-period’ functions (agile maneuvers over small distances, sensing, mechanical manipulation). The recent servicing missions described above use the same vehicle for both long-period and short-period functions, which results in a spacecraft larger than strictly necessary for servicing. Instead, we propose the Bandit concept, which splits the long-period and short-period functions between a host vehicle and a drone vehicle. Bandit has the following enabling elements: • A very small (< 10kg), maneuverable drone capable of independent (or lightly supervised) operation on 10 or more sorties lasting up to 2 hours each • A host vehicle (possibly the service recipient) with the following interfaces: - A launch containment system to carry the drone to orbit - An on-orbit docking system to allow a drone to “sleep” between sorties - A recharging (and possibly refueling) system in conjunction with the dock. - A short-range, low-power communications link to the drone This concept also creates “responsive” engineering education; early student teams create the platform and design/test infrastructure, and successive generations improve on the design. We have already seen the benefits of this approach over the past four years. At present, Bandit-C is being developed as part of the AFRL/NASA/AIAA University Nanosat-4 student satellite competition. This paper outlines the Bandit mission in more detail, including current design, prototyping activities and functional/environmental testing. Special emphasis is placed on hardware testing using a 3DOF air-bearing testbed and operations/autonomous control testing on the 6DOF software-based simulator. Design of the 25 kg host spacecraft Akoya is also discussed. We conclude by presenting sample missions for future Bandits. | |
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| Paper Number RS4-2006-3005: ORS Phase III Bus Standards Status | |
| J. Christopher Garner (US Naval Research Laboratory), Michael Hurley (US Naval Research Laboratory), Gurpatap S. Sandhoo (US Naval Research Laboratory), Eric J. Finnegan (Johns Hopkins University/Applied Physics Laboratory), Patrick A. Stadter (Johns Hopkins University/Applied Physics Laboratory), Brian Kantsiper (Johns Hopkins University/Applied Physics Laboratory) | |
| View/Download:Presentation | Paper | |
| Abstract: The U.S. Naval Research Laboratory and the Johns Hopkins University/Applied Physics Laboratory are collaborating with many industry partners to write bus standards for responsive spacecraft buses as part of the ORS/JWS Phase III. The next Phase, Phase IV led by SMC, will use the standards as input to the procurement of responsive spacecraft buses in 2008. More than 8 industry partners (Spectrum-Astro, Design-Net, Swales, Orbital, Raytheon, Loral-Microcosm, and Microsat Systems Inc) are under contract to NRL to participate in the integrated systems engineering team (ISET). The ISET has been meeting since June 2005 and has produced the first drafts of the payload developers guide (PDG) and the bus standards documents. Currently, an NRL/APL team is working to develop a prototype spacecraft bus to mature portions of the standards and supply the spacecraft bus for the TacSat 4 mission. This paper will discuss the ISET team process in developing the bus standards and the progress of experimentation with the prototype bus. Phases I-III of this effort are funded by OSD’s Office of Force Transformation, Phase IVeffort will be funded by SMC. | |
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| Paper Number RS4-2006-4003: Development of the Tactical Satellite 3 for Responsive Space Missions | |
| Thomas M. Davis (AFRL), Capt. Stanley D. Straight (USAF AFRL) | |
| View/Download:Presentation | Paper | |
| Abstract: Numerous Department of Defense studies show implementing a responsive satellite capability provides for significant military utility to augment or surge current space capabilities. The TacSat concept explores the capability/technological maturity of small, low-cost satellites with the most prominent efforts currently being conducted within the Science and Technology (S&T) Program. In addition to providing for ongoing innovation and demonstration in this important technology area, these S&T efforts also help mitigate technology risk and establish a concept of operations (CONOP) for future acquisitions. TacSat efforts underway by the Air Force Research Laboratory (AFRL) and the Naval Research Laboratory (NRL) are focused on demonstrating small (<500kg), operationally responsive, low-cost satellite and launch capabilities to support warfighter. AFRL’s Space Vehicles Directorate is leading the Tactical Satellite 3 (TacSat-3) team and partners include Space and Missiles Center Detachment 12, the Army Space Battle Laboratory, the Air Force Space Warfare Center, the Office of Naval Research, and the DoD Office of Force Transformation. Building on the experiences with TacSats 1 and 2, TacSat-3’s mission was vetted through a formal payload selection process with Air Force Space Command (AFSPC) and Combatant Commands (COCOMs). TacSat-3’s mission was selected for specific capabilities to meet user needs, and to demonstrate those capabilities within cost and schedule constraints. A building block for Operationally Responsive Space, TacSat-3 will experiment with a Hyperspectral Imaging (HSI) capability direct to the tactical warfighter within 10 minutes of a collection opportunity. The TacSat-3 demonstration features a low cost “plug and play” modular bus and low cost militarily significant payloads – a Hyperspectral Imager and a secondary payload demonstrating data exfiltration provided by the Office of Naval Research. TacSat-3 will demonstrate evolutionary steps and traceability towards objective system goals for the capabilities and processes including rapid response to a user defined need for material detection and identification, and battle damage assessment. Additionally, it will demonstrate traceability to enable launch processing at the launch base faster than 7 days. Finally, it will feature a rapid development of the space vehicle and integrated payload and spacecraft bus by using components and processes developed by the Operationally Responsive Space Modular Bus program. Design constraints established for the TacSat-3 program include a total program cost to be less than $50M, to fit on a low cost responsive space booster and a satellite weight of less than 400 kilogram, with a build time for payload and modular bus of less than 18 months. The TacSat-3 CONOPS breaks old paradigms and gives COCOMs first realistic opportunity for responsive, dedicated space capabilities at the operational and tactical level. The TacSat-3 spacecraft will collect and process images and then downlink material ID text and geolocation or downlink full data image using a Common Data Link. An in-theater tactical ground station will have the capability to uplink tasking to spacecraft and will receive full data image. | |
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| Paper Number RS4-2006-5003: Reconnaissance Payloads for Responsive Space | |
| Stanley Kishner (Goodrich Optical and Space Systems Division), David Flynn (Goodrich Optical and Space Systems Division), Charles Cox (Goodrich Optical and Space Systems Division) | |
| View/Download:Presentation | Paper | |
| Abstract: A key mission fueling the interest in Responsive Space is optical reconnaissance. Minimizing the cost and delivery schedules of optical reconnaissance payloads having true operational capability will be key to success of these missions. Modification of existing proven airborne reconnaissance payloads provides a practical path for achieving this Responsive Space capability. In addition to space sensors such as the Multispectral Thermal Imager developed for Sandia Laboratories and launched in early 2000, Goodrich currently provides a range of imaging sensor systems and services for airborne reconnaissance. Goodrich has provided the reconnaissance cameras for the U-2 since 1957, with the current SYERS electro-optical system providing a robust set of outputs supporting IMINT and MASINT missions. The capabilities of the SYERS system have continually improved through our P3I program. From a low earth orbit of 300 kilometers, a SYERS system modified for use in space could provide a ground sample distance of approximately 1-meter. It is this system and its functional elements that form the basis for our Responsive Space Reconnaissance (RSR) approach. The Goodrich approach for producing payloads for RSR can be visualized as pulling from our “product stream” of airborne sensors to build an inventory of RSR payloads that can be made available upon short notice. The major effort for adapting the SYERS sensor system for responsive space is associated with the focal plane and electronics. Retaining the current operational functionality and architecture could be implemented with parts and processes compatible with a short lived vacuum environment and aimed at reducing the power consumption for compatibility with the low-cost spacecraft buss. Our vision for RSR Payloads is to establish a pre-positioned, rapid-response process that can adapt our continually evolving product line of high acuity airborne sensors for responsive space missions as the need for such missions is identified. In this paper we will describe the SYERS sensor, its modification for use in space and interfaces to candidate spacecraft. We will also address the CONOPS that will allow a modified SYERS sensor to meet responsive space needs. In summary, optical imaging payloads for Responsive Space can be evolved from our operationally-proven line of tactical and strategic airborne sensors, which have demonstrated on-demand support to our warfighters. These existing airborne systems emphasize operational availability and can be readily adapted for RSR missions. This philosophy and capability is directly aligned with Responsive Space needs. | |
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| Paper Number RS4-2006-6002: Java-based Plug-n-Play (Flight) Control Systems for Responsive Space | |
| Constantine Orogo (Lockheed Martin Advanced Technology Center), Michael Enoch (Lockheed Martin Advanced Technology Center), Donald Flaggs (Lockheed Martin Advanced Technology Center) | |
| View/Download:Presentation | Paper | |
| Abstract: A major challenge to achieving a usable and useful “6-day spacecraft” for Operationally Responsive Space is the ability to rapidly compose the system to perform both the needed mission- and spacecraft-oriented functionality using the available "Plug-N-Play" (PnP) spacecraft components. Physical assembly of the PnP spacecraft components is a necessary, but insufficient condition for achieving a fully realized operational system. The assembled system needs to provide the functional capabilities to support the intended mission and also needs to provide the functional capabilities to ensure the operational health and safety of the resulting spacecraft. A preliminary service-oriented spacecraft architectural model to provide a reusable infrastructure is under development as part of the AFRL Responsive Space Testbed effort. The Lockheed Martin ATC is pursuing the development of a Java-based distributed architecture environment that supports this service-oriented, reference spacecraft architectural model. This work draws on the research experience at the LM ATC since the mid-1990s directed towards the problem of performing multi-fidelity, “composable” simulations, with the ultimate objective being the ability to simulate the entire life cycle of a space system. A key component of this approach involves a simulation architecture that is based on spacecraft services, much like the service-oriented models now widely used in the consumer marketplace. It is but an evolutionary step to extend this approach from simulation to operations. To seamlessly span the entire range from simulation to operations, a single vertically integrated software architecture was needed. The Java-based distributed architecture provided such an environment with its evolution from desktop, to enterprise, to mobile devices, and now to real time systems. The Java environment addresses the complexity needed for operational simulations and ultimate deployment for integrated spacecraft flight and payload control systems. The Real Time Specification for Java (RTSJ) supports hard real time, soft real time and non-real time processes all interoperating within the same virtual machine. Initial prototyping is being done using IBM’s Real Time Java (RTJ) implementation of the RTSJ. Along with the development of the Java platform came the development of a multitude of supporting APIs, and one in particular, the JINI protocol, which supports the operation of dynamically changing networks of distributed services (and devices). Using JINI, running as a non-real time process within IBM’s RTJ, provides the rich set of Plug-N-Play capabilities needed to demonstrate both automatic configuration, as would be needed for I&T, as well as for operational fault tolerance and reconfigurability needed for on-orbit operations. | |
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| Paper Number RS4-2006-6003 alt: Implications of Responsive Space on the Flight Software Architecture | |
| Jonathan Wilmot (NASA Goddard Space Flight Center) | |
| View/Download:Presentation | Paper | |
| Abstract: The Response Space initiative has several implications for flight software that need to be addressed not only within the run-time element, but the development infrastructure and software life-cycle process elements as well. The run-time element must at a minimum support “Plug & Play”, while the development and process elements need to incorporate methods to quickly generate the needed documentation, code, tests, and all of the artifacts required of flight quality software. Very rapid response times go even further, and imply little or no new software development, but using only pre-developed and certified software modules that can be integrated and tested through automated methods. These elements have typically been addressed individually with significant benefits, but it is when they are combined that they can have the greatest impact to Responsive Space. The Flight Software Branch at NASA’s Goddard Space Flight Center has been developing the run-time, infrastructure and process elements needed for rapid integration with the Core Flight software System (CFS) architecture. The architecture consists of three main components; the core Flight Executive (cFE), the component catalog, and the Integrated Development Environment (IDE). This paper will discuss the design of the components, how they facilitate rapid integration, and lessons learned as the architecture is utilized for an upcoming spacecraft. | |
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| Paper Number RS4-2006-7002: Responsive Payload Accommodations and Integration Operations for Dedicated CubeSat Missions | |
| John M. Garvey (Garvey Spacecraft Corporation), Dr. Jordi Puig-Suari (California Polytechnic State University), Lori Brooks (California Polytechnic State University) | |
| View/Download:Presentation | Paper | |
| Abstract: A key factor to achieving responsive space operations is the availability of standardized payload accommodations that can simplify integration tasks and reduce costs. Several such standards are beginning to emerge in the very small end of the payload market that is characterized by the so-called CubeSat class of spacecraft. These also happen to be compatible with proposed nanosat launch vehicle (NLV) concepts that are intended to enable dedicated CubeSat missions that are free from the operational constraints associated with traditional secondary payload manifest opportunities. The Poly-Picosat Orbital Deployer (P-POD) under development by California Polytechnic University, San Luis Obispo (Cal Poly SLO) is one such system that is now transitioning to flight status. The viability and merits of such dedicated CubeSat missions was highlighted recently during flight testing of the Prospector 7 prototype reusable launch vehicle (RLV) that was developed by Garvey Spacecraft Corporation (GSC) and California State University, Long Beach (CSULB)). In this case, an engineering prototype of the P-POD unit manifested and then deployed a set of three simulated CubeSats twice within a period of just 3.5 hours. The entire program, from authority to proceed through launch, took only six months, as compared to lead times that are measured in years for larger launch systems. Future plans envision extending the operational environments that the P-POD will be evaluated under as the NLV development program transitions to higher-performance test vehicles. Besides continued evaluation of refined payload accommodations and integration techniques, it is anticipated that future CubeSat payloads will help monitor and characterize NLV payload environments. Throughout this endeavor, students from Cal Poly SLO, CSULB and other participating academic institutions will continue to gain valuable experience with flight hardware integration and responsive launch operations. | |
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| Paper Number RS5-2007-2003: Spacecraft Slewing/Guidance Algorithm for Hyper Spectral Imagers | |
| Morris Frayman (Star Technologies Corporation), Robert R. Strunce Jr. (Star Technologies Corporation) | |
| View/Download:Presentation | Paper | |
| Abstract: Hyper Spectral Imagers (HSI) are planned as payloads on future Responsive Space missions. Typically, the mission requirements dictate a specific scanning rate that the spacecraft must support in order to body point the HSI such that a single pixel covers a specific Ground Sample Distance (GSD) size in one integration frame. The actual focal plane array is typically one pixel wide by several hundred pixels long. The frame rate (integration time) is the time to scan the length of the array in one GSD. All too often the requirements are stated as: a fixed GSD size (~3 meters); a fixed altitude (~400k); a fixed scanning rate which is usually constrained parallel to the spacecraft trajectory. Even for a nadir target, these specifications severely compromise the quality of the image and overly constrain the image swath which could lead to minimal utility to the war fighter on the ground. And in particular, following these requirements will produce increasingly poor image quality the more the target is off-nadir. While working on the Navy Research Laboratory’s (NRL) Naval Earth Mapping Observatory (NEMO), a generic guidance algorithm for body pointing its HSI was developed that overcomes these restrictions and provides a high level of image quality regardless of target location or swath direction with respect to the image swath orientation. This algorithm was implemented and simulated for a TacSat3 spacecraft. A unique spacecraft slewing algorithm was developed (including sun avoidance) that rapidly moves the spacecraft into the proper orientation and settles the HSI LOS jitter at the beginning of imaging swath. The guidance algorithm is based on a spacecraft line-of-sight (LOS) coordinate frame and a target coordinate frame at the target site. The target frame is defined as a function of the Image swath to be collected. The HSI telescope relationship with the spacecraft coordinate frame is defined with respect to the spacecraft LOS frame. The relationship between the LOS and target frames guarantees that the HSI scan line or focal plane remains perpendicular to the scan direction independent of the target swath orientation with respect to the spacecraft trajectory. The focal plane speed over the ground (SOG) is calculated based on the HSI parameters: output pixel angle; HSI camera frame rates; and the size of the GSD which is based on look angle to the target and the earth’s geoid. The slewing and guidance algorithms are developed in this paper and simulation results are presented. | |
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| Paper Number RS5-2007-3002: ISET ORS Bus Standards and Prototype | |
| Tom Doyne (OSD's Director of Defense Research and Engineering), Patrick Stadter (Johns Hopkins University Applied Physics Laboratory), Cheryl Schein (Johns Hopkins University Applied Physics Laboratory), Eric Finnegan (Johns Hopkins University Applied Physics Laboratory), Steve Vernon (Johns Hopkins University Applied Physics Laboratory), Paul Schwartz (Johns Hopkins University Applied Physics Laboratory), George Moretti (SMC Space Development Group), Kellie Turner (SMC Space Development Group), Gurpartap S. Sandhoo (Naval Research Laboratory), Mark Johnson (Naval Research Laboratory), Mike Hurley (Naval Research Laboratory) | |
| View/Download:Presentation | Paper | |
| Abstract: A necessary element of an Operationally Responsive Space (ORS) system is spacecraft bus standardization allowing increased modularity for a rapid, tailored response of the space element. Advancing sound and accepted bus standards is the objective of the OSD’s ORS Bus Standards Initiative. This effort involves multiple government, industry, and academia participants assembled into an Integrated System Engineering Team (ISET). The core ISET industry team members include AeroAstro, Boeing, Design Net Engineering, General Dynamics, Loral, Microcosm, MicroSat, Orbital, Raytheon, and Swales. Government and FFRDC team members include the NRL, JHU/APL, SMC, AFRL, MIT/LL, SMDC, and Space Dynamics Lab. The ISET generates recommended standards for ORS Spacecraft and uses two prototype bus builds (including one that is open to the ORS community and specify implemented for the development of these standards) to evaluate and mature the bus standards. The ISET has recently made their second major release of the bus standards documents which are available at this conference. This ISET team is also complemented by an open membership Business Team who provides business case factors for consideration in the standards definition, as well as for input to the acquisition transition plan. This paper describes the status of the ORS Bus Standards developed by the ISET to date including prototype lessons learned. The RSC papers #2005-1006 and #2006-3005 provide additional history and context for this paper. | |
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| Paper Number RS5-2007-4004: Responsive Spacecraft Bus Implementation for Unique HEO Missions Based on Standard Interfaces | |
| P.A. Stadter (Johns Hopkins University Applied Physics Laboratory), C.S. Schein (Johns Hopkins University Applied Physics Laboratory), M.T. Marley (Johns Hopkins University Applied Physics Laboratory), C.T. Apland (Johns Hopkins University Applied Physics Laboratory), R.E. Lee (Johns Hopkins University Applied Physics Laboratory), B.L. Kantsiper (Johns Hopkins University Applied Physics Laboratory), B.D. Williams (Johns Hopkins University of Applied Physics Laboratory), E.D. Schaefer (Johns Hopkins University of Applied Physics Laboratory), S.R. Vernon (Johns Hopkins University of Applied Physics Laboratory), P.D. Swartz (Johns Hopkins University of Applied Physics Laboratory), E.J. Finnegan (Johns Hopkins University of Applied Physics Laboratory), J. Chris Garner (Naval Research Laboratory) | |
| View/Download:Presentation | Paper | |
| Abstract: This paper will provide details of the implementation of an Operational Responsive Space spacecraft bus to be used by the TacSat-4 CommX mission in a highly elliptical orbit (HEO). Through this discussion, two primary themes of the RS5 conference will be addressed: applications that lend themselves to solution by small spacecraft in HEO orbits and implementation of a critical element of the overall mission within the context of responsive capabilities. Specifically addressed will be the challenge of developing a spacecraft bus as a platform designed to a set of defined interface standards, while faced with the unique requirements of a particular payload. This will include a discussion of the driving requirements for the bus to provide operations in the HEO orbital environment and the user applications that can take advantage of such a platform given candidate payloads. The paper will provide details on design and implementation decisions that were made to accommodate standards, and places where proposed standards were not able to be addressed for the particular implementation. The technical details included will provide insights into the bus implementation well after Critical Design Review, but prior to space vehicle Integration and Test, thus system designs will be mature and near completion. | |
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| Paper Number RS5-2007-5005: Generic Nanosatellite Bus for Responsive Mission | |
| F.M. Pranajaya (University of Toronto Institute for Aerospace Studies), Robert E. Zee (University of Toronto Institute for Aerospace Studies) | |
| View/Download:Presentation | Paper | |
| Abstract: To achieve responsiveness, satellites must be developed quickly and launches must be available on demand. In order to develop satellites quickly, an obvious approach is to use a common bus that can be customized to fit a range of missions. However, designing a common bus that is both capable and suitable for rapid development and production is a challenge given the diversity in mission requirements. The common bus must achieve the most stringent requirements across all missions. The Generic Nanosatellite Bus (GNB) from the Space Flight Laboratory (SFL) at the University of Toronto Institute for Aerospace Studies (UTIAS) meets the common-bus challenge for a wide range of nano-class missions (up to 12 kg and potentially bigger), and provides a platform for state-of-the-art, high-performance applications not previously achievable with nanosatellites. A typical GNB is a 20 cm cube and consists of one ARM7 housekeeping computer, two ARM7 computers for attitude/propulsion control and payload operations, CMOSimagers, a power system with Triple-Junction cells and Lithium-ion batteries, passive thermal control, UHF uplink, a 32-256 kbps S-band downlink, and a 1 arc-min three-axis attitude control system consisting of tiny reaction wheels, sun sensors, magnetometer and star tracker. A cold-gas propulsion system may also be added to the GNB. The GNB is currently being used to implement the BRight Target Explorer (BRITE) space astronomy mission and the CanX-4&5 dual satellite formation-flying mission at SFL. These demanding missions illustrate the diversity in application for the GNB. In order to facilitate rapid launches, SFL has adopted an approach to build customizable separation systems for any nanosatellite. These separation systems can be integrated with the satellites prior to launch site delivery and hence make launch coordination easier. The SFL “XPOD” separation system interfaces the GNB-based spacecraft to practically any launch vehicle. Spacecraft up to 12 kg may be accommodated in existing XPOD designs. This paper describes the GNB, its available subsystems and their performance, the XPOD family of customizable separation systems, and the upcoming science and technology-demonstration missions in 2008 and 2009 that will make use of the GNB technology. | |
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| Paper Number RS6-2008-1001: Numerical Optimization Study of LEO to LEO Aeroassisted Orbital Transfer for Small Satellites | |
| Arthur Scherich (University of Florida), Anil V. Rao (University of Florida), Skylar Cox (MicroSat Systems), Todd J. Mosher (MicroSat Systems) | |
| View/Download:Presentation | Paper | |
| Abstract: Operationally responsive space (ORS) is an area of growing interest to the U.S. space community. ORS refers to the ability to enhance capability, increase flexibility, and reduce execution time of operational spacecraft. A desirable capability for ORS is the development of spacecraft that can accomplish multiple distinct missions by having the ability to change its orbit. Designs for orbital transfer generally fall into one of two categories: all-propulsive transfers (where the orbit is changed completely using on-board fuel) or transfers that combine propulsive maneuvers with atmospheric flight maneuvers (where a portion of the orbital transfer is accomplished using propulsion while the remainder of the orbital transfer is accomplished using aerodynamic control via flight through the atmosphere). The latter category of orbital transfer is called aeroassisted orbital transfer. In the case of small satellites, the on-board fuel constraints will render all-propulsive maneuvers infeasible for many missions, thereby requiring the use of atmospheric flight maneuvers. Thus, it is important to study the problem of aeroassisted orbital transfer for ORS. Optimal aeroassisted orbital transfer for high-mass lifting bodies has been studied extensively (but never flown) over the past several decades. In these studies, several types of aeroassisted maneuvers, such as aerocruise with propulsive maneuvers and aeroglide without propulsive maneuvers, have been discussed. It has been found that the heating rate constraint is one of the key parameters in determining the performance of the aeroassisted orbital transfer (i.e. the sustainable heating rate directly affects the amount of inclination change that can be achieved by the aeroassisted maneuver) and the overall mission cost (i.e. the amount of fuel required for the mission). Due to the complexity of the atmospheric maneuvers and the need for performance (e.g. minimization of fuel), aeroassisted orbital transfer problems are often posed as optimal control problems. Moreover, because these optimal control problems cannot be solved analytically, it is necessary to obtain solutions using numerical methods. Numerical methods for solving optimal control problems fall into two general categories: indirect methods and direct methods. The merits of these two approaches will be discussed in this paper. In recent years a new class of direct methods that have shown promise in the numerical solution of optimal control problems are orthogonal collocation or pseudospectral methods. In an orthogonal collocation method, the state is approximated using a basis of polynomials. Several different orthogonal collocation methods exist including the Legendre pseudospectral method, the Chebyshev pseudospectral method, the Radau pseudospectral method, and the Gauss pseudospectral method. In this research we are interested in applying the Gauss pseudospectral method to the problem of low-Earth orbit (LEO) to LEO aeroassisted orbital transfer. In this paper accurate numerical solutions are presented to the problem of LEO to LEO aeroassisted orbital transfer for a small spacecraft with constraints on inclination change, heating rate, and total heat load is considered. The spacecraft is chosen to be of a size that can be launched on a modern day small launch vehicle (e.g. Falcon or Minotaur). Furthermore, we consider orbit transfers where the size, shape, and line of apsides of the terminal orbit are constrained. In particular, the constraint on the line of apsides makes it possible to locate the apogee of the orbit over a strategic point on the Earth for intelligence, surveillance, or reconnaissance (ISR) purposes. The aeroassisted orbital transfer problem is posed as a three-phase nonlinear optimal control problem and is solved using the software GPOCS9 which is a MATLAB® implementation of the aforementioned Gauss pseudospectral method. The optimal trajectories obtained in this study provide insight into the possibilities that these types of orbits could provide ORS missions in the future. | |
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| Paper Number RS6-2008-2005: Rapid Assembly of Spacecraft Structures for Operationally Responsive Space | |
| Roopnarine (Honeybee Robotics Spacecraft Mechanisms Corporation), Shazad Sadick (Honeybee Robotics Spacecraft Mechanisms Corporation), Irene Yachbes (Honeybee Robotics Spacecraft Mechanisms Corporation), Brandon Arritt (Air Force Research Lab/Space Vehicles Directorate) | |
| View/Download:Presentation | Paper | |
| Abstract: A goal of the Operationally Responsive Space thrust is to enable turn-around of a tactical satellite within six days, from mission call-up to on-orbit operation. The ability to produce such a spacecraft would be a vast departure from the norm of large, complex, costly custom spacecraft that require a period of years to deploy. While progress has been made in developing shorter timescale, more modular small satellites, every aspect of the spacecraft development process needs to be reassessed in order to achieve the goals of a truly responsive tactical space program to rapidly meet the tactical needs of the warfighter. Assembly, integration and test typically account for 6 months to 2 years of the spacecraft production cycle. This process could be drastically reduced by stocking component-ready modular panels for assembly. Even with the pieces of a spacecraft bus and payload prepared for integration, the assembly of the structure itself needs to be sped from the typical process of securing panels with dozens of mixed-size fasteners and the associated verification, tooling, and documentation. Likewise, assembly of the structure also must take into consideration the need to pass electrical and thermal connections across panels of the bus. A rapid method for providing a stiff mechanical attachment across panels of a spacecraft bus while simultaneously providing electrical and thermal continuity would help to further realize the goals of ORS. It will also be crucial to demonstrate quick disassembly of bus panels in order to swap out faulty components, accommodate upgrades or support last-minute component changes to satisfy changing mission needs. In collaboration with the Air Force Research Laboratory/Space Vehicles Directorate, Honeybee Robotics Spacecraft Mechanisms Corporation has developed a fastening strategy for enabling rapid assembly of a spacecraft bus structure using our patented Quick Insertion Nut (QIN) technology. With this approach, a standard bolt can be rapidly inserted into the QIN and then about one turn is required to preload the connection, without significant support equipment or operator skill. These QINs are embedded in manifolds which reside at each edge inside the spacecraft bus (the manifold includes panel-to-panel electrical interconnects) that together comprise a skeleton for the spacecraft panels. When the panels are assembled to the manifolds, a robust structural, electrical and thermal connection for the bus is achieved. The QIN resembles a standard bolt joint typically used in aerospace applications and therefore benefits from using a similar analysis. Our approach is conducive to a quick, reliable electrical connection and passive thermal conduction between panels. While the concept is simple, when extrapolated across the multiple fasteners in a typical spacecraft bus (the time for threading of each bolt alone is eight to ten times faster), this results in a revolutionary decrease in the amount of time required for spacecraft assembly. This approach therefore represents a paradigm shift in spacecraft development, helping to enable ORS. View the video of our feasibility demonstration using a subscale prototype at: http://www.honeybeerobotics.com/roop/Assembly4.mov - a representative “corner” of three satellite panels assembled using the QIN fasteners, manifold and electrical connection. | |
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| Paper Number RS6-2008-3006: Vehicle Based Independent Tracking System (VBITS): A Small, Modular, Avionics Suite for Responsive Launch Vehicle and Satellite Applications | |
| Edmund Burke (Space Information Laboratories, Inc.), Edwin Rutkowski (Space Information Laboratories, Inc.) | |
| View/Download:Presentation | Paper | |
| Abstract: Space Information Laboratories, Inc. (SIL) Vehicle Based Independent Tracking System (VBITS), is a pioneering, open stackable, modular bus architecture that allows for use of state of the art technologies in GPS receivers, Inertial Measurement Units (IMUs), computing, communication, and Radio Frequency (RF) transmission technologies for direct and/or beyond line of site data retrieval applications. The VBITS is a multipurpose avionics technology that can be used to satisfy multiple aerospace vehicle applications like GPS metric tracking, autonomous flight termination systems, and space-based range. It can also be used for satellite bus systems and payload experiment data and control applications. VBITS can be used in conjunction with a space-based range through LEO/GEO satellites and autonomous flight safety techniques that have great potential to reduce user and DoD test range costs and can be used to meet Operational Responsive Space (ORS) objectives and requirements. The VBITS open stackable modular technology allows quick unit repair. The VBITS unit has internal sensors and diagnostics, and can be remotely monitored with a portable computer prior to flight. The VBITS design also allows for ease of manufacturability and system/unit level testing. Five VBITS production units were fully qualified and acceptance tested to Minuteman III levels based on requirements in the Eastern/Western Range GPS RCC 324-01 and EWR 127-01. The VBITS units are currently going through full qualification and acceptance testing to United States missile levels for the DoD. A tailored GPS RCC 324-01 document has been developed in coordination with 30SW Range Safety office that defined environmental test requirements for use at all DoD test ranges. One year of GPS Simulation runs were performed at Applied Physics Laboratory in coordination with SIL, flying off many COTS GPS Receivers capable of uninterpolated 20Hz position, velocity and time and raw GPS data for downlink. Test results and GPS receiver lessons learned from the many APL GPS simulator runs for a wide body missile with two patch antennas are presented. SIL is currently working with the Air Force Research Laboratory’s Space Vehicles Directorate to interface VBITS with additional ORS aerospace vehicle avionics, launch vehicles, experiment payload, and satellite plug-and-play applications. The goal is to offer a modular avionics system to new launch vehicles and to modify VBITS technology for small satellite applications. | |
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| Paper Number RS6-2008-4006: Standardization Promotes Flexibility: A Review of CubeSats’ Success | |
| Alexander Chin (California Polytechnic State University), Roland Coelho (California Polytechnic State University), Lori Brooks (California Polytechnic State University), Ryan Nugent (California Polytechnic State University), Puig-Suari (California Polytechnic State University) | |
| View/Download:Presentation | Paper | |
| Abstract: This paper will focus on the continuing development of the Poly Pico-satellite Orbital Deployer or P-POD, and its leading role in standardizing CubeSat satellite development. The paper will reflect on past mission successes and look to the future capability and promise of larger standards for access to space. Cal Poly’s role on the standardization of CubeSats through the P-POD will be explained. The initial creation of the P-POD was driven by the need for consistency in pico-satellite development. The P-POD protects the launch vehicle and the primary payload as well as the CubeSats, and is compatible with many launch vehicles, making integration repeatable and cost-efficient. The P-POD can accommodate pico-satellites as long as they meet the 1kg and10x10x10cm dimensional CubeSat standard. Mass producing a stock deployment device creates reliability in flight heritage and decreases design, manufacturing and testing costs. The P-POD provides a framework for developers to design around, and enforces adherence to the CubeSat specification. In turn, the P-POD is designed with the capability to integrate onto multiple launch vehicles. The advantages of this system are most evident in creating flexibility for CubeSat developers to launch on multiple rockets as secondary payloads. Since most satellite manufacturers must coordinate directly with the launch vehicle provider, CubeSats can find it difficult to find launches as secondary payloads. The P-POD can group multiple CubeSats to provide a competitive basis for launch as a viable secondary payload. This has allowed CubeSat developers to develop their system without a preset launch. A review of the P-POD flights over the past 5 years, and an outline of future launches consistently show the value of regulations and the benefits of flexibility. One of the main keys to the success of the CubeSat Program has been its strict adherence to the initial standard. Cal Poly, NASA Ames, and other organizations are looking to incorporate similar standards to larger satellites in an effort to bring low-cost access to space for a wider range of spacecraft. These efforts will utilize the efficiency of the P-POD and will incorporate outside influence in developing future standards. CubeSats provide a unique flexibility in the aerospace industry opening up quicker and cheaper mission opportunities than ever before. In addition, the research at the CubeSat level offers a unique paradigm shift in design operations. This means that the structure and hardware are designed first, while software development comes second. Ultimately missions can focus on meeting the standard and developing satellites and not on launch logistics and integration. | |
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| Paper Number RS6-2008-4003: Responsive Spacecraft Bus Implementation for HEO Missions Designed to Bridge Prototype and Operational Systems | |
| P.A. Stadter (Johns Hopkins University Applied Physics Laboratory), M.T. Marley (Johns Hopkins University Applied Physics Laboratory), C.T. Apland (Johns Hopkins University Applied Physics Laboratory), C.T. Apland (Johns Hopkins University Applied Physics Laboratory), R.E. Lee (Johns Hopkins University Applied Physics Laboratory), B.D. Williams (Johns Hopkins University Applied Physics Laboratory), E.D. Schaefer (Johns Hopkins University Applied Physics Laboratory), P.D. Schwartz (Johns Hopkins University Applied Physics Laboratory), B. Kantsiper (Johns Hopkins University Applied Physics Laboratory), E. J. Finnegan (Johns Hopkins University Applied Physics Laboratory), W. Raynor (Naval Research Laboratory), G.S. Sandhoo (Naval Research Laboratory) | |
| View/Download:Presentation | Paper | |
| Abstract: This paper will provide details of the implementation, integration, and test of an Operationally Responsive Space spacecraft bus to be used by the TacSat-4 CommX mission in a highly elliptical orbit (HEO). It will provide details of near-term results of the development, integration and test, and the implementation of standard interfaces to facilitate a bridge between prototype and operational systems. The paper details the means by which the technology and system engineering inform further use for future operational systems, including specifically the work of the Integrated System Engineering Team of industry, laboratory, and government participants. Details of the qualification of the spacecraft up through completion of integration and test will be provided, as will lessons observed that specifically translate into information for future operational satellite builds. This will include a discussion of the driving requirements for the bus to provide operations in the HEO orbital environment and the user applications that can take advantage of such a platform in a timely manner given candidate payloads. Aspects of complementary analyses of how similar operationally responsive space systems can military utility will also be presented. | |
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| Paper Number RS6-2008-4004: Design and Use of a Variable Thermal Layer (VTL) for Rapid Satellite Component Intergration | |
| William Hafer (Infoscitex Corporation), Nicholas Vitale (Infoscitex Corporation), Chris Macris (Enerdyne Solutions), Robert Ebel (Enerdyne Solutions), John McCullough (Enerdyne Solutions), Andrew D. Williams (Air Force Research Laboratory, Space Vehicles Directorate) | |
| View/Download:Presentation | Paper | |
| Abstract: Operationally Responsive Space (ORS) requires the design and assembly of small tactical satellites in greatly reduced timeframes. This capability can be achieved with a generic plug-n-play satellite bus implementing modular structural, electronic and thermal interfaces for payload and supporting components. Modular thermal interfaces are particularly difficult to implement, due to the wide range of thermal characteristics of spacecraft components. To address this need, Infoscitex is developing the Variable Thermal Layer (VTL), a modular interface component for insertion between the satellite bus and a range of critical components. The VTL functions as a thermal gasket that is inserted between the bus and the component's baseplate. The thermal behavior of the VTL can be varied to allow precise control of thermal flux into or out of the component. VTL is implemented as an array of thermo-electric devices (TEDs) embedded in an otherwise insulating matrix, such as an MLI or aerogel blanket layer. Each TED can actively pump heat in either direction, either warming or cooling the component. Maximum heat loads on the VTL occur during the spacecraft hot cycle, when heat must be removed from the component down the thermal gradient and into the bus. Working with the thermal gradient allows the TEDs to operate efficiently, at a coefficient of performance (COP) of 5 or greater, meaning that the heat is removed from the component is five times the power supplied to the TEDs. By applying active thermal pumping over baseline conduction, the VTL can achieve “effective” thermal conductivities ranging from a minimum of 10 W/m2-K, up to a maximum of 700 W/m2-K. This performance range addresses many spacecraft components relevant to a tactical satellite bus. Some components whose thermal requirements exceed VTL capabilities, such as some electromagnets, can be integrated with the use of a thermal doubler or similar mechanism to spread the thermal load. The distributed nature of the TED array allows the VTL to conform to the hot-spot distribution of a given component, as well as matching the variation of the component’s thermal requirements in time due to changes in operating mode and orbital position. The footprint of the VTL (L, W) is sized specifically to each component. All other attributes of VTL are unchanged from component to component and spacecraft to spacecraft. | |
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| Paper Number RS6-2008-4005: Spacecraft Functional Sensitivity Study | |
| Tim Havard (Advatech Pacific, Inc.), Mark Sutton (Advatech, Inc.), Deganit Armon (Advatech, Inc.), Jerry Sellers (Rocket Science Solutions, Inc.) | |
| View/Download:Presentation | Paper | |
| Abstract: This paper describes the results of a spacecraft functionality study aimed at quantifying the effects of subsystem-by-subsystem mass and power reductions. To guide future investments in space vehicle research, Air Force Research Laboratory’s Space Vehicles Directorate, Kirtland Air Force Base, N.M. commissioned a study to analyze the impact, at the system-level, of breakthrough technology advances at the subsystem-level. The overall goal was to determine the best path for achieving an overall reduction size, weight and power of responsive-space class satellites over current state-of-the-art. To this end, engineers at Advatech Pacific, Inc. applied detailed spacecraft system modeling tools to assess the effects of reducing the mass and power of each subsystem and on the overall system mass and power. These effects were analyzed using real-world data on two current AFRL tactical satellites. Results indicate that most system mass/power reduction effects follow logically from the allocated percentage of mass/power for each subsystem. Furthermore, as payload is typically a large percentage of system mass and power, breakthroughs in payload technology could achieve large bus and overall spacecraft mass and power reductions. However, significant reduction in overall system SWAP could only be achieved by reducing the mass/power of more than one subsystem/payload simultaneously. However, to best leverage these potential technology advances, and guide the selection of new ones, rigorous systems engineering should focus on cross-subsystem functionality. | |
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| Paper Number RS6-2008-6001: Spacecraft-to-Spacecraft Subsystem Modularization for Operationally Responsive Space | |
| Manny Nimelman (MDA), Andrew Allen (MDA ), Catherine Erkorkmaz (MDA), Andrew Ogilvie (MDA) | |
| View/Download:Presentation | Paper | |
| Abstract: This paper discusses the need for a modular and reconfigurable architecture for rendezvous, proximity operations, docking, and spacecraft servicing operations to support Operationally Responsive Space (ORS) missions. The motivation for ORS is be able to respond quickly to new satellite needs, either through rapid exploitation of existing assets in orbit, or through rapid assembly and launch of satellites with requisite payloads while exercising the bare minimum of commissioning and check-out of the integrated flight configuration. Many ORS missions will have tactical needs that include situational awareness, inspection, and close-range proximity operations with other spacecraft, and even docking and servicing operations to support upgrades of existing satellites and assets in support of ORS mission operations. These missions have a spectrum of requirements for fault tolerance, imaging capability, range-to-target, accuracy, and cost, but they can share a common set of sensor systems, avionics, physical interfaces, and numerous other spacecraft subsystems. As a specific example, ORS would greatly benefit from a common modular and reconfigurable architecture permitting users to select payloads from a common pool of plug-and-play elements to support rendezvous, proximity operations, inspection (passive or active), and/or docking. The paper discusses the requirements of an ORS reconfigurable architecture, focusing on the modules required for this particular set of operations. The starting point for this architecture is a spacecraft infrastructure that includes standardized interfaces for power, data, video, and standardized operating protocols for architecture elements. MDA has identified necessary elements for this architecture which include standard satellite navigation and attitude control sensors including GPS, startrackers, IMUs/IRUs, range and pose sensor solutions, imaging systems, fault-tolerant avionics, and physical interfaces for docking and servicing. The architecture is extensible to include different configurations of a given sensor, with a baseline solution and a set of plug-and-play upgrades based on required sensor capabilities. This paper also presents technologies developed through MDA programs that support this architecture and identifies areas requiring further technology development. MDA capabilities combine demonstrated spaceflight experience in the design of high reliability, mission critical systems with experience in decoupling system functions in order to package avionics and sensor systems as modular Orbital Replacement Units (ORUs) and in designing standardized, non-proprietary interfaces. These capabilities can address both the ORS objectives of quickly integrating new satellites, and being able to adapt existing on-orbit assets. In addition, MDA’s experience in the development of many high-TRL sensor solutions promising benefits for ORS missions will be used to illustrate the ORU design process. Finally, the paper addresses a concept for a high fidelity plug-and-play simulator that would allow for rapid prototyping and straightforward integration and test of different payload configurations. Such a simulator would allow a set of launch-ready subsystems to be integrated, verified, and launched in minimal time. | |
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| Paper Number RS6-2008-5006: Payload Design Criteria for the DoD Space Test Program (STP) Standard Interface Vehicle | |
| Chris Badgett (Space Development and Test Wing), Mike Marlow (Space Development and Test Wing), Hallie Walden (Ball Aerospace & Technologies Corporation), Mike Pierce (Ball Aerospace & Technologies Corporation) | |
| View/Download:Presentation | Paper | |
| Abstract: The Space Development and Test Wing (SDTW) of the Space and Missile Systems Center (SMC) is midway through the development of a new means of spaceflight access for the science and technology (S&T) community. The goal is to make available to the entire space community a standard spacecraft (SC) to payload (PL) interface on which to base PL designs and enable access to space in a shorter timeframe, with less cost and reduced risk. Rather than designing a unique SC for each payload; the STP Standard Interface Vehicle (SIV) is a recurrent SC bus with adaptable interfaces to accommodate a range of payloads. The SC will accommodate one to four payloads totaling up to 60 kg mass and 100 watts orbit average power mounted to an external payload interface plate. The space vehicle is designed for orbits ranging from 400 to 850 km and inclinations of 0 to 98.8 degrees. The program offers a Payload User’s Guide which defines the mechanical, thermal, power and data interfaces to help facilitate PL design and integration. This paper focuses on the PL design criteria to meet the standard interface and the adaptable capabilities of the SC to perform a variety of low earth orbit (LEO) missions. | |
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| Paper Number RS7-2009-2002: Developing Autonomy Systems in ORS Timescales | |
| George Cancro (John Hopkins University Applied Physics Lab) | |
| View/Download:Presentation | Paper | |
| Abstract: Multiple methods exist for the development of on-board autonomy systems for spacecraft. Methods include rule-based systems, table-driven systems, scripting, and more advanced systems such as automated planning and model-based autonomy. Each of these autonomy development systems has unique advantages and disadvantages, but all of these systems fail to meet the timeliness requirements imposed by ORS. In order to develop autonomy systems in ORS timescales, a high-level user must be able to rapidly construct the autonomy system, rapidly test the autonomy system to ensure that the system will react correctly to faults, and then integrate the autonomy system into the spacecraft without interfering with other assembly processes. This paper describes a new development system, called ExecSpec, which rapidly develops and tests autonomy systems for Tier 2 or 3 ORS spacecraft. ExecSpec enables a high-level user to visually create spacecraft autonomy systems by drawing diagrams representing desired behavior or rapidly assembling an autonomy system from a diagram library. Once the diagrams are assembled, the autonomy system can be rapidly tested using visual stimulation of the diagrams by the user or through model checking, an advanced technique that performs an exhaustive search to find counter-examples where the diagrams violate requirements. Following testing, the diagrams are then loaded directly to the spacecraft via command into a generic, mission-independent, on-board interpreter. This enables a flexible method of integrating autonomy into the spacecraft process flow. This feature also allows a user to change diagrams post launch to support ORS Tier-1 activities or to modify the spacecraft functionality to work around post-launch issues. During operations, mission controllers can monitor execution of the system by viewing the design diagrams, which are animated according to telemetry from the on-board interpreter. Since the same diagram context is preserved from design through operations, mission operators can suggest changes directly to the design diagrams rather than writing change requests that have to interpreted and implemented by someone not involved in operations. This paper will contrast the ExecSpec system with current autonomy development methods in relationship to ORS development timescales. The paper will then describe the current state of the ExecSpec system which is currently at TRL 5. The paper will also detail a timeline for developing an autonomy system for an ORS mission with ExecSpec, showing each step of the process and how long it will take. Finally, this paper will demonstrate the process of modifying the same autonomy system in-flight. | |
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| Paper Number RS7-2009-3001: Leveraging the First ORS Mission into ORBCOMM and the Implications for Future ORS Missions | |
| Todd Mosher (Sierra Nevada Corporation Space Systems), H. ‘Lad’ Curtis (Sierra Nevada Corporation Space Systems) | |
| View/Download:Presentation | Paper | |
| Abstract: On December 16, 2006, Tacsat-2 launched from Wallops Island, Virginia atop a Minotaur rocket to become the first Operationally Responsive Space (ORS) Satellite. After one year of operation it went beyond its 6 months of design life to make several notable accomplishments. First, Tacsat-2 was launched in 36 months after both an orbit and launch vehicle change during System Integration and Test. Second, it demonstrated command uplink and image downlink to a mobile tactical ground unit, collected color and panchromatic images from space with less than 1 m resolution, and provided a unique ELINT sensor demonstration from space. Third, it was operated in a different way with commanding of the spacecraft via the Internet, on-board autonomous orbit maintenance software calculating the orbit and thruster firings, and on-board autonomous task planning and execution. This mission was enabled by a capable spacecraft that supported fourteen successful payloads/experiments on one 361 kg spacecraft for a 58% payload mass fraction. It also had several key technologies including on-board image processing, the first U.S. Hall Effect thruster flown in space with new data on in-situ performance, the demonstration of a low power transceiver, and a thin-film solar array deployment demonstration. It has been recognized by Aviation Week and Space Technology Magazine with its Small Company Product Breakthrough Award and the American Institute of Aeronautics and Astronautics with its Space Systems Award. The success of this program has been parlayed into the ORBCOMM replenishment program (OG-2). As outlined by ORBCOMM, the current order for 18 satellites with a possible option for 30 more is a proactive approach to replacing the currently flying constellation of 29 low-orbiting satellites that provide two-way messaging service. As stated publicly by ORBCOMM officials on December 4, 2007 ORBCOMM expects to pay about $6.3 million for each of the 18 satellites and to launch them between 2010 and 2012. The similarity of SNC’s response to these cost and schedule requirements to the goals of ORS is not a coincidence. In fact, if anything the programmatics are more aggressive than many of the ORS goals and as they are successfully accomplished they should cause ORS to revise the expectations of what is possible. MicroSat Systems also won one of the four ORS BAA-3 awards to study the next generation of ORS satellites that move beyond the Tacsats to have more operational capabilities. With its partners it has developed innovative strategies to make these satellites better respond to ORS requirements such as urgent needs with the ability to be built rapidly and ready for launch in a depot like operation. While this has required some new approaches, it also has been influenced by Tacsat-2 heritage and lessons learned as well as commercial practices that come out of the ORBCOMM program. In this paper, Sierra Nevada Space Systems (the new entity created through the merger of MicroSat Systems and SpaceDev) approach to learning from the first ORS mission, Tacsat-2, and parlaying this into the ORBCOMM program will be discussed. Similarities of the ORBCOMM single payload, low cost, launch-on-demand mission to ORS objectives will be discussed and the implications of what the success of the ORBCOMM program may have for future ORS missions will be projected. | |
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| Paper Number RS7-2009-3002: ISET Compliant Modular Multi-Mission Space Vehicle Design Feasibility for Rapid Response | |
| Steven Schenk (Comtech AeroAstro, Inc.), Stanley Kennedy (Comtech AeroAstro, Inc.) | |
| View/Download:Presentation | Paper | |
| Abstract: Comtech AeroAstro, Inc. and its teammates recently completed a successful space vehicle architecture preliminary design study to determine the feasibility of designing a modular, low earth orbit multi-mission space vehicle that satisfies a variety of ORS missions while minimizing the large overhead typically associated with integration and test operations. One of the most innovative aspects of this study was the ability to accommodate a wide variety of payloads and launch vehicles through the use of a modular and flexible spacecraft bus that had strict compliance with the recently established Integrated Systems Engineering Team (ISET) standards contained in the General Bus Standard (GBS). Using the standard ISET payload and launch vehicle interfaces, the spacecraft bus requires minimal redesign or reconfiguration regardless of mission selection, allowing for rapid integration and test measured in the number of days or weeks versus number of months or years for typical spacecraft timelines. The modular spacecraft design also implements a space plug-and play avionics (SPA) architecture that allows quick reconfiguration of spacecraft bus components and flight software with little or no redesign or retest required. This approach enables the utilization of efficient and effective Chile Works depot operations currently envisioned by the ORS office for storage and integration of spacecraft components to support quick reaction urgent needs. Another innovative aspect of this study is that a standard/modular bus, in order to meet the variety of missions, payloads, and launch vehicles, will have various inefficiencies or “overdesigns” in numerous areas. A perfect example of this is the structure mass fraction, in which a standard/modular bus will have a higher mass fraction versus specific mission point designs in that the structure needs to meet different launch vehicle frequency requirements while supporting generous payload mass and center of gravity variances. However, the findings of this study show this specific example as acceptable as the majority of missions considered had adequate mass margin in relation to launch vehicle mass capability to orbit. Another example is designing a spacecraft for all low earth orbit inclinations between 0° and sun synchronous. Low inclinations drive solar array sizing and attitude control design (especially with torque rods and reaction wheels), which are “overdesigned” in a sun-synchronous orbit. A third innovative aspect of this study was the successful design of a modular imaging payload that allowed for numerous optical payload missions (VNIR, MWIR, SWIR, HSI) utilizing a common core telescope assembly, allowing for rapid payload reconfiguration at any point in the space vehicle integration and test timeline. We will discuss, in detail, many pertinent results from this study, including the advantages and disadvantages of designing a spacecraft and payload that meets a large number of missions, mission orbits, launch vehicles, and mission utility, compliance with the ISET standards (and notable deviations as our design was not 100% compliant with the ISET standards), novel approaches for minimizing the integration and test logistics footprint and timelines, and the feasibility of incorporating the space plug-and-play architecture into the spacecraft design. | |
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| Paper Number RS7-2009-3003: Autonomous Instrument Checkout and Calibration Built In Test | |
| Brian A. Bauer (Johns Hopkins University Applied Physics Lab) | |
| View/Download:Presentation | Paper | |
| Abstract: An ORS mission design strategy is most constrained by the time and budget allocated for assembling and launching the asset. In order to achieve the team performance required to achieve all goals within the resource constraints, many required activities must be automated. A set of software tools that automate sensor checkout and calibration on the ground and in orbit would greatly reduce the time and personnel required to perform these costly tasks. Furthermore, emphasis is generally placed on reducing the time required to design, integrate and launch an asset; while the time required to perform initial operations checkout and sensor calibration will also delay the date at which the spacecraft can be used for its intended purpose. This paper will begin with a brief discussion of the tests and activities performed to checkout and certify imagers on the ground and in orbit. In general, the test procedures for individual instruments have been tailored to that instrument; therefore, emphasis will be placed on the similarities between the procedures in order to define a common set of core tests. Using these core tests, the next section will discuss automation techniques which may prove useful to performing these activities. In this discussion, we will examine the costs of developing these tools and compare them to the estimated savings in manpower and schedule. Additional attention will be paid to the computational cost of performing the checkout activities onboard in relation to current technology. Concluding discussion will recap the benefits and challenges of automating the checkout and calibration suite and discuss a roadmap for developing these tools. | |
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| Paper Number RS7-2009-3004: A New Paradigm for Responsive Space Missions | |
| Bill Jackson (Sierra Nevada Corporation Space Systems) | |
| View/Download:Presentation | Paper | |
| Abstract: SpaceDev (a subsidiary of Sierra Nevada Corporation) built and delivered the Trailblazer microsatellite as the first Operationally Responsive Space (ORS) “Jumpstart” mission, which was intended to demonstrate rapid assembly, integration, test, and launch processes. This “Jumpstart” mission was a multi-pronged effort to fly responsive payloads on the SpaceX Falcon 1 Flight 003 launch vehicle, which launched in August 2008 from Kwajalein Atoll in the Marshall Islands. This launch opportunity became available because the original payload for this launch had been de-manifested. The spacecraft was assembled, integrated, and tested at SpaceDev’s Poway, California facility. The rapid call-up time on this mission presented a number of difficult technical problems; including the absence of several critical long-lead items, a late flight radio change, a late requirement for encryption, lack of a ground station and mission operations center, and lack of any mission operations procedures. The SpaceDev Trailblazer team adopted an extremely aggressive “skunk works” approach that used a small, empowered, multi-disciplined team to meet difficult technical and schedule challenges. The SpaceDev team was able to demonstrate unusual flexibility and responsiveness by tailoring engineering processes to meet the demanding schedule. SpaceDev not only successfully delivered the Trailblazer satellite on budget and on schedule, but also developed a Mission Operations Center in Poway, and fielded much of the Ground Station equipment on Kwajalein. The Trailblazer satellite was launched on August 2, 2008. Unfortunately, a technical problem with the Falcon second stage separation sequence resulted in catastrophic failure of the launch vehicle, and the Trailblazer satellite did not achieve orbit. Despite the unfortunate launch vehicle failure, the Trailblazer program nevertheless made a number of significant accomplishments: ? Responsive spacecraft build and test (4 months) ? Responsive spacecraft-to-launch vehicle integration (< 1 week) ? Responsive Ground Station development ? Responsive Mission Operations Center development ? Responsive contracting and administration ? Successfully demonstrated an end-to-end launch call-up within 7 months of standing up the ORS office This paper will detail some of the many technical and programmatic challenges of this fast-paced program, and will discuss how SpaceDev was able to ultimately deliver a fully-functional spacecraft to ORS in just four months. | |
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| Paper Number RS7-2009-3005: Synergy through Diversity: The Benefits of Applying the Lessons from Microspace to Achieving the Goals of Responsive Space | |
| Ray Zenick (Comtech AeroAstro, Inc.) | |
| View/Download:Presentation | Paper | |
| Abstract: The need for fast response tactical military intelligence space systems can no longer live with the 3–5+ years currently needed for space vehicle fabrication, integration, and launch. Urgent mission requirements demand turn-around times for space missions to be drastically reduced at the bus, subsystem, and component level in order to make Responsive Space a reality. Comtech AeroAstro (CAA) has been working to achieve quick turn around in the development cycle of complex subsystems and smaller more efficient satellite busses for completely different reasons: to satisfy the customer’s needs, improve our return on investment, reduce our cost to produce, and deliver our products on timelines far quicker than typical of the aerospace industry. Whether the latest requirement is a transponder, star tracker, radiation-tolerant GPS, software defined radio, or filling the need for a highly complex anti-spoofing waveform generator, we are beginning to overcome long design cycles with components cleverly designed to allow product versatility and multiple uses in the harsh environment of space. CAA’s approach to meeting the needs of these two different missions while still providing innovative, rapid, and cost-effective products is to design for the moment, design for versatility, and design with off-the-shelf components. In most cases, these simple rules preclude the use of the more typical technologies such as ASIC, MMIC designs or LSI because of their cost and generally long development cycles. So the challenge for small aerospace companies and their engineers is to cope with not applying the latest technology. Several successful recent designs have sought to downplay the prototypical radiation hard microprocessor that is inherently insensitive to radiation and single event upsets, and is dependent on an operating system, and replace these time consuming designs with those designed around the off-the-shelf FPGA. Through clever applications of today’s FPGA-based controllers and pipe line processors embedded in our subsystem designs, we are able to offer relatively high radiation tolerance, programmable versatility, and miserly power consumption for complex yet compact designs in a minimum amount of time. Utilizing this strategy, we are able to implement sophisticated functions and processes for space situational awareness and space intelligence applications in months rather than years. In most of these cases, utilizing the FPGA state machine or pipe line processor opens the door to the quadrature digital processing techniques required of today’s software defined radio, image processor, or autonomous system demanded by the emerging responsive space aspect of our industry. When we put these goals together, Responsive Space becomes a great opportunity for small aerospace companies while at the same time acting as a very positive enabler for achieving the goals of the ORS Office. This paper will detail how numerous advanced subsystems and components for space applications have been implemented with minimum development time, thereby filling the need for responsive in Responsive Space. | |
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| Paper Number RS7-2009-3006: PCPMU: A Modular, Multi-Use Payload Electronics Architecture for Affordable, Responsive Missions | |
| Sasa Trajkovic (MDA Corporation), George Tyc (MDA Corporation), Kenneth James (MDA Corporation), Daniel Schulten (MDA Corporation), Peter Allan (MDA Corporation), Ed Ahad (MDA Corporation), Richard Allen (MDA Corporation), Simon Ladouceur (MDA Corporation) | |
| View/Download:Presentation | Paper | |
| Abstract: Compact, multi-use, software-reconfigurable electronic modules enable affordable, highly responsive missions. These modules will permit rapid assembly, alignment and integration of electro-optical or RF payloads with a standard bus. A conventional Payload Electronics (PE) assembly usually encompasses a multitude of custom designed electronic units, based on space qualified parts and procured from different suppliers, leading to long lead time, high cost, and labor-intensive assembly and validation process. The Payload Controller, Processor and Memory Unit (PCPMU) challenges the conventional approach to PE by incorporating a modular architecture utilizing highly integrated, multi-purpose components (based on reconfigurable, state-of-the-art FPGA technology) and standardized interfaces. This approach introduces a scalable, multi-mission capability providing for rapid, low-cost PE assembly, test and integration that results in a reduction in complexity, mass and volume, and a corresponding increase in reliability. This paper describes the PCPMU architecture which has been developed to support optical, SAR and communication missions. The key low-cost enablers, along with results of the current development, are presented. | |
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| Paper Number RS7-2009-3009: Thermal Subsystem Design Methodology for Responsive Space Missions” | |
| M. Eric Lyall (Air Force Research Laboratory, Space Vehicles Directorate), Andrew D. Williams (Air Force Research Laboratory, Space Vehicles Directorate), Derek Hengeveld (Purdue University), Quinn Young (Utah State University, Space Dynamics Laboratory) | |
| View/Download:Presentation | Paper | |
| Abstract: The traditional approach to satellite design is a customized and highly optimized satellite bus. The primary design driver is to minimize mass but often at the expense of time and money. To meet the goals of Operationally Responsive Space (ORS), the satellite must be adaptable to different missions, changing threats, and emerging technologies. One of the subsystems that will be challenging for the development of robust and modular architectures is the Thermal Control Subsystem (TCS). To design the TCS, virtually every aspect of the mission, the satellite, and the components must be known. The overall goal of the engineer is to reduce the mass of the system by trading cost and engineering time. As a result, every design is unique and requires extensive design, modeling, analysis, and test programs. One philosophical approach to achieve the goals of responsive space in the near term is to separate the design and engineering of the payload from the bus. The bus would have a standard design providing a specific set of baseline capabilities and would have limited upgradeability. The disadvantage with most standardized bus development programs is that the bus eventually becomes obsolete and must be completely redesigned as new technologies are developed. One of the goals of the ORS program is the development of technologies that provide robust and flexible bus designs. The Space Avionics Plug-and-Play (SPA) system in development by Air Force Research Laboratory, Space Vehicles Directorate addresses the software and electrical interfaces, but other efforts are needed to address the mechanical and thermal interfaces. For responsive space, the ideal TCS would be modular and robust to accommodate the wide range of orbits, components, and payloads with minimal survival heater power. In addition, the design and assembly time must be dramatically decreased. The ultimate goal would be a TCS with an inherent plug-and-play capability. One hindrance is that the missions, payloads, and requirements for ORS are still somewhat nebulous. As a result, bus architectures and specific components have not been identified, which makes it difficult to derive even initial thermal system requirements. To provide a baseline for the TCS design and to help bound the problem for the development of thermal plug-and-play systems, the range of external and internal heat loads for small satellites are evaluated. From this analysis, the worst hot and cold cases are identified. Using these two cases, various thermal control architectures are evaluated and a one-size-fits-most solution methodology is developed. | |
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| Paper Number RS7-2009-4001: GIST: Our Strategy for Globalizing and Internationalizing ORS Standards and Technology | |
| Robert D. Pugh (Think Strategically, LLC), Doug Harris (Operationally Responsive Space Office), James C. Lyke (Air Force Research Laboratory, Space Vehicles Directorate), Denise Lanza (Science Applications International Corporation) | |
| View/Download:Presentation | Paper | |
| Abstract: The Operationally Responsive Space mission depends on innovative ways to acquire components, satellites, and launch vehicles; assemble and test space systems; and perform launch and mission operations. Many US allies are already focused on developing and exploiting small satellites and streamlined mission operation strategies for civil and defense applications. From the beginning, the ORS Office and AFRL have had the goal of collaborating with our allies to develop a single set of responsive space standards and enlisting their abilities and expertise in realizing the ORS vision. Toward that end, AFRL has initiated the “Globalize and Internationalize [ORS] Standards and Technology” or GIST program to develop and document the legal foundation and establish an international team to participate in the development of ORS standards. GIST has developed a strategy that will establish a collaborative environment to allow government, industry, and academic entities in the US and multiple ally countries to work together and forge the set of standards necessary to enable the responsive spacecraft assembly and test that is a keystone of the ORS vision. The strategy incorporates two different approaches, one that addresses DoD requirements for international technology development activities, and one that authorizes the exchange of so-called “export-controlled technical data” between US non-defense entities and their foreign counterparts (industry, academia, and civil space organizations). The full paper describes how this international approach aligns with the ORS Office plans for further standards development; how the GIST approach ensures ITAR-compliance by employing existing treaties, export licensing processes, and defense international project agreements; and how the program will increase opportunities for international collaboration in developing ORS components and provide export opportunities for the US space industry. | |
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| Paper Number RS7-2009-4002: A Roadmap for Responsive Software Systems | |
| Ed Birrane (Johns Hopkins University Applied Physics Laboratory), Brian Bauer (Johns Hopkins University Applied Physics Laboratory) | |
| View/Download:Presentation | Paper | |
| Abstract: Responsive systems provide dynamic operational benefits within short timeframes and tight cost constraints. Fundamental to this concept is engineered adaptability which is atypical of many heritage systems. This is achieved by reducing the coupling between components in a system – a modularizing approach to system decomposition that creates discrete functional units that can be recombined during construction, deployment, or on orbit. Practically, this capability requires an infrastructure investment to ensure that multiple domain vendors build interoperable and re-usable systems. In the hardware domain this infrastructure work is underway, but similar infrastructure construction must be commenced for software. Flight software is a critical component of any re-usable system design as hardware components are difficult to alter once integrated and nearly impossible to alter post-launch. Software is the mechanism through which spaceborne capabilities are extended. While there are certain guidelines for success – open-architectures, open standards, modularity, and cross-vendor interoperability – there has been little work in understanding the specific enablers necessary to actually construct reusable software. This level of modularity requires a reasoned, vendor-neutral approach to interoperable, re-usable software systems. Lacking this, subsequent integration issues risk failure in meeting response times despite the presence of modular hardware. Our research has identified five critical enablers for flight software systems: software certification procedures, low-level architectures and frameworks, systems-level architectures and patterns, integration environments, and an evolving software library. It is notable that there are dependencies between these enablers: building a software re-use library without common test/certification procedures will result in code that is far less re-usable across missions. This paper further defines these enablers, the milestones necessary to mature each one, a brief review of the state of the industry relevant to these milestones, and recommended priorities in the maturation of these technologies. The goal of this work is to publish a framework for consideration by performers in the responsive space community to more rapidly converge on software-adaptable capabilities. | |
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| Paper Number RS7-2009-4004: A Novel Spacecraft Standard for a Modular Nonsatellite Bus in an Operationally Responsive Space Environment
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| David Voss (Boston University) | |
| View/Download:Presentation | Paper | |
| Abstract: A truly modular satellite design must be approached from a systems perspective in order to meet the Operationally Responsive Space (ORS) community’s goal of a rapidly responsive spacecraft. In such a design approach, the modular mechanical bus system must complement a modular data bus system. Through the University Nanosatellite V competition, Boston University and Taylor University have built a comprehensive nanosatellite bus where each subsystem has followed a standardized electrical and mechanical interface. This standard is based off of the CubeSat concept but expanded to nanosatellites. It allows for instrument and subsystem designers to know the mechanical and electrical interfaces their instruments or subsystem must conform to prior to a mission, providing for ease of reuse for subsequent missions. Examples from the recently constructed Boston University Student Satellite for Application and Training (BUSAT) are given to illustrate the proposed standard and its capability for a rapid response. | |
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| Paper Number RS7-2009-5001: P-n-P Attitude Control System for Responsive Space Missions | |
| Frederick Leve (University of Florida), Vivek Nagabhushan (University of Florida), Norman Fitz-Coy (University of Florida) | |
| View/Download:Presentation | Paper | |
| Abstract: The past several Responsive Space Conferences (RS1-RS6) have developed and debated the concept of responsive space missions and the potential utility of small satellites to carry out such missions. Typical missions for responsive satellites will include earth disaster management, earth monitoring, directional communication and imaging. Systems and technologies that would augment the functionality of small satellites and extend their utility to these missions need to be developed. Also these system components should be available in prepackaged units with standardized interfaces so that they can be quickly configured into satellites that best meet the requirements of the responsive mission. One such key component discussed in this expose would enable rapid retargeting and precision pointing (R2P2) of small satellites thus enabling them for responsive missions. This feature (R2P2) is an absolute necessity for small satellites since they typically are in low earth orbits, have smaller swath, and larger angular speeds. The paper presents a fully developed “black boxed” attitude control system (ACS) using control moment gyroscopes (CMG) prepackaged into a ½ U1 plug-n-play device with standard electrical and mechanical interfaces. A CMG-based ACS is well suited for the three axis attitude control of R2P2 small satellites since they have the capacity to generate large torques, are capable of high precision, and consume relatively low power. A comparison justifying the choice of CMG over reaction wheels, magnet coils, and other types of actuators is detailed in the paper. Due to their low inertia small satellites are more susceptible to perturbations due and thus pose a challenge for attitude control systems. The analysis of the dynamics of the CMG itself presents a challenge, as the effects of bearing friction, gimbal accelerations, variable inertia, and perturbations due to flywheel eccentricities on the performance of the system can no longer be neglected. Packaging of the hardware into a compact space under acute mass constraints poses mechanical design challenges. The paper discusses in detail the design and approaches adopted in overcoming these challenges and developing a fully functional plug-n-play attitude control device for small satellites. The paper also discusses control strategies and steering logics for improving the performance of the device, some simulations and experimental test results. | |
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| Paper Number RS7-2009-5002: Building SPA PnP Satellites | |
| Donald Fronterhouse (PnP Innovations, Inc), Maurice Martin (AFRL RVSE) | |
| View/Download:Presentation | Paper | |
| Abstract: The space community, led by AFRL, started developing spacecraft plug and play concepts and standards in 2004 and has resulted in the Space Plug and play Avionics (SPA) Standards. AFRL has undertaken two efforts in small satellite development to both solidify the technology and to demonstrate the benefits. The Plug and Play Satellite (PnPSat) utilizes the SPA-S interface standard and demonstrated that rapid development, integration and testing is possible. The second effort is PnPSat-2 that uses the next generation of SPA components for a larger bus focused on ORS needs to make real the promise of custom performance at commodity prices. The SPA standard interface has proven critical to the development of design tools that both select (based upon performance requirements) and place (based upon restrictions such as mass and power balance) components. The Satellite Data Model (SDM) method of query and discovery enables the development of modular, single purpose applications that support autonomous flight software in a distributed computing system. The utilization of a data centric architecture (as opposed to component centric) insolates software developers from both specific hardware components and data network topology. The SPA standard interface reduces the need for many specialized test methods resulting in major reductions in test time. This paper will present the steps used in designing, building, and testing SPA PnP satellites. | |
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| Paper Number RS7-2009-6003: The 7-Day Solution: How ORS Will Answer The Rapid Call-up Challenge | |
| Charles J. Finley (DoD Operationally Responsive Space Office), Apoorva Bhopale (Millenium Engineering and Integration Company) | |
| View/Download:Presentation | Paper | |
| Abstract: The Operationally Responsive Space (ORS) Office leads the community in the development of enabling capabilities providing assured space power focused on the timely satisfaction of Joint Force Commanders’ needs by employing current systems and deploying or developing new systems to augment and replenish capabilities. Our efforts are focused on achieving three categories of solutions. Tier-1 solutions use or “employ” existing or on-station capabilities to provide highly responsive space effects through the employment/modification/revised application of existing, fielded space capabilities. The targeted time period for application of Tier-1 solutions is immediately-to-days from the time at which the need is established. Therefore, these solutions focus on existing ground and space systems, operations, and processes. If a Tier-1 solution is unachievable, a Tier-2 solution is considered. Tier-2 solutions rapidly call-up or utilize field-ready capabilities or “deploy” new or additional capabilities that are field-ready. The targeted timeframe for delivering usable Tier-2 solutions is days-to-weeks from the time at which the need is established. The focus of activities in Tier-2 solutions is on achieving responsive exploitation, augmentation, or reconstitution of space force enhancement or space control capabilities through rapid assembly, integration, testing, and deployment of a small, low cost satellite. Finally, when an expressed need is not be addressable through existing capabilities (Tier-1) or through the rapid deployment of field-ready capabilities (Tier-2), ORS efforts must focus on the rapid development and deployment of a new Tier-3 capability. Once developed, Tier-3 capabilities will be responsively deployed and employed in the same way as Tier-2 assets. The goal for execution 7of Tier-3 approaches is months-to-one year from established need to presentation of operational capability.2 Within the three-tiered suite of ORS responses, the Tier-2 response requires the largest paradigm shift, i.e., departure from the current space enterprise best practices and concept of operations. How do you respond to any urgent need within a week with a 100% reliable, perfectly effective solution with minimal non-recurring cost? The simple answer is, “You don’t.” The Tier-2 problem is solved as much by defining boundaries, as it is by enforcing interface standards or developing innovative technologies and techniques, more akin to an aircraft depot than a one-of-a-kind spacecraft development. This paper will start by addressing the boundaries that transform the insurmountable task referenced above into a potentially solvable problem. It will explain how these boundaries not only manage expectations, but serve to focus the ORS Office Tier-2 activities. Next, the paper will present the ORS Office’s envisioned Tier-2 End State and detail the steps the Office is taking over the next six years to achieve our vision: 1) It will explain how the ORS Office is championing interface standards as enablers to the Tier-2 vision and how it will make these standards meaningful by winning industry buy-in through ORS-funded missions that clearly demonstrate the business case; 2) It will explain how the ORS Tier-2 CONOPS is incorporating best-practices from non-space industries, which currently solve problems very similar to the Tier-2 challenge; 3) It will present analysis to show how the solution was derived from a detailed trade of schedule, cost, flexibility of the solution to meet the portfolio of anticipated urgent needs, and reliability and effectiveness of the ultimate solution; and 4) Finally, it will summarize the status and results of ongoing ORS Office Tier-2 efforts. | |
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