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Past Conference Papers:
Technology
| Paper Number RS1-2003-2001: A Docking Solution for On-Orbit Satellite Servicing: Part of the Responsive Space Equation | |
| Pete Tchoryk, Jr. (Michigan Aerospace Corporation), Anthony B. Hays (Michigan Aerospace Corporation), Jane C. Pavlich (Michigan Aerospace Corporation) | |
| View/Download:Paper | |
| Abstract: The ability to service satellites on orbit (i.e., refuel or replenish consumables, reboost, repair, or upgrade) would increase both their lifetime and utility. In the case of tactical space assets, there is an ever-growing need to maneuver satellites to new orbits on a regular basis. The ability of a satellite to change orbits allows it to be repositioned to observe new crisis areas, reduces the need for multiple-satellite constellations to provide global coverage, and also makes it less vulnerable to attack. Michigan Aerospace Corporation has developed a mechanism that enables autonomous docking and servicing of space assets. Based on the probe/cone concept, the docking mechanism provides a softdocking capability and can be used with any size asset, from nanosats to full-size spacecraft. A prototype of the mechanism has been successfully tested at Marshall Space Flight Center’s flat floor facility and on NASA’s KC-135 microgravity aircraft. The simplicity of the concept has resulted in a low-cost design that minimizes size, weight, and power, while maximizing tolerance to misalignment, which means that it can be integrated with a minimal effect on the spacecraft. This paper will discuss the benefits of the docking mechanism as an enabling technology for military, scientific and commercial autonomous spacecraft applications and as one of the elements in a responsive space architecture. | |
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| Paper Number RS1-2003-2003: Developments in Commercial Near-Space Systems | |
| Jerry Knoblach (Space Data Corporation), Jerry Queenville (Space Data Corporation) | |
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| Abstract: An Arizona company is nearing commercial deployment of a wire less communications system based on platforms which free-drift on balloons at 31,500 meters (100,000 feet). The system provides two-way wireless messaging to existing commercial devices in a 570-km (360-mile) diameter coverage circle. The system is based on a constellation of small (less than six pounds), low cost (a few hundred dollars each), short life (24 hours) expendable electronics packages launched on conventional weather balloons, which free drift with the uniform winds in the stratosphere. A seventy-platform constellation covering the continental United States (US) is planned to enter commercial operations within a year. This new type of constellation can service several traditional space missions on a lowcost and responsive basis because the major barriers related to rocket launches are eliminated. Also since the platform is over twenty times closer to the earth than even a low earth orbit satellite, radio link budgets and optical resolutions are greatly improved. These stratospheric platforms offer several advantages in the areas of development time, cost, logistics, and responsiveness. Since the platforms are over 20 times closer to the user than a satellite, only move at tens of miles per hour, and are not subject to long-term exposure to radiation in the space environment, they can be developed using the components and processes widely used in the commercial electronics industry. Since the platforms make use of mass production techniques and mass-produced components, the cost per platform can be very low cost. Since the platforms are launched on weather balloons which have been in operation on a daily basis from thousand of sites for over half a century and are produced using readily available contract manufacturing resources, the logistics are well established and current launch sites provide coverage to virtually the entire landmass of the earth. The platforms can be prepared for launch and put on station in only two hours from a site a couple hundred miles from the desired area of coverage. With Moore’s Law continuing to increase the capability and reduce the costs of electronics, the velocity of technology is everyday increasing the advantage of systems that can be developed and deployed quickly. This paper details the development of the current wireless communications network currently in precommercial testing and then discusses the range of possible applications for the technology including: remote imaging, wireless voice and high data rate applications, signal intelligence, bi-static radar and others. It also offers several comparisons of traditional satellites versus this new technology on important capability metrics. | |
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| Paper Number RS1-2003-2004: Operational Concepts and Payoffs for Responsive Space Systems | |
| John M. Borky (Tamarac Technologies), Robert E. Conger (Microcosm) | |
| View/Download:Presentation | Paper | |
| Abstract: The ability to rapidly emplace space assets with tailored payloads and orbits or to deliver a variety of terrestrial systems globally and at orbital speeds, providing specific operational support to a variety of military missions, would lead to revolutionary operational concepts. Three complementary elements are needed for this operationally responsive military space capability (referred to in the paper as “Responsive Space”): truly responsive and affordable launchers, highly modular and standardized satellites, and tactical reentry systems. We first consider a set of operational concepts that explore the potential roles of such systems in various military scenarios. We make preliminary estimates of size, weight, performance and cost. We then examine the technologies that make these concepts feasible in the near- to mid-term, say 5 to 10 years from the start of properly funded development programs. We show that much of the enabling technology portfolio is in hand or well along toward demonstration. The inescapable conclusion is that a military capability of genuinely revolutionary impact is not only available but essential to the realization of the kind of information-enabled operations that are at the core of Joint Vision 2020 and, indeed of military transformation in general.* | |
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| Paper Number RS1-2003-3002: Advanced Lightweight Electronically Steered Antennas for Responsive Space Payloads | |
| Wyman L. Williams (EMS Technologies) | |
| View/Download:Presentation | Paper | |
| Abstract: Due to the mix of missions that Responsive Space payloads will be called upon to execute, and the available launcher technology, it appears likely that initial implementations will consist of constellations of small LEO satellites. It is likely that most Responsive Space payloads will need to maintain multiple simultaneous data links to the battlefield, even if the spacecraft’s primary mission is something other than providing tactical communications. In order to maintain broadband communications with multiple fixed or mobile terminals distributed across a theater of interest, these LEO satellites will require agile high gain communications antennas. Providing such an antenna within the size, mass, and cost constraints of a Responsive Space payload presents a number of technical challenges. This paper presents some of the design trades involved in selecting an antenna technology for this application. It also gives examples of some recently developed antenna and beam forming technologies that show promise for addressing Responsive Space payload requirements. The devices shown have adequate performance for a wide variety of missions, were developed for space or military airborne applications, and can be qualified for the Responsive Space mission with low risk. | |
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| 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 RS1-2003-4005: Affordable, Responsive Space Asset Delivery by Electro Dynamic Delivery Express (EDDE) | |
| Joseph Carroll (Thether Applications) | |
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| Paper Number RS1-2003-5000: Technology for Responsive Space Capability | |
| Robert Pugh (AFRL) | |
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| Paper Number RS1-2003-6004: Pistonless Dual Chamber Rocket Fuel Pump | |
| Steve Harrington (Flometrics) | |
| View/Download:Presentation | Paper | |
| Abstract: A positive displacement pistonless rocket fuel pump uses two pumping chambers alternately filled and pressurized in sequence to maintain a steady flow of pressurized propellant to a rocket engine. This pump fills the gap between pressure fed and turbopump rockets by making a lower cost rocket feasible without the weight of a pressure fed design or the high cost and complexity of a turbopump. Thrust to weight ratios are calculated for the pump using typical fuel combinations. For a 2219 aluminum LOX/RP-1 pump at 4 MPa the thrust/weight ratio of the pump is ~700. Design and test data for a prototype which pumps water at 3.5 MPa and 1.2 kg/s is presented. The low cost and simple construction of the pump allow for low cost, mass producible propulsion systems to fulfill a responsive space requirement. | |
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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 RS2-2004-8000: Almost There: Responsive Space | |
| Grant Williams (SpaceDev), Frank Macklin (SpaceDev), Marti Sarigul-Klijn (SpaceDev), Nesrin Sarigul-Klijn (SpaceDev), Jim Benson (SpaceDev) | |
| View/Download:Paper | |
| Abstract: SpaceDev is developing a hybrid-based small launch vehicle called StreakerTM. In addition to providing a low cost (<$5M total) alternative for launching microsats, Streaker is being designed to fulfill multiple responsive space requirements. By their very nature, the building blocks of the Streaker hybrid propulsion system, Nitrous Oxide and HTPB rubber, are inert, safe to handle, and readily storable. Range safety complexity is reduced by reduced toxic material handling requirements and by the Streaker thrust termination capability. Building on extensive hybrid motor testing conducted by AMROC and SpaceDev, the Streaker vehicle is transitioning into a modular launch system family, with configurations capable of either air or ground launch. As part of this development, the infrastructure is being developed to accommodate a mobile ground launch capability, and provision for rapid (hours) launch turnaround. Through a number of on-going programs at SpaceDev, the necessary components of the Streaker, i.e. the Common Core Booster (CCB), hybrid upper stage, and hybrid transfer stage (MoTV) are maturing, targeting a first launch in 2007-08. | |
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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 RS2-2004-7000: Tether Balute Low Cost De-Orbit and Recovery Project | |
| Radu Dan Rugescu (University of Bucharest) | |
| View/Download:Paper | |
| Abstract: Balute recovery systems for sounding vehicles are well known. Such systems were proposed also for the end recovery in the US Gemini program. Tether aerodynamic planetary capture was also investigated in Seattle. Now the project TEBLOR is offered to make profit of these potentials in a new de-orbit and reentry system for recovery of small-sized payloads in a cost-effective manner. In the computational model perturbations were considered as produced by aerodynamics, the Moon and the Sun gravity fields. The system is able to achieve the initial de-orbit maneuver without reaction motors. The tethered air-brake is electrically actuated and exploits the dynamics of orbital two-body tethered systems. Potentiality to use electro-dynamic effects is suggested. Regarding the technological concept, it makes use of carbon nano-tubes in a similar, but a less risky manner than the space lift system. After the non-powered de-orbit, the high altitude, tethered air brakes are further used. Here a major impact upon the descent trajectory and landing site position is introduced by the system dynamics that causes the main concern. The last phase of low altitude, lower speed descent is running in a routine manner where the tethered balute transforms into a usual recovering device. The system should be transformed into a re-usable application for small size, low mass Earth satellites. The study consists in the appraisal and evaluation of the TEBLOR project. | |
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| Paper Number RS2-2004-3005: Using IDOS To Develop Eagle into "Real" Flight Software to Support Responsive Missions | |
| D.T. Chen (Universal Space Lines) | |
| View/Download:Presentation | Paper | |
| Abstract: The Integrated Development and Operations System (IDOS) is intended to be a rapid development and operations platform for new launch and atmospheric vehicles. A “real life experience” of its use to develop an advanced guidance algorithm is described here. The Evolved Acceleration Guidance Logic for Entry (EAGLE) is an advancement to existing entry guidance algorithms (such as that flown on the Space Shuttle) and has the potential to reduce the amount of pre-mission design effort, increase the range of entry opportunities and contribute to achieving “aircraft like” operations (i.e., responsive missions). EAGLE is based on the concept of planning and tracking aerodynamic acceleration profiles, a concept developed and proven effective in the Apollo and Shuttle programs. The most distinguishing feature of EAGLE relative to the Shuttle entry guidance is its ability to plan a three-dimensional trajectory, taking into account the longitudinal and lateral dimensions, and thereby handling entries as well as aborts with significant cross range motion. Development of EAGLE was started at the University of California, Irvine (UCI) under the supervision of Dr. Kenneth Mease. Part of its development was coordinated under the Marshall Space Flight Center’s (MSFC) Advanced Guidance and Control (AG&C) Technologies effort such that competing designs could be compared and analyzed by MSFC. EAGLE consistently obtained high scores in the various test cases defined for the AG&C effort. During development of IDOS, EAGLE was transitioned into the IDOS development environment and has continued development and testing. Development, implementation and test within IDOS required less than one “equivalent person” labor effort and was our “pathfinder” implementation to drive the IDOS development down a path that best supports the advanced algorithm development goal. Details of this development in IDOS will be presented, as well as results of its performance (flight quality objectives as well as performance on real time flight like processor hardware). Comparison to development time using traditional software development tools will be provided. This “case study” is just one example of how IDOS is an enabling technology for responsive missions. | |
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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 RS2-2004-4003: A Conformally-Mounted SatCom Antenna System To Support Stars Phase-2 Testing | |
| W.P. Cooke (EMS Technologies), M.G. Guler (EMS Technologies), L.A. Cintron (EMS Technologies), J.P. Montgomery (EMS Technologies), D.E. Whiteman (Dryden Space Flight Research Center), R.D. Sakahara (Dryden Space Flight Research Center), R.J. Franz (Dryden Space Flight Research Center) | |
| View/Download:Presentation | Paper | |
| Abstract: Maintenance and staffing of down-range tracking and telemetry stations is a large component of the cost of current space launches. While the data gathered by such telemetry stations are vital to the technical assessment of the launcher’s performance and safety, more cost effective methods of gathering telemetry data must be developed in support of Responsive Space missions. This paper reports on the development of a telemetry subsystem that uplinks data from a launch vehicle directly to the Tracking and Data Relay Satellite System (TDRSS), reducing the need for ground-based telemetry | |
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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 RS2-2004-6005: Challenges, Enabling Technologies and Technology Maturity for Responsive Space | |
| Kevin G. Bowcutt (Boeing), S. Jason Hatakeyama (Boeing) | |
| View/Download:Presentation | Paper | |
| Abstract: The goals of responsive space may be best met with pure rocket technology, a combination of rocket and hypersonic air-breathing technology, or all-rockets followed by application of air-breathing propulsion at a later date. The optimal approach is difficult to determine today due to uncertainty and immaturity of the key technologies required by either approach. An added complication is the large number of technology options that can be combined in numerous ways to achieve responsive space systems. Propulsion is the primary technological challenge for responsive launch systems -- either low-maintenance, long-life, high thrust-to-weight ratio rocket engines or air-breathing turbine engines combined with scramjets. Additional technologies contributing both to vehicle performance and operational utility must also be matured before responsive space can be realized. In this paper, primary challenges and enabling technologies will be outlined for both rocket and air-breathing reusable launch vehicles, many of which are common to both, and a maturity assessment will be made for each. Technologies that improve vehicle performance, operational utility and other contributors to space launch responsiveness and low life cycle cost will be included in the assessment. Also outlined will be a recommended approach for determining which propulsion system or combination of propulsion systems will best meet the ultimate requirements of responsive space access based upon actual flight data. | |
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| Paper Number RS2-2004-7001: Highly Operable Propulsion Technologies and Propulsion System Approaches for Operationally Responsive Space Systems | |
| Claude Russell Joyner II (Pratt & Whitney), Patrick McGinnis (Pratt & Whitney), Richar Hagger (Pratt & Whitney) | |
| View/Download:Presentation | Paper | |
| Abstract: In 2001-2002, the USAF developed an Operationally Responsive Spacelift Mission Needs Statement (ORS MNS) that defined the requirement for responsive, on-demand access to, through, and from space. This requirement not only encompassed the spacelift mission of delivering payloads rapidly to or from orbit, and their operation on orbit; it also states the systems will address operational issues related to launch on-demand, provide mission flexibility, and be cost effective. The ORS MNS indicates that future operational systems are needed that must be able to launch within hours of need, be low cost relative to today’s launch architectures, and provide greater onorbit mission flexibility. Previous Responsive Space workshops have also indicated that operationally responsive space systems must also provide a roadmap toward overcoming the business barriers that inhibit low cost operation. Propulsion technologies and the propulsion system have been shown to be one of the primary enablers of meeting the requirements that formulate Operationally Responsive Spacelift architectures. Pratt & Whitney has explored several promising individual and integrated propulsion technologies that exist today or are in development that would provide the USAF with the capability to launch with-in hours of a mission determined need. These propulsion technologies include use of modular or integrated air-breathing booster systems, employment of soft-cryogenic fuels and oxidizers, and hybrid motors. Part of the employment of the soft-cryogenic fuels (i.e. liquid methane) could include the use of an Integrated Thermal Management Unit (ITMU) that would provide a stable storage environment during pre-launch and during launch. This unit would also have synergy with the use of methane for on-orbit systems. The air-breathing propulsion would have synergy with current USAF systems and act to provide a capability to launch from CONUS bases to deploy assets on demand to any inclination. These booster systems would be derived from current hardware, thus reducing the cost to develop and operate them as part of a USAF evolutionary responsive launch system. The expendable, low cost hybrid boosters could be used as alternative booster or strap-on stages, as launch assist systems, much like the “JATO/RATO boosters” of the 1950’s or low cost expendable upper stages for in-space systems. The use of and advanced integrated engine health management system (EHMS) working with the vehicle management system (IVHM) creates a “system of systems”. It creates a more responsive system by increasing the overall reliability by integrating propulsion or engines systems health management with the overall vehicle health management system. Consequently, this paper will discuss the attributes and applicability of specific propulsion technologies and systems that could provide a robust performance capability as well as the capability meet a military aircraft-type operational responsiveness via horizontal take-off boosters. | |
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| Paper Number RS2-2004-7003: A Status Report on the Development of a Nanosat Launch Vehicle and Associated Launch Vehicle Technologies | |
| John M. Garvey (Garvey Space Corporation), Eric Besnard (CSULB) | |
| View/Download:Presentation | Paper | |
| Abstract: Much current attention on responsive launch capabilities is focusing on small launch vehicles that can deliver on the order of several hundred kilograms to low Earth orbit. However, for developers and operators of even smaller spacecraft with masses on the order of 10 or less kilograms, i.e. “nanosats” and “picosats,” this class of launch systems is still oversized and the costs are at least a magnitude too great. Consequently, an effort is presently under way to address this evolving launch service niche. Through the California Launch Vehicle Education Initiative (CALVEIN), a joint academic-industry team is developing and flight testing a series of prototype vehicles that are demonstrating and evaluating candidate technologies and operations for a notional nanosat launch vehicle – an “NLV” – that could deliver 10 kg to low polar orbit. To date, this program, which is hosted by the California State University, Long Beach, in partnership with Garvey Spacecraft Corporation, has developed four reusable LOX/ethanol launch vehicles and participated in seven flight tests as well as a similar number of static fire tests at the Mojave Test Area. Principal accomplishments include the first-ever powered flight test of a liquid propellant aerospike engine. In parallel, several classes of aerospace engineering students have gained invaluable experience working with actual flight hardware. Photo by Tony Richards This paper reports on the results of several recent NLVrelated research and development activities. These include the pioneering aerospike flight tests, ongoing development of low-cost thrust vector control subsystems and an initial static fire test using propylene as an alternative hydrocarbon fuel. | |
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| Paper Number RS2-2004-7004: Launch Vehicle and Spacecraft System Design Using the Pistonless Pump | |
| Steve Harrington (San Diego State University) | |
| View/Download:Paper | |
| Abstract: The application of a pistonless pump to a launch vehicle or spacecraft can provide cost and reliability improvements over standard pressure fed or turbopump fed designs. Calculations show that in a first stage launch vehicle application, a system which uses the pistonless pump has comparable performance to gas generator turbopump designs. The performance can be improved by using lowpressure liquid helium which is pumped using a pistonless pump to high pressure and then heated at the engine. This allows for lower pressurant tankage weight. This system uses less than 1% of the fuel mass in liquid helium, which offers a performance advantage over comparable gas generator turbopump powered rockets. A complete overall vehicle design is presented which shows how the various systems are integrated and how much each component weighs. The vehicle uses LOX/hydrocarbon propellants at moderate to high pressures to achieve high performance at low weight and low cost. The pump is also shown to have significant performance and flexibility increases for spacecraft when combined with high-pressure storable propellant engines. The proposed pump is also applicable to pumping gelled propellants. | |
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| Paper Number RS2-2004-7005: The Lightband As Enabling Technology For Responsive Space | |
| Walter Holemans (Planetary Systems Corporation) | |
| View/Download:Presentation | Paper | |
| Abstract: The Lightband separation system used to separate space vehicles is presented as means to enable responsive space. It does this by simplifying payload integration, lowering cost, delivering rapidly and by the utility of its advanced features. The Lightband can reduce integration time from days to minutes. Low cost is achieved by PSC’s focus on the Lightband and all of its associated production tasks including: detailed design, procurement, testing and sales. A by-product of this focus is the upcoming capacity to stock and deliver Lightbands within a day. The advanced features, which resulted from five years and $1.5M yielded a litany of improvements that allows users to save weight, increase payload height, complete inexpensive full-system testing, reduce shock, buy an integrated solution, eliminate temperature dependence and eliminate hazards associated with pyrotechnics and fracture. | |
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| Paper Number RS2-2004-A021: High Power All Electric Power-Actuation Technology for Responsive Space Missions | |
| Eric Knight (Lockheed Martin) | |
| View/Download:Presentation | |
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| Paper Number RS2-2004-A023: Launch Vehicle and Spacecraft System Design Using Pistonless Pump | |
| Steve Harrington (Flometrics) | |
| View/Download:Presentation | |
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| Paper Number RS3-2005-1003: On-Demand Science Missions | |
| John J. Webb, Jr. (Instarsat) | |
| View/Download:Presentation | Paper | |
| Abstract: Over the last four decades, robotic space explorers have yielded a wealth of scientific discoveries about our solar system and its origins. However, the resources required to design, develop, launch, and operate such missions is enormous. The highly prohibitive nature of established design practices and long development cycles significantly precludes responsive science investigations. Historically, robotic science missions flown in the last forty years have been highly limited in scope and capability. This paper briefly reviews the current practices in use for developing science missions, including; mission design, spacecraft design, and cost estimating. In contrast, today’s science missions must be more responsive to changing circumstances. The advances in space related technologies make ondemand science missions even more relevant and desirable. The spacecraft capabilities, capacity, and cost effectiveness are essential deterministic factors enabling successful on-demand science missions. This paper will focus on defining these factors within the context of a responsive space system. This paper discusses the emergence of new space-related technologies that will accelerate the development of on-demand science missions. This discussion includes an overview of current advances in materials, communications, propulsion, and onboard autonomous systems that can play a critical role in the successful design, development, and operation of on-demand science missions. Finally, this paper discusses an on-demand science mission life cycle scenario. | |
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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 RS3-2005-2005: Operational Responsive, Persistent Space Systems | |
| Terrance A. Weisshaar (DARPA), Ivan Bekey (Bekey Designs) | |
| View/Download:Paper | |
| Abstract: A new paradigm for operationally responsive space system development is discussed, together with four examples of these types of systems. The current approach to responsive space systems and future systems with multiple roles operating in hostile environments is unlikely to meet future needs. In particular, small satellites with specialized missions and larger monolithic satellites are important, but very limited, means of achieving affordable responsiveness to unforeseen threats or desired flexibility. The new paradigm of reconfigurable/morphing systems includes modular techniques that exploit new technologies and can be developed quickly, taking advantage of spiral development processes, and also adapt semiautonomously to new requirements, saving serial development time and cost. Among the features that drive reconfigurable systems are greater degrees of self-awareness of the system and surrounding environment, together with autonomous features not seen today. Equally important is the ability to operate a multi-module system as a single connected unit or as a set of separated units that can exchange information, power and propellants when required to perform alternate missions. These systems will appear in the next decade if the special technologies required for their development and operation are matured beyond current embryonic stages. Some principal technologies required are identified and discussed. | |
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| Paper Number RS3-2005-3001: Cubesats As Responsive Satellites | |
| Armen Toorian (Cal Poly, San Luis Obispo), Emily Blundell (Cal Poly, San Luis Obispo), Jordi Puig Suari (Cal Poly San Luis Obispo), Robert Twiggs (Stanford) | |
| View/Download:Presentation | Paper | |
| Abstract: California Polytechnic State University, in coordination with Stanford University, has developed the CubeSat standard to provide inexpensive and timely access to space for small payloads. These picosatellites, built mostly by universities, are 10 centimeter cubes with a mass of 1 kilogram. Of the 40 or so participating universities and private firms, more than 60% of CubeSat developers reside in the United States. Our goal is to make launching these satellites easy and cost effective by coordinating launches and providing a reliable deployment system. This paper will discuss Cal Poly’s role in the CubeSat program, and the characteristics of the project which create practical, reliable, and costeffective launch opportunities. | |
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| Paper Number RS3-2005-3002: KUTESAT-2, A Student Nanosatellite Mission for Testing Rapid-Response Small Satellite Technologies in Low Earth Orbit | |
| Trevor Sorenson (University of Kansas), Glenn Prescott (University of Kansas), Marco Villa (University of Kansas), Dewayne Brown (National Nuclear Security Administration), John Hicks (National Nuclear Security Administration), Arthur Edwards (AFRL), James Lyke (AFRL), Thomas George (JPL), Sohrab Mobasser (JPL), JPL (JPL), Scott Tyson (Space Microsystems) | |
| View/Download:Presentation | Paper | |
| Abstract: The Air Force Research Laboratory (AFRL) is interested in using nanosats to perform space experiments, demonstrate new technology, develop operational systems, and integrate advanced responsive space system technology. One potential operational application of nanosats is using clusters of microsatellites that operate cooperatively to perform the function of a larger, single satellite. Each smaller satellite communicates with the others and shares the processing, communications, and payload or mission functions. This type of a distributed system has several advantages: (1) systemlevel robustness and graceful degradation, and (2) distributed capabilities for surveillance and science measurements built into the system architecture. There are a number of technology advancements needed to operationalize and enable tactical missions. These advancements include modular ‘plug-n-play’ satellite architectures and components; high performance tactical downlinks; adaptable, agile propulsion systems, and lean manufacturing, assembly and test. The Kansas Universities’ Technology Evaluation Satellite (KUTESat) program originated at the University of Kansas (KU) in 2002. The technical objective of the program is the development and operation of miniature satellites that can demonstrate and test technologies and techniques necessary to accomplish various government missions. The first satellite, KUTESat-1 Pathfinder, was designed to perform imaging and measure radiation from orbit. The design and construction of this 1-kg satellite helped KU to develop the capability to produce and operate small research satellites. Pathfinder is due for launch in mid-2005. Nanosats are a rapid and low-cost technology platform for the space testing of a broad range of micro-electro-mechanical systems (MEMS) and nanotechnologies as well as new mission architectures. The KUTESat program offers a low-cost solution to the problem of acquiring “space heritage” for new technologies and concepts. These programs can undertake higher risk missions that would be otherwise avoided by more conservative mission planners. Thus new MEMS and nanotechnologies related to avionics, guidance and control, communications, imaging, maneuvering, and instrumentation are offered a rapid and low-cost approach to space testing that will help realize a rapid response space force. The objective of the current program is to develop and fly a nanosatellite to test components, technologies, and concepts that are of use to the AFRL, the National Nuclear Security Administration (NNSA) and the National Aeronautics and Space Administration (NASA), while providing a valuable contribution to the education of students who will soon be entering the space workforce. KU is leading a team consisting of the NNSA Kansas City Plant, the AFRL, and NASA Jet Propulsion Laboratory (JPL) to design and execute the KUTESat-2 mission using a 16-kg nanosatellite based on the Pathfinder satellite with much commonality in the avionics and ground system. The major technologies to be tested include: a miniature distributed and adaptive S-band transceiver; a miniature maneuvering control system; standardized interface (“plug and play”) electronic modules; various MEMS technologies, including a single-axis MEMS gyroscope; a micro sun sensor; an array of miniature dosimeters; and a miniature imager. New capabilities to be tested include a Tracking and Data Relay Satellite (TDRS) communication demonstration with the Sband transceiver, and demonstration of target inspection capability using a deployed inflated target. The KUTESat-2 will be prepared for a launch in 2007. | |
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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 RS3-2005-5001: 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 RS3-2005-5002: PNP Transceiver-FNT | |
| K.D. Brown (NNSA KCP) | |
| View/Download:Presentation | Paper | |
| Abstract: The National Nuclear Security Administration’s (NNSA) Kansas City Plant (KCP) has supported design, development, and production of weapon flight test telemetry technology in conjunction with the national weapons laboratories for decades. The notion of a responsive satellite is predicated on bus components that are flexible, pre-qualified, and interoperable with other bus subsystems across a quickconnect interface bus. This paper describes a novel, standardsbased, modular, scalable, dynamically configurable, wireless network transceiver architecture that can support integration into a satellite interface bus to provide flexible network data links for a range of network centric requirements A plug and play interface is under development and will be integrated into the Flexible Network Transceiver (FNT). The paper will describe the transceiver’s development, architecture, and capabilities which support responsiveness. | |
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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 RS3-2005-5004: Reconnaissance Payloads for Responsive Missions | |
| Charles Cox (Goodrich Optical and Space Systems Division), Stanley Kishner (Goodrich Optical and Space Systems Division), Richard Whittlesey (Goodrich Optical and Space Systems Division), Fredrick Gilligan (Goodrich Optical and Space Systems Division) | |
| View/Download:Presentation | Paper | |
| Abstract: Operationally responsive Electro-Optical (EO) imaging capability exists and is routinely used to provide intelligence information to the tactical war fighter. This capability is provided by Goodrich Reconnaissance systems having “plug and play” interfaces to strategic (i.e., U-2) and tactical airborne platforms. These operational systems have visible, IR and multispectral capability, and the resulting data readily interface into an existing infrastructure providing timely information to theater commanders. These airborne operational systems can be modified to provide reconnaissance capabilities from space to support the Operationally Responsive Space (ORS) vision. This paper describes these systems, summarizes some of the utility provided by them, and discusses hardware modifications and operational scenarios consistent with lowcost mission requirements. This approach, the modification of existing airborne operationally responsive EO imaging systems, provides a low cost alternative to top-down special-purpose development and leverages a continually evolving product stream to provide ORS payloads. | |
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| Paper Number RS3-2005-5006: How Not to Design an Avionics System | |
| Jason E. Holt (Brigham Young University) | |
| View/Download:Presentation | Paper | |
| Abstract: In 1995, four Utah Universities launched a hybrid sounding rocket at the Utah Test and Training Range. In December 2003, they successfully launched a much larger successor to that rocket. We describe the design, construction, deconstruction, redesign and reconstruction of the avionics package during the 8 year period between flights, then describe the system which was actually flown. That package used COTS hardware worth less than $1000, was substantially redesigned within weeks of the launch, and was completely destroyed after an entirely successful flight upon an otherwise soft, vertical landing. Although the package met only simple requirements and used no cutting-edge hardware, we feel that the lessons we learned from both technical and social standpoints will be useful to others who wish to rapidly develop avionics systems despite severely limited resources. Furthermore, we describe a new, straightforward design for the core control system which is a result of the lessons we learned, and which we hope will be flexible enough to meet the continuing demands of our project and potentially many other projects as well. | |
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| Paper Number RS3-2005-6001: Aquarius Low Cost Launch Main Engine Study | |
| Andrew E. Turner (Loral) | |
| View/Download:Presentation | Paper | |
| Abstract: Congressional funding has been provided to study the main engine for the new low-cost launch vehicle Aquarius. The goals and background to this study are discussed in this paper, and an overview of the planned approach to developing and testing a concept for a new lower-cost engine is presented. The new engine concept is the Vortex Cooled Chamber Wall engine by ORBITEC, which has been evaluated only at small scales up to this point. The Aquarius vehicle, which incorporates this new engine and other innovations promises to reduce launch cost by an order of magnitude. | |
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| Paper Number RS3-2005-6002: Operationally Responsive Space Is Here Now Using Minotaur Class Vehicles | |
| Randall Riddle (Kirtland Air Force Base), Robert Kelsey (Aero Thermo Technologies), Mitch W. Elson (Rocket Systems Launch Program), Steven J. Buckley (Kirtland Air Force Base), Scott Schoneman (OSC) | |
| View/Download:Presentation | Paper | |
| Abstract: Responsive space takes on different meanings depending on who you talk to, but the common denominator is that it must be done quicker. Not only must the missions be completed faster, but they should also cost less than current systems. There are high hopes that technological breakthroughs in launch vehicles are going take costs to the hundreds of dollars per pound/kilogram for low earth orbit. There are serious technical hurdles to achieving that type of capability, the toughest being a new propulsion system. No current propulsion theory is going to revolutionize access to space, nor are there any in the near future with that promise before 2020. There are technology improvements in areas like turbo pumps, vapor pressure systems, solid rocket fuels, liquid fuels, guidance software, flight computers, materials and hundreds of other components, and while those offer cost savings and greater performance, they are not the revolutionary concept that is going to change the paradigm. The research and development efforts should continue, and in the time frame between now and 2020, the community should look to ways of using assets that are currently available as a bridge to future programs that are cost effective and require no development, which allows developmental funds to be applied to what ever concept is going to be our access to space in 2020. In April, 2003, the Rocket Systems Launch Program (RSLP) attended the 1st Annual Responsive Space Conference held in Redondo Beach California. Since then, RSLP has looked into what is possible with the space launch vehicles on contract today. Current space launch vehicles on contract with RSLP are derived from the US Air Force’s decommissioned ICBMs that were developed to be launched on demand in defense of the country. Motor configurations are all solid propellant derivatives providing a truly call-up launch capability. The paper addresses options to meet reduced mission timelines, providing options ranging from a nominal 18 months to within 48 hours. The paper discusses how as timelines are decreased, overall mission risk remains low. Typically as the mission timeline decreases, so does the flexibility of the payload to make changes. The paper will explore and propose more efficient processes inherent to the launch vehicle that would help alleviate the loss of flexibility. Timelines for various responsive scenarios will be presented as well as the associated costs broken out between non-recurring, developmental, recurring and yearly funds required to maintain readiness when applicable, and launch costs. | |
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| Paper Number RS3-2005-6003: The Evolution of Space Launch Booster Designs In the 21st Century | |
| Matt Steele (OSC), Warren Frick (OSC) | |
| View/Download:Presentation | Paper | |
| Abstract: The future technical approaches for space launch vehicles in the 2010-2050 timeframe will mirror the development and evolution of the US and Soviet ICBM fleet during the cold war. The result of objective trade studies yields simple, solid propellant launch systems that cost less, are more responsive, and more reliable than today’s liquid propellant space launch fleet. As future space payloads mature and become more standardized and miniaturized, the pressure for lower cost launch platforms will increase. The result will be a move away from expensive, high performance liquid fueled systems and towards systems that can deliver payloads to space inexpensively and reliably. For launch systems where responsiveness is crucial, the use of a solid propellant system will be seen as the ultimate solution to the problem, just as today’s nuclear-tipped ICBMs are the evolution of a similar problem faced by the US Air Force. For the launches where low cost and flexibility in orbital inclination is required, a hybrid system, with a re-usable first stage and an expendable solid propellant system for the upper stages will be the logical choice. Once the launch rate can justify the fixed investment of the reusable stage, the marginal cost per flight can significantly lower overall launch costs. Incremental technological advances in materials, processes, and launch procedures will combine to lower costs and reduce risk of these systems. Companies that can exploit these technologies in a cost effective manner will maintain a competitive advantage over those tied to older, more complex and failure-prone systems or those exploring high risk, high payoff breakthrough technologies. | |
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| Paper Number RS3-2005-6005: Operationally Responsive Space: The Vision Launch Architecture Is Dependent On The Requirements | |
| Slater B. Voorhees (Lockheed Martin), N. Patrick Hansen (Lockheed Martin), Heather J. Swenson (Lockheed Martin) | |
| View/Download:Presentation | Paper | |
| Abstract: There have been numerous studies in recent years addressing the government’s desire for a rapid response launch system. An upfront focus on detailed requirements is imperative to determine an optimal launch architecture but often times an architecture type, being expendable, reusable or hybrid (reusable first stage and expendable upper stage), is chosen before a complete set of requirements has been defined. Prematurely choosing the architecture type could result in a solution that is less than optimal from an economic and design standpoint. A requirement that is prominent in today’s aerospace environment is the overall architecture life cycle cost (LCC) which is believed to be driven by the individual launch cost of the system. This paper presents a LCC tool that can be utilized to help determine the optimal architecture solution. Using representative design concepts of a fully reusable, fully expendable, and hybrid launch architectures, we have modeled the effects on LCC when varying requirements such as the life of the program, theoretical first unit cost, and nonrecurring cost. Specific examples will be presented that illustrate that the appropriate choice of the requirements, input parameters, and figures of merit (FOM) on the optimal architecture is imperative to performing a launch architecture analysis. Slight variations in the input parameters or improper choice of a FOM can result in an architecture being identified as optimal incorrectly. | |
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| Paper Number RS3-2005-6006: A Rocket-Powered Technology Demonstrator for Responsive Access to Space | |
| Daniel P. Raymer (Conceptual Research Corporation), Jess Sponable (AFRL), Timothy Fry (University of Dayton Research Institute), Jeremy Zanzig (Analytical Methods), Jared R. Ethier (Composite Engineering), Mitchell Burnside Clapp (Pionneer Rocketplane) | |
| View/Download:Presentation | Paper | |
| Abstract: A six-company team headed by Conceptual Research Corporation is developing a design concept for a rocket-powered technology demonstrator under funding from USAFWPAFB, with administrative and technical assistance from the University of Dayton Research Institute. This “Micro-X” demonstrator offers affordable and incremental demonstration of responsive space access system concepts and enabling technologies, and will demonstrate high-tempo reusability in an operational environment. To keep the program affordable, the Micro-X demonstrator will have an empty weight of less than 10,000 lbs yet will be capable of reaching space altitudes and hypersonic speeds. | |
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| Paper Number RS3-2005-A015: Technology for Responsive Space | |
| Jaime Esper (GSFC), Bruce Underwood (GSFC) | |
| View/Download:Presentation | |
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| Paper Number RS3-2005-A016: Technologies Necessary for Joint Warfighting Space | |
| Peter M. Wegner (AFRL) | |
| View/Download:Presentation | |
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| Paper Number RS4-2006-1002: Fractionated Space Architectures: A Vision for Responsive Space | |
| O. Brown (Defense Advanced Research Projects Agency), Paul Eremenko (Booz Allen Hamilton) | |
| View/Download:Presentation | Paper | |
| Abstract: The advent of the integrated circuit some four decades ago set off mankind’s insatiable thirst for computational power. The quest to quench this desire led to the development of increasingly more sophisticated computers. Microchips sprouted ever greater numbers of transistors, choking buses, and forcing memory banks to struggle to keep up. The novelty of micro- and minicomputers was quickly trumped by the sheer computational prowess of supercomputers. And so the trend continued. In a matter of two decades, however, this drive towards greater processing power culminated in mammoth mainframes whose rapidly increasing complexity, fragility, and cost quickly outpaced the capability gains. A scant few years into the second decade of the era of the integrated circuit, the availability of inexpensive, mass-produced microcomputers, and the advent of fast, seamless internetworking ensured the relegation of the large monolithic mainframes to obsolescence and obscurity. Spacecraft have followed a trajectory that is uncannily parallel (and, of course, technologically intertwined) to the history of high-end computing. Borrowing the historical analogy, we posit that the era of distributed space architectures has likewise arrived. The gargantuan monolithic systems deployed to orbit today have grown too large, too complex, too fragile, and consequently much too expensive; furthermore, these trends have not been offset by commensurately rapid growth in capability. We propose a fractionated architecture for space systems, whereby a satellite is decomposed into a heterogeneous set of components which interact wirelessly. In the extremum, the fundamental functionality of most space systems is the reflection of photons back to earth. Thus, assuming that the requisite photon collection, processing, and re-radiation can be accomplished, the spacecraft need be nothing more than a collection of free-floating “pixie dust.” In the realm of the foreseeable technological future, however, there are a handful of schema for severing and distributing the functionality of a monolithic spacecraft. Perhaps the most basic is fractionating the spacecraft along its data channels, resulting in a loose cluster of networked spacecraft modules. Somewhat more challenging is also fractionating the power system and disseminating power wirelessly among the modules. At the technological horizon is also fractionating the propulsion and stationkeeping functionality, also necessitating the wireless transmission of forces and torques. The fractionated architecture is likely to incur an aggregate mass impact versus its monolithic counterpart (although it is noteworthy that at least one massive component may shrink – the flywheels necessary to ensure payload pointing accuracy need only be responsible for stabilizing and pointing the payload module, not the entire spacecraft). The impact on overall system cost is ambiguous since the cost impact due to greater system mass is at least partially offset by learning curve and mass production effects across the multitude of modules. For a constant required level of functionality, however, the fractionated architecture dramatically outperforms its monolithic counterparts in its value proposition. It affords its user/operator greater flexibility in the form of system scalability, reconfigurability, and adaptability (including multi-payload functionality). It dramatically increases robustness and survivability. It allows the isolation of the payload for both improved security and increased pointing accuracy. It lowers possible increases in lifecycle cost and decreases schedule risk by decorrelating failure probabilities of the various component subsystems and multiple payloads. It improves responsiveness by allowing incremental capability deployment, by enabling the utilization of small launch vehicles for the emplacement of massive orbital capabilities, and by shifting the deployment decision chain from the strategic to the tactical level. Perhaps most importantly – and much like the internetworked microcomputer – it commoditizes the space industry and transforms it from an exotic boutique to a customer-driven, cost-competitive enterprise. The technologies needed to make fractionated space systems a reality are well within reach. They potentially include responsive and inexpensive small launch vehicles, highly secure ultra wideband inter-module data links (which may also provide relative navigation capabilities for the spacecraft modules), efficient radio frequency power transmission, passively stable Keplerian cluster orbits, and mass-produced, inexpensive, space-qualified satellite components (many with their legacy in the newly-emergent field of unmanned aerial vehicles). More esoteric technology options include very high frequency power beams, laser power transmission, and remote force and torque transmission through electromagnetic induction. The Defense Advanced Research Projects Agency (DARPA) has been studying the fractionated architecture concept and is poised to commence an initiative entitled F6 – short for Future Fast, Flexible, Fractionated Formation-Flying Spacecraft utilizing Information eXchange, and incidentally a tornado of unimaginable strength on the Fujitsu scale – that will mature the associated technological, architectural, and organizational advancements necessary for an onorbit demonstration of a fractionated spacecraft. A brief discussion of the vision for F6 concludes. | |
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| Paper Number RS4-2006-1006: Using Proven Aircraft Avionics Principles to Support a Responsive Space Infrastructure | |
| Randy Black (Honeywell Space Applications) | |
| View/Download:Presentation | Paper | |
| Abstract: Creating an engineering environment that supports responsive space involves a variety of interrelated disciplines. Included within these disciplines is the ability to quickly re-configure systems through plug-and-play hardware and software. Plug-and-play hardware as a concept has been progressing well throughout the electronics industry. Plug-and-play software has been somewhat more challenging. While some advances have been made through object-oriented architectures and model-based autocoding, software lags behind hardware in this area. Honeywell has experienced significant success for the past ten years fielding plug-and-play software at the application level. Using a combination of time and space partitioning, table-driven operations, and robust off-line development tools, Honeywell’s Integrated Modular Avionics (IMA) has produced significant savings in development cost and schedule. More importantly, modifications to either hardware or software are quickly and easily integrated into the overall system with minimal re-certification required. During the past decade, Honeywell has produced multiple implementations of this advanced avionics technology. One lesson learned is that specific implementation details are not as important as designing to key architectural principles. This paper describes several of those principles that have a proven track record of enabling rapid reconfiguration of system architectures. Architectural principles that support plug-and-play software applications, as well as minimizing the impact of hardware modifications, provide the core of a system design that is integral to an overall responsive space infrastructure. | |
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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 RS4-2006-5003: 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 RS4-2006-5005: On-orbit Calibration and Focus of Responsive Space Remote Sensing Payloads | |
| Thomas Chrien (Raytheon Space and Airborne Systems), Stephen Schiller (Raytheon Space and Airborne Systems) | |
| View/Download:Presentation | Paper | |
| Abstract: Radiometric, spectral, and spatial response performance and calibration knowledge requirements have been traditional cost drivers for remote sensing payloads. Performance has a direct relationship to the usefulness of the information product, including (1) bore-sight calibration in order to achieve geolocation accuracy, (2) optimization of focus to maximize spatial resolution, and (3) absolute spectral and radiometric calibration for effective atmospheric compensation. Meeting strict requirements prior to launch is problematic. Careful (and costly) compensation must be made for gravity effects, and thermal vacuum test conditions can only approximate on-orbit thermal environments. Furthermore, the trauma of launch and subsequent space contamination can invalidate a “perfect” pre-launch calibration. An alternative approach is to fine tune focus and calibration after the payload is on orbit using vicarious calibration techniques. This reduces cost and schedule by relieving the accuracy requirements and complexity of pre-launch calibration measurements. Cost / benefit rationale as well as conceptual approaches to pre-launch testing and on-orbit focus and vicarious calibration will be presented. | |
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| Paper Number RS5-2007-7001: A Network Broadcast Service for SpaceWire Plug and Play | |
| Allison Bertrand (Southwest Research Institute), Sandra G. Dykes (Southwest Research Institute), Robert Klar (Southwest Research Institute), Christopher C. Mangels (Southwest Research Institute) | |
| View/Download:Presentation | Paper | |
| Abstract: The foundation for low-cost responsive missions depends on the existence of standardized components with the capability to be quickly tailored to the current need. The automation of network configuration and communication services is necessary to rapidly assemble and deploy Plug and Play spacecraft. On-board communications over SpaceWire networks are currently being expanded to support the Plug and Play model. The term “Plug and Play” is used to describe automatic device recognition as well as automatic network configuration and the discovery of services at the application layer. For example, the Address Resolution Protocol (ARP) and the Dynamic Host Configuration Protocol (DHCP) can be considered Plug and Play services. In this paper, we focus on such automatic network services for SpaceWire. SpaceWire is a lightweight switched network that was developed for the space environment. It supports the current trend moving from shared bus architectures with mission-specific protocols towards on-board switched networks with established network protocols. The use of standard protocols such as Internet Protocol (IP) and operating system (OS) network stacks as part of the SpaceWire standard would reduce mission development time, errors, and cost. Ethernet is the dominant switched network technology on Earth and is one contender for the space environments. In Ethernet, the switches are responsible for link-layer broadcasts. To avoid forwarding loops, the switches must continuously learn network topology and build a minimum spanning tree for message delivery. However, these algorithms add complexity and circuitry which increases mass and power requirements. SpaceWire, an ESA standard, is gaining popularity because of its simple circuitry, low power consumption, and high-link speed. Despite SpaceWire’s many advantages, it does not have a link-layer broadcast mechanism. Broadcasts are necessary to support automatic network configuration and discovery software such as the ARP and the DHCP. Consequently, address resolution and host IP assignments for SpaceWire networks currently require manual configuration. This presentation will describe the design and implementation of a link-layer broadcast service for SpaceWire. Our link-layer broadcast mechanism is implemented in the hosts rather than in the SpaceWire switches. The design is compliant with the SpaceWire specification and can be implemented solely within the interface driver software. No changes are required to the SpaceWire specification, routers, or host interface hardware. The algorithm uses the concept of a SpaceWire subnet, which consists of a router and its directly connected hosts. The advantage of a subnet approach is that broadcast messages are propagated over a large multi-router SpaceWire network using only port addressing or the SpaceWire packet distribution mechanism. We believe that supporting standard protocols such as ARP, DHCP, and other broadcast-based protocols greatly simplifies the administration of spacecraft on-board networks and reduces mission development time. Standard communication protocols also provide compatibility and interoperability between a wide range of existing network applications. This is an important step towards building Plug-and-Play spacecraft with automatic network configuration, resulting in more rapid development and deployment of SpaceWire-based responsive missions. | |
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| Paper Number RS5-2007-7006: Dallas EEProm Equipment Profile for Rapid Integration and System Modeling | |
| Forrest Rogers-Marcovitz (Washington University in Saint Louis), Phelps Williams (Santa Clara University) | |
| View/Download:Presentation | Paper | |
| Abstract: One definition of Responsive Space is the ability for mission-specific payloads and support systems to be rapidly integrated within a short period. However, as components are added to the spacecraft, the complex interactions between subsystems must be noted and, if possible, modeled; this process is extremely time consuming and, when done poorly (or not at all), is a major contributor to spacecraft failure. A new paradigm is needed for Rapid Integration and System Modeling. At the 2006 Conference on Small Satellites in Logan Utah, Washington University and Santa Clara University demonstrated Rapid Integration and Testing by functionally combining their respective satellites, Akoya and Onyx; both vehicles were connected via a common power and data wiring harness, allowing one spacecraft to operate any device on either vehicle. Despite possessing minimal prior knowledge about the other school’s subsystems, functional integration was achieved in less than thirty minutes. Each satellite uses a distributed computing architecture with a standardized interface and communication protocol. This architecture allows each subsystem to be developed separately and rapidly integrated into the spacecraft. The success of this experience led to an improved design for subsystem-level embedded operational intelligence. The Dallas EEProm Equipment Profile (DEEP) Architecture extends this standardized bus to include improved support for rapid integration and system modeling. DEEP is a protocol standard using the Maxim/Dallas 1-Wire bus allowing for low level control and monitoring of the spacecraft using commercially-of-the-self devices including memory and sensor devices. DEEP specifies a standard with which subsystem functionality is encoded within the subsystem itself allowing for the creation of a satellite-wide model in parallel with physical integration of the spacecraft. This allows a stockpile of flight DEEP enabled subsystems, ready to be rapidly composed into a functional spacecraft. Each subsystem includes a subsystem model, with parameters such as thermal and power characteristics, allowing an anomaly management system to identify off-nominal conditions through model-based reasoning. DEEP is currently being developed at Santa Clara University and Washington University in Saint Louis as part of the University Nanosatellite competition operated by the Air Force Research Laboratory. This paper describes the current success of both universities with rapid integration, current development of the DEEP architecture, and future advances regarding responsive space. | |
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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 RS6-2008-2002: Responsive Coverage Using Propellantless Satellites | |
| George E. Pollock (Purdue University), Joseph W. Gangestad (Purdue University), James M. Longuski (Purdue University) | |
| View/Download:Presentation | Paper | |
| Abstract: Traditional reconnaissance satellites, which are fixed in their orbits, are limited in responding to rapidly-evolving conditions on the battlefield. For example, typical satellites cannot alter their time of arrival over a current region of interest. Further, as one conflict subsides and another emerges, these satellites cannot change their inclination to cover different latitudes. Thus, to support the warfighter in dynamic battlespaces, significant on-orbit maneuvering capability may be highly desirable. In this paper, we introduce the Lorentz spacecraft, a near-term propellantless vehicle, which can change orbit inclination, arrival time, altitude, and other orbit characteristics. The spacecraft modulates an electrostatic charge that interacts with Earth’s magnetic field to induce a propulsive Lorentz force. Assuming a conventional satellite power system (e.g. solar panels or RTGs), this spacecraft has inexhaustible maneuvering capability. In a matter of days a satellite’s orbit can be reconfigured to provide coverage of new theaters, perform flyby inspection of foreign space assets, and evade foreign tracking. We demonstrate the feasibility of Lorentz spacecraft by 1) characterizing the orbit dynamics of a charged spacecraft in Earth’s magnetic field, 2) deriving control laws for a variety of responsive space applications, and 3) providing an overview of hardware considerations and development efforts. | |
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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 RS6-2008-3007: A PCI-Based, Multiple-Payload Processing System for the DSX Flight Experiment” | |
| James King (QinetQ North America/PSI), Robert Gillis (QinetQ North America/PSI), Scott Greeley (QinetQ North America/PSI), Lawrence Davis (QinetQ North America/PSI), Lt. Col. Jon Schoenberg (Air Force Research Laboratory), Capt. Mark Scherbarth (Air Force Research Laboratory) | |
| View/Download:Presentation | Paper | |
| Abstract: The need rapidly to accommodate a variety of diverse payloads, each with its own data processing and command/telemetry protocols, is increasingly common in the emerging responsive space environment. On the AFRL Demonstration and Science Experiments (DSX) spacecraft, the Experiment Computer System (ECS) provides communication interfaces, mass data storage, and complex processing for ten different payloads, each from a different organization, some of legacy design and some developmental, while providing a single, simple interface to the host spacecraft. In this paper we discuss the hardware and software design of the ECS with a view to clarifying how a single architecture has supported and isolated the complexity of a variety of different payloads. Of equal importance, we show how an aggressive program of early integration testing, using hardware and software prototypes representing the payloads at various stages of development, has reduced the risk of payload integration, and isolated the host spacecraft from the vicissitudes of payload development. | |
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| Paper Number RS6-2008-6004: Programmable Satellite Transceiver (PST) for Responsive Space | |
| Jason Phillips (Real Time Logic, Inc.), Bill Asiano (Real Time Logic, Inc.) | |
| View/Download:Presentation | Paper | |
| Abstract: As the demand increases for more responsive lower costs space operations, the leadtime for fielding space assets will need to significantly decrease. The Air Force Research Laboratory (AFRL) Operation Responsive Space (ORS) initiative is pushing the boundary of quick deployment of space assets with an objective goal of the six day satellite. To meet the goals of ORS, the satellite must be adaptable to different missions, changing threats, and emerging technologies. In order to minimize integration efforts and meet these objectives, satellite subsystems will be intelligent modules that support a plug-¬in-play interface architecture between modules and the satellite bus. The Space Plug-and-Play Avionics (SPA) system in development by AFRL Space Vehicles Directorate addresses the space craft subsystems software and electrical interfaces. Two critical subsystem modules are the S-band Command, Control, and Telemetry (CC&T) and UHF/S-Band Mission Data radios. This paper introduces the Programmable Satellite Transceiver (PST) design concept for the radios. The PST as one of many Plug and Play (PnP) components supports the Space Plug-in-Play Avionics Universal Serial Bus (SPA¬U) interface. The PST design is adaptable to be used in standard satellite configurations, but it can also support AFRL’s ORS PnP configuration architecture. A PnP component contains Self¬Defining Data Sheet (xTEDS) which has all the data products, commands, interfaces, services, telemetry, and standard commands to define that subsystem. With a comprehensive intelligent protocol for each of the subsystems, satellite integration and test is significantly reduced and simplified. Traditional single frequency radios are set at the factory to a specific frequency and modulation type. The PST provides a modular software radio designed for space operations combining frequency agility and software configured signal processing functions in a re-programmable transceiver. Radiation tolerant parts and radiation mitigation techniques are used to enable the configurable operation in a space environment. For CC&T radio applications the modulator and demodulator can be independently tuned to any SGLS or USB frequency combination under software control. As a Mission Data radio both two-way UHF is supported along with S-Band transmit for TDRS applications using staggered PSK modulation schemes. All baseband Digital Signal Processing (DSP) functions are performed by a reconfigurable Xilinx Field Programmable Gate Array (FPGA). FPGA signal processing allows future upgrade to virtually any waveform set without hardware modifications. The ORS program objectives direct the development of technologies that support robust and flexible satellite bus designs. The PST using the SPA-U interface allows for rapid integration within these designs and provides a flexible yet modular and adaptive solution for CC&T/Mission Data radio functions. The PST design concept enables the ORS to provide a cost effective approach to rapid space asset deployment, operations, and maintenance over the life cycle of DoD space missions. | |
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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 RS6-2008-6002: Satellite Component Load Reduction Using SoftRide | |
| Raman Johal (CSA Engineering, Inc.), Paul S. Wilke (CSA Engineering, Inc.), Conor D. Johnson (CSA Engineering, Inc.) | |
| View/Download:Presentation | Paper | |
| Abstract: For Operationally Responsive Space (ORS) to be a success, a satellite must be able to launch into orbit on any available launch vehicle (LV) with only a few days (or weeks) notification. To further enable responsiveness, these spacecraft might need to have previously unplanned mission critical instruments or components added to the bus just prior to launch. For ORS to be as flexible as intended, these late-stage components must not fail due to the launch loads, yet these components most likely will not be included in a coupled loads analysis cycle to adequately quantify their risk of failure. To help enable ORS, one of two things must occur: a) prior to designing the component, all possible loads from all possible launch vehicles must be surveyed (since each has its own characteristic launch loads) and the mounts for the components must be designed structurally to withstand the highest possible set of loads, or b) a whole-spacecraft vibration and/or shock isolation system could be integrated between the spacecraft and the LV to greatly reduce the transmission of vibration and shock loads from whichever vehicle is chosen. The major drawback to the first solution is that the resulting components will be structurally over-designed and much heavier than necessary. The second solution offers much more flexibility for the mission commander in terms of a lighter, higher capability spacecraft; the reduced structural mass can be used more effectively in the form of more late-stage instruments onboard and the spacecraft can be launched on any available LV. Whole-spacecraft vibration and shock isolation systems (SoftRide) have been developed and flown to attenuate dynamic loads for several launch vehicles, and are currently in development for other launch vehicles and loading environments. To further enable responsive space, advanced knowledge of the dynamic environments of the launch vehicles and various size satellites could be gained by developing generic satellite models that cover the size ranges under consideration. Coupled-loads analyses could be performed that will develop the characteristics (stiffness, damping, and strength requirements) of a whole-spacecraft vibration and/or a shock isolation system that will mitigate the effects of the dynamic loads for each combination of satellite and LV. From these characteristics, a family of two, three, or four (if needed) isolation systems that will satisfy the requirements for all combinations could be developed and manufactured. The resulting isolators could be on-the-shelf and a simple chart of key spacecraft and LV characteristics could determine which SoftRide system to install in the field. Implementing a whole-spacecraft isolation system can provide benefits for responsiveness for other aspects of the mission, not just spacecraft design and launch. Using SoftRide during payload vibration testing will reduce the risk of experiencing a test failure, thereby decreasing delays due to repair and re-test. Including SoftRide in the spacecraft’s shipping container further protects against component failure due to shock and vibration loads from transport and handling. This will facilitate smooth spacecraft integration onto the LV. | |
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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 RS7-2009-3010: Nanosatellite Tracking Ships: Responsive, Seven-Month Nanosatellite Construction for a Rapid On-Orbit Automatic Identification System Experiment
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| Freddy Pranajaya (Space Flight Laboratory), Robert E. Zee (Space Flight Laboratory), Jeff Cain (COM DEV Limited), Richard Kolacz (COM DEV Limited) | |
| View/Download:Presentation | Paper | |
| Abstract: The Space Flight Laboratory (SFL) at the University of Toronto Institute for Aerospace Studies and COM DEV Ltd have developed a low Earth orbit nanosatellite in less than seven months to perform rapid turnaround experiments in space to detect and study Automatic Indentification System (AIS) signals transmitted by maritime vessels. The satellite, known as "Nanosatellite Tracking Ships" (NTS) leverages both SFL's CanX-2 nanosatellite technology and Generic Nanosatellite Bus (GNB) mechanical design to house a custom AIS receiver payload developed by COM DEV Ltd. NTS was developed under an extremely tight schedule, with on-orbit results required within a year from contract start. NTS have successfully met all of its mission objectives and continues to operate in orbit. This paper outlines how SFL and COM DEV were able to rapidly design, construct and deploy a custom satellite to respond to the opportunity to bring on-orbit AIS detection services to the international community. | |
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| Paper Number RS7-2009-2006: µSWaP INFOSEC Technology for ORS Secure Communication Links | |
| Ken Clauss (General Dynamics C4 Systems) | |
| View/Download:Presentation | Paper | |
| Abstract: General Dynamics C4 Systems (GDC4S) is a defense industry developer and supplier of Information Security (INFOSEC) subsystems for satellites and ground systems. ORS satellites use telemetry, tracking and commanding (TT&C) links and mission data links as part of their communications subsystems. These links need to be secured to prevent unauthorized control of the satellite, and to protect sensitive data as it is transmitted to the ground. Securing these links requires low size, weight, and power (SWaP) equipment to minimize burden on the ORS satellite. The equipment must also be space qualified and NSA certified. GDC4S is researching innovative methods used to meet the needs for a µSWaP, space qualified, NSA approved crypto device. This paper will discuss the trades in developing space qualifiable and NSA certifiable crypto, which includes the use of technology for mission flexibility (for example, ASICs vs. FPGAs), as well as the algorithms and keying requirements stipulated by NSA. Combining all these variables is non-trivial. The audience will gain an understanding about the details of implementation to meet the crypto security, while keeping the design as simple as possible to minimize SWaP. One implementation solution is the use of AES256 for the operational algorithm in space qualified radiation tolerant FPGA hardware to minimize power. This algorithm keeps the equipment unclassified until keyed prior to launch eliminating costly security controls until absolutely needed. The radiation tolerant space qualified FPGA’s also allow a single hardware design that can implement different functions (TT&C, mission data, etc.) to meet different missions as needed for various satellites. This paper highlights GDC4S’ efforts to develop an ultra-low SWaP ORS Crypto Module (OCM) with the aforementioned characteristics, including the use of use open standard interfaces for ease of integration, and data rate operation up to 300 Mbps. | |
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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 RS7-2009-3005: 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 RS7-2009-4002: 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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