General Information -- Computer System Architectures

General Information -- Computer System Architectures
General
Information -- Computer System Architectures
Company Name
SPACE PHOTONICS, INC.
Title
Geolocation and Attitude
Determination from Laser
Communication Systems
DOD SBIR Phase I
Quad Chart
To realize the benefits of formation flying multi-spacecraft clusters, four key elements must be present: 1. Broadband inter-spacecraft communications to enable the transport
of high resolution sensor data. 2. Precision inter-spacecraft timing synchronization to enable precise multi-sensor sampling throughout the cluster. 3. Precision inter-spacecraft
relative range to enable the determination of spacecraft and sensor spacing within the cluster. 4. Precision inter-spacecraft relative position and attitude determination to
enable geolocation of cluster craft and sensor planes. An inter-spacecraft laser communications network is the only integrated subsystem that can provide all four of these key
elements. The proposed SBIR Phase I effort will leverage several existing SPI programs to extend the resulting models, analyses, and designs beyond the normal scope of a
Phase I effort, and--should a follow-on Phase II be awarded--produce fully-functional prototype hardware at a minimum cost. Space Photonics and its subcontractors will
develop the analytical models, algorithms, calibration processes, and VHDL code for the inter-spacecraft ranging, timing synchronization and inter-spacecraft position and
attitude determination; and will (sub-scale) demonstrate these functions in our gimbal-less, free space optical, lasercom testbed (LaserFire® testbed). However, the
algorithms, calibration processes, and VHDL code developed for this SBIR Phase I program can be applied to any gimbaled or gimbal-less lasercom crosslink system. These
functions can even be applied to RF crosslink systems (the RF implementation will suffer significant loss of inter-spacecraft position accuracy due to the larger RF
beamwidths).
General
Thermal—Control Instrumentation
Company Name
Mainstream Engineering
Corp.
Title
Demonstration of a Plug and Play
Approach to Satellite Thermal
Control System Development
Field Center
GSFC
NASA SBIR Phase I
Quad Chart
Mainstream is proposing a methodology to reduce the development time and cost, and improve the reliability of future thermal control systems for the next
decade of deep space missions by utilizing a Plug-and-Play (PnP) thermal control strategy. A unique modular thermal control approach is described in the
proposal as well as discussions why such an approach is ideally suited to reducing the development time and cost of deep space satellite thermal control
systems as well as power conversion systems. Phase I will experimentally demonstrate that the proposed PnP approach can be configured to provide design
flexibility, long life and extremely high reliability. With potential missions to Europa, Venus, and Titan, as well as primitive bodies, the PnP approach can simplify
thermal control design while utilizing very compact, reliable, two-phase thermal control bus architecture. In this Phase I effort, Mainstream will demonstrate a
unique PnP Tool-Box approach where the thermal control system is easily built-up into a complete integrated system, as well as being integrated into the power
conversion components. Highly integrated systems will reduce cost, reduce system size, improve reliability, and tighten the thermal control.
General
Thermal – Cooling
Company Name
Advanced Cooling Technologies, Inc.
Title
PRESSURE CONTROLLED HEAT
PIPE FOR PRECISE
TEMPERATURE CONTROL
Field Center
GSFC
NASA SBIR Phase II
Quad Chart
The principal Phase II objective is to refine and further develop the prototype pressure controlled heat pipe (PCHP) into a useful thermal management tool.
The Phase I program established the feasibility of thermal control, an axially-grooved heat pipe with a variable-volume reservoir. The follow-on Phase II
program will address control system optimization, component longevity, reductions in mass and power, and show that the device can be flight qualified.
Nanosat Launch Vehicle Technologies
Avionics and Astrionics – Attitude Determination and Control
FC
Glenn Research Center
Title
Pulsed Plasma Thruster (PPT) Technology
Earth Observing-1 PPT Operational and
Advanced Components Being Developed
NTRS
Abstract
In 2002 the pulsed plasma thruster (PPT) mounted on the Earth Observing-1 spacecraft was operated successfully in orbit. The two-axis thruster system is fully
incorporated in the attitude determination and control system and is being used to automatically counteract disturbances in the pitch axis of the spacecraft. The first tests
conducted in space demonstrated the full range of PPT operation, followed by calibration of control torques from the PPT in the attitude control system. Then the
spacecraft was placed in PPT control mode. To date, it has operated for about 30 hr. The PPT successfully controlled pitch momentum during wheel de-spin, solar array
acceleration and deceleration during array rewind, and environmental torques in nominal operating conditions. Images collected with the Advanced Landsat Imager during
PPT operation have demonstrated that there was no degradation in comparison to full momentum wheel control. In addition, other experiments have been performed to
interrogate the effects of PPT operation on communication packages and light reflection from spacecraft surfaces. Future experiments will investigate the possibility of
orbit-raising maneuvers, spacecraft roll, and concurrent operation with the Hyperion imager. Future applications envisioned for pulsed plasma thrusters include longer life,
higher precision, multiaxis thruster configurations for three-axis attitude control systems or high-precision, formationflying systems. Advanced components, such as a "dry"
mica-foil capacitor, a wear-resistant spark plug, and a multichannel power processing unit have been developed under contract with Unison Industries, General Dynamics,
and C.U. Aerospace. Over the last year, evaluation tests have been conducted to determine power processing unit efficiency, atmospheric functionality, vacuum
functionality, thruster performance evaluation, thermal performance, and component life.
Nanosat Launch Vehicle Technologies
Avionics and Astrionics – Guidance, Navigation, and Control
Company Name
SiWave, Inc.
Title
ULTRA-SMALL, LOW-COST EARTH HORIZON SENSOR
Field Center
GSFC
NASA SBIR Phase II
Quad Chart
A MEMS Earth horizon sensor, when compared with current state-of-the-art Earth horizon sensors provides for three orders-of-magnitude
reductions in size, reduced power consumption, lower cost, and increased functionality and flexibility.
Nanosat Launch Vehicle Technologies
Avionics and Astrionics – Guidance, Navigation, and Control
USPTO
Patent
Number
US6380526
Title
Assignee
Abstract
Employing Booster
Trajectory in a Payload
Inertial Measurement Unit
Honeywell International
Inc.
A payload launching method measures the attitude of a booster during a launch with an inertial measurement unit on the payload to compare the actual
booster trajectory with a desired trajectory to control the payload at the deployment location, the inertial measurement unit on a booster carrying the
payload using said desired trajectory to reach the deployment location
US6550720
Aerobraking Orbit Transfer AeroAstro
Vehicle
US6921051
System for the Delivery
None
and Orbital Maintenance of
Micro Satellites and Small
Space-Based Instruments
US7246775
System and method of
substantially autonomous
geosynchronous timeoptimal orbit transfer
Lockheed Martin
Corporation
The excess space and weight capacity that is typical of a launch of large satellites to high-energy orbits, such as a geosynchronous orbit, is used to
deploy small satellites at a substantially lower-energy orbit, such as a low-earth orbit. An orbit-transfer vehicle provides the navigation, propulsion, and
control systems required to transport a payload satellite from a geosynchronous-transfer orbit (GTO) to a predetermined low-earth orbit (LEO).
Depending upon the particular configuration, upon achieving the low-earth orbit, the orbit transfer vehicle either releases the payload satellite, or
remains attached to the payload satellite to provide support services, such as power, communications, and navigation, to the payload satellite. To reduce
the fuel requirements for this deployment via the orbit-transfer vehicle, the orbit-transfer vehicle employs aerobraking to bring the satellite into a lowearth orbit. The aerobraking is preferably performed at a nominal altitude of 150 km above the earth, where the atmosphere is dense enough to allow for
a reasonably sized drogue device, yet rare enough to avoid the need for special purpose heat-shielding materials. In a preferred operation, the provider of
the orbit-transfer vehicle identifies and secures available excess capacity on geosynchronous-transfer launch vehicles, and allocates the excess capacity
to the satellites requiring low-earth orbit deployment, thereby providing a deployment means that is virtually transparent to the purchaser of this
deployment service.
A low cost, on demand, dedicated launch system is provided for placing micro satellites or space-based instruments at orbital and sub-orbital altitudes
and velocities. The invention describes a space launch vehicle (SLV) that incorporates a single, integrated guidance, navigation, and control unit
(GNCU) that performs all guidance and control for the SLV from main stage ignition to orbital insertion. The GNCU can remain with the payload after
orbital insertion to provide satellite station keeping and orbital maneuvering capability. The use of a single integrated avionics unit for all guidance,
navigation, and control simplifies the SLV, reducing weight and significantly reducing cost. In addition, this architecture allows for a combined launch
and satellite bus system as the GNCU can also be used as a satellite bus. This further reduces cost and increases the payload capacity to orbit by
optimizing the use of launch vehicle and satellite bus subsystems and reducing non-instrument mass delivered to orbit. All support functions are
provided by the IDMV. This approach represents a significant improvement over conventional systems, especially with respect to the orbital launch of
payloads less than about 100 kg.
Method of and system for on-board substantially autonomous control for transferring a spacecraft from an initial orbit to a final geosynchronous orbit, by
a trajectory that minimizes remaining transfer time and orbit transfer fuel. The spacecraft determines its orbit using a GPS-based system to determine the
spacecraft orbital elements. Based on the measured orbit error, corrected co-state parameters are calculated and used to generate an updated thrust
trajectory. The corrections are calculated using an innovative numerical procedure, carried out repetitively at a fixed interval until the target
geosynchronous orbit is achieved.
Nanosat Launch Vehicle Technologies
Avionics and Astrionics – Guidance, Navigation, and Control
U.S. Patent Applications
Patent Number
Title
US2002171011A1
System for the
Delivery and Orbital
Maintenance of
Micro Satellites and
Small Space-Based
Instruments
Assignee
None
Abstract
A low cost, on demand, dedicated launch system is provided for placing micro satellites or space-based instruments at orbital and sub-orbital altitudes and velocities.
The invention describes a space launch vehicle (SLV) that incorporates a single, integrated guidance, navigation, and control unit (GNCU) that performs all guidance
and control for the SLV from main stage ignition to orbital insertion. The GNCU can remain with the payload after orbital insertion to provide satellite station keeping
and orbital maneuvering capability. The use of a single integrated avionics unit for all guidance, navigation, and control simplifies the SLV, reducing weight and
significantly reducing cost. In addition, this architecture allows for a combined launch and satellite bus system as the GNCU can also be used as a satellite bus. This
further reduces cost and increases the payload capacity to orbit by optimizing the use of launch vehicle and satellite bus subsystems and reducing non-instrument mass
delivered to orbit. All support functions are provided by the IDMV. This approach represents a significant improvement over conventional systems, especially with
respect to the orbital launch of payloads less than about 100 kg.
Nanosat Launch Vehicle Technologies
Avionics and Astrionics – Guidance, Navigation, and Control
WIPO
Patent Number
Title
WO02077660A2
A System for the
Delivery and Orbital
Maintenance of Micro
Satellites and Small
Space-Based
Instruments
WO07024655A2
System and Method
for Propellantless
Photon Tether
Formation Flight
Assignee
Space Launch
Corporation
Abstract
A low cost, on demand, dedicated launch system is provided for placing micro satellites or space-based instruments at orbital and sub-orbital altitudes and
velocities. The invention describes a space launch vehicle (SLV) that incorporates a single, integrated guidance, navigation, and control unit (GNCU) that
performs all guidance and control for the SLV from main stage ignition to orbital insertion. The GNCU can remain with the payload after orbital insertion to
provide satellite station keeping and orbital maneuvering capability. The use of a single integrated avionics unit for all guidance, navigation, and control
simplifies the SLV, reducing weight and significantly reducing cost. In addition, this architecture allows for a combined launch and satellite bus system as
the GNCU can also be used as a satellite bus. This further reduces cost and increases the payload capacity to orbit by optimizing the use of launch vehicle
and satellite bus subsystems and reducing non-instrument mass delivered to orbit. All support functions are provided by the IDMV. This approach represents
a significant improvement over conventional systems, especially with respect to the orbital launch of payloads less than about 100 kg.
Bae Young K
The invention is a system and method for propellantless, ultrahigh precision satellite formation flying based on ultrahigh precision intracavity laser thrusters
and tethers with an intersatellite distance accuracy of nanometers at maximum estimated distances of tens of kilometers. The repelling force of the
intracavity laser thruster and the attracting force of tether tension between satellites form the basic forces to stabilize matrix structures of satellites. Users of
the present invention can also use the laser thruster for ultrahigh precision laser interferometric metrology, resulting in simplification and payload weight
reduction in integrating the thruster system and the metrology system.
Nanosat Launch Vehicle Technologies
Avionics and Astrionics – Telemetry, Tracking, and Control
Company Name
ASRC Research &
Technology
Solutions
Title
Autonomous Flight
Termination & Satellite
Based Telemetry System
for Launch Vehicles
Quasonix, LLC
Common Flight Test
Module (CFTM)
DOD SBIR Phase I
Quad Chart
Space Based Autonomous Flight Termination and Telemetry System (SBAFTTS) for launch vehicles will develop algorithms and hardware needed to implement a new tracking/telemetry
paradigm that enables lower cost and more responsive launches. Current space launch operations require an extensive network of ground-based tracking and communications sites that
provide support for a limited geographic area. These sites require long turnaround times and are costly to maintain due to obsolete equipment, staffing cost and remote locations. By using
space-based assets for tracking and communications along with autonomous intelligent software flight termination capabilities, future ranges will have lower cost, reduced complexity and
faster turnaround time while continuing to provide unsurpassed safety. This proposed system will provide tracking data to the onboard Autonomous Flight Termination System (AFTS) part of
this proposed system and to the ground over satellite without deploying down range assets. An optional Range Safety tracking and termination command link to/from the range launch head
and/or over satellite will be provided to add an extra margin of safety during the initial phase of launch, the most critical. The AFTS can have the primary flight termination responsibility or
take over if either the Range Safety communications is lost or when out of range.
All DoD aircraft, airborne munitions, and weapon systems are subjected to extensive flight testing. This testing requires a core set of instrumentation onboard the test article, including TimeSpace Position Information (TSPI), flight termination (FT), and telemetry (TM) functions. At present, each test article employs a combination of unique modules to support these capabilities.
The time and expense to develop, flight-qualify, and maintain these unique modules is enormous, so a common module which provided these standard functions for a variety of test articles
would result in substantial improvements in efficiency within the flight test community. The goal of the Subminiature Flight Safety System (SFSS) program is just such a common module.
Designing a common module to support weapon system flight testing is a conceptually straightforward task. However, the steady progression toward ever-smaller munitions makes the
transition from concept to reality risky. The emergence of new paradigms for some capabilities, such as the two-way telemetry links envisioned for the integrated Network Enhanced
Telemetry (iNET) program, adds further complexity and technical risk. The research proposed here will greatly mitigate the technical, cost, and schedule risks associated with the
development of future range safety and telemetry systems such as SFSS.
Nanosat Launch Vehicle Technologies
Avionics and Astrionics – Telemetry, Tracking, and Control
FC
DFRC/KSC
Title
Ku- and Ka-Band Phased Array Antenna for the
Space-Based Telemetry and Range Safety
Project
NTRS
Abstract
The National Aeronautics and Space Administration Space-Based Telemetry and Range Safety study is a multiphase project to increase data rates and flexibility and
decrease costs by using space-based communications assets for telemetry during launches and landings. Phase 1 used standard S-band antennas with the Tracking and
Data Relay Satellite System to obtain a baseline performance. The selection process and available resources for Phase 2 resulted in a Ku-band phased array antenna
system. Several development efforts are under way for a Ka-band phased array antenna system for Phase 3. Each phase includes test flights to demonstrate performance
and capabilities. Successful completion of this project will result in a set of communications requirements for the next generation of launch vehicles.
Nanosat Launch Vehicle Technologies
Avionics and Astrionics – Telemetry, Tracking, and Control
Company Name
NAL Research
Corporation
Title
Small Satellite
Transceiver for Launch
Vehicles
Field Center
GSFC
NASA SBIR Phase I
Quad Chart
NAL Research Corporation proposes to develop a small, light-weight, low-cost transceivers capable of establishing satellite communications links for telemetry and control during
the launch and ascent stages of flight. The proposed transceiver will offer continuous and truly global coverage. When data are sent from a launch vehicle, the signals are received
immediately by one of the LEO satellites and relayed in real-time to command and control center via either Public Switched Telephone Network/Public Data Networks (PSTN/PDN),
directly to another transceiver, through the Internet or through a direct IP address. The entire process can take a fraction of a second. This will provide electronic global access to
airborne vehicles from any place.
Nanosat Launch Vehicle Technologies
Electronics –Highly Reconfigurable
FC
Goddard Space Flight
Center
Title
SpaceWire Plug and Play
NTRS
Abstract
The ability to rapidly deploy inexpensive satellites to meet tactical goals has become an important goal for military space systems. In fact, Operationally Responsive Space (ORS) has
been in the spotlight at the highest levels. The Office of the Secretary of Defense (OSD) has identified that the critical next step is developing the bus standards and modular interfaces.
Historically, satellite components have been constructed based on bus standards and standardized interfaces. However, this has not been done to a degree, which would allow the rapid
deployment of a satellite. Advancements in plug-and-play (PnP) technologies for terrestrial applications can serve as a baseline model for a PnP approach for satellite applications. Since
SpaceWire (SpW) has become a de facto standard for satellite high-speed (greater than 200Mbp) on-board communications, it has become important for SpW to adapt to this Plug and
Play (PnP) environment. Because SpW is simply a bulk transport protocol and lacks built-in PnP features, several changes are required to facilitate PnP with SpW. The first is for Host(s)
to figure out what the network looks like, i.e., how pieces of the network, routers and nodes, are connected together network mapping, and to receive notice of changes to the network.
The second is for the components connected to the network to be understood so that they can communicate. The first element, network topology mapping and amp change of status
indication, is being defined (topic of this paper). The second element describing how components are to communicate has been defined by ARFL with the electronic data sheets known as
XTEDS. The first element, network mapping, is recent activities performed by Air Force Research Lab (ARFL), Naval Research Lab (NRL), NASA and US industry (Honeywell, Clearwater,
FL, and others). This work has resulted in the development of a protocol that will perform the lower level functions of network mapping and Change Of Status (COS) indication required by
Plug 'n' Play over SpaceWire. This work will be presented to the SpaceWire working group for standardization under European Cooperation for Space Standardization (ECSS) and to
obtain a permanent Protocol ID (see SpaceWire Protocol ID What Does it Mean to You IEEE Aerospace Conference 2006). The portion of the Plug 'n' Play protocol that will be described
in this paper is how the Host(s) of a SpaceWire network map the network and detect additions and deletions of devices on a SpaceWire network.
Nanosat Launch Vehicle Technologies
Electronics –Highly Reconfigurable
Company Name
AeroAstro, Inc.
Title
FLEXIBLE AND EXTENSIBLE BUS FOR SMALL SATELLITES
(FEBSS)
DOD SBIR Phase II
Quad Chart
AeroAstro proposes to further develop modular stackable spacecraft system; the Plug'n'Sense software architecture standard which allows spacecraft
subsystems to interact without a great deal of custom hardware or software; the Universal Small Payload Interface to make launch vehicle integration
more transparent; and finally, the Flexible Extensible Bus for Small Satellites.
Nanosat Launch Vehicle Technologies
Information – Data Acquisition and End-to-End Management
U.S. Patent Applications
Patent Number
Title
US2006192057A1
Spacecraft Adapter
Having Embedded
Resources, and
Methods of Forming
Same
Assignee
None
Abstract
A spacecraft adapter having embedded resources for supporting a non-primary payload on a launch vehicle. The spacecraft adapter includes a battery, a power
distribution and control system, and an interface circuit for interfacing with the non-primary payload. Other modules/subsystems such as data storage, sensor and data
interface and communications may be included to suit the needs of a particular non-primary payload and/or particular mission of the non-primary payload. The
adaptor does not require any interfacing with the bus of the primary payload and can be scaled/modified as needed to provide only that degree of functionality needed
for a given non-primary payload being carried by the launch vehicle.
Nanosat Launch Vehicle Technologies
Information – Expert Systems
FC
ARC
Title
Distributed Web-Based Expert System for
Launch Operations
NTRS
Abstract
The simulation and modeling of launch operations is based on a representation of the organization of the operations suitable to experiment of the physical, procedural, software,
hardware and psychological aspects of space flight operations. The virtual test bed consists of a weather expert system to advice on the effect of weather to the launch operations. It
also simulates toxic gas dispersion model, and the risk impact on human health. Since all modeling and simulation is based on the internet, it could reduce the cost of operations of
launch and range safety by conducting extensive research before a particular launch. Each model has an independent decision making module to derive the best decision for
launch.
Nanosat Launch Vehicle Technologies
Information – Software Tools for Distributed Analysis and Simulation
Company Name
Mobile Foundations, Inc.
Title
AN AUTOMATED TOOL TO ENABLE THE DISTRIBUTED
OPERATIONS OF AIR FORCE SATELLITES
DOD SBIR Phase II
Quad Chart
Mobile Foundations developed and deployed the software that it prototyped in Phase I, when it showed the feasibility of building an
enhanced version of its Spacecraft Emergency Response System (SERS) software, which NASA uses for autonomous satellite
operations. The enhanced SERS will meet the more rigorous demands of Air Force operations laid out in the Air Force Space
Command Strategic Master Plan.
Nanosat Launch Vehicle Technologies
Materials -- Composites
FC
GRC
Title
SiC SiC Ceramic Matrix Composites Developed for HighTemperature Space Transportation Applications
NTRS
Abstract
Researchers at the NASA Glenn Research Center have been developing durable, high-temperature ceramic matrix composites (CMCs) with silicon carbide (SiC)
matrices and SiC or carbon fibers for use in advanced reusable launch vehicle propulsion and airframe applications in the Next Generation Launch Technology
(NGLT) Program. These CMCs weigh less and are more durable than competing metallic alloys, and they are tougher than silicon-based monolithic ceramics.
Because of their high specific strength and durability at high temperatures, CMCs such as C SiC (carbon- fiber-reinforced silicon carbide) and SiC SiC (silicon-carbidefiber-reinforced silicon carbide) may increase vehicle performance and safety significantly and reduce the cost of transporting payloads to orbit.
Nanosat Launch Vehicle Technologies
Materials – Metallics
Company Name
Plasma Processes, Inc.
Title
INNOVATIVE TUNGSTEN ALLOYS
FOR ADVANCED PROPULSION
SYSTEMS
Field Center
MSFC
NASA SBIR Phase II
Quad Chart
Innovative process for fabricating net shape, tungsten-rhenium-hafnium carbide alloy components are proposed. Development of these materials will
allow the production of components with unique properties and reduce the size, weight, and cost of propulsion systems.
Nanosat Launch Vehicle Technologies
Materials – Multifunctional/Smart Materials
FC
GRC
Title
Epoxy Crosslinked Silica Aerogels (XAerogels)
GRC
Mechanically Strong Lightweight
Materials for Aerospace Applications
(x-aerogels)
NTRS
Abstract
NASA is interested in the development of strong lightweight materials for the dual role of thermal insulator and structural component for space vehicles freeing more weight for useful
payloads. Aerogels are very-low density materials (0.010 to 0.5 g cc) that, due to high porosity (meso- and microporosity), can be, depending on the chemical nature of the network, ideal
thermal insulators (thermal conductivity approx. 15 mW mK). However, aerogels are extremely fragile. For practical application of aerogels, one must increase strength without compromising
the physical properties attributed to low density. This has been achieved by templated growth of an epoxy polymer layer that crosslinks the "pearl necklace" network of nanoparticles the
framework of a typical silica aerogel. The requirement for conformal accumulation of the epoxy crosslinker is reaction both with the surface of silica and with itself. After cross-linking, the
strength of a typical aerogel monolith increases by a factor of 200, in the expense of only a 2-fold increase in density. Strength is increased further by coupling residual unreacted epoxides
with diamine.
The X-Aerogel is a new NASA-developed strong lightweight material made by reacting the mesoporous surfaces of 3-D networks of inorganic nanoparticles with polymeric crosslinkers. Since
the relative amount of the crosslinker and the backbone are comparable, X-Aerogels can be viewed either as aerogels modified by templated accumulation of polymer on the skeletal
nanoparticles, or as nanoporous polymers made by templated casting of polymeric precursors on a nanostructured framework. The most striking feature of X-Aerogels is that for a nominal 3fold increase in density (still a ultralightweight material), the mechanical strength can be up to 300 times higher than the strength of the underlying native aerogel. Thus, X-Aerogels combine a
multiple of the specific compressive strength of steel, with the thermal conductivity of styrofoam. XAerogels have been demonstrated with several polymers such as polyurethanes polyureas,
epoxies and polyolefins, while crosslinking of approximately 35 different oxide aerogels yields a wide variety of dimensionally stable, porous lightweight materials with interesting structural,
magnetic and optical properties. X-Aerogels are evaluated for cryogenic rocket fuel storage tanks and for Advanced EVA suits, where they will play the dual role of the thermal insulator
structural material. Along the same lines, major impact is also expected by the use of X-Aerogels in structural components thermal protection for small satellites, spacecrafts, planetary
vehicles and habitats.
Nanosat Launch Vehicle Technologies
Propulsion – Beamed Energy
Company Name
Exquadrum, Inc.
Title
Technology to Enable Rapid Application of Laser
Propulsion
DOD SBIR Phase I
Quad Chart
The objective of the proposed research and development effort is to demonstrate the feasibility of laser propelled systems with greatly simplified hardware
and improved performance. This will be accomplished through application of an advanced propellant system. This technology will help to enable a
responsive, low-cost launch system for micro-satellites, based on the combination of a high efficiency multi-stage electromagnetic gun and laser beam
propulsion. The technology will be experimentally demonstrated during the research program.
Nanosat Launch Vehicle Technologies
Propulsion – Beamed Energy
Company Name
Physics, Materials & Applied Math
Research
Title
SPACE PROPULSION THROUGH ULTRASHORT PULSE
LASER ABLATION
DOD SBIR Phase II
Quad Chart
This test matrix can identify the thrust that can be obtained in space, using both individual pulses and pulse trains from nanosecond
and/or Joule-level ultra short pulse lasers. This information will identify missions that can be supported by laser-propulsion
techniques either from ground- or space-based sources.
Nanosat Launch Vehicle Technologies
Propulsion – Chemical
Company Name
Advanced
Mechanical
Technology, Inc
Title
High Efficiency
Regenerative Helium
Compressor
Field Center
GRC
Pioneer
Astronautics
Nitrous Paraffin Hybrid
MSFC
NASA SBIR Phase I
Quad Chart
Helium plays several critical rolls in spacecraft propulsion. High pressure helium is commonly used to pressurize propellant fuel tanks. Helium cryocoolers can be used to subcool and thereby densify cryogenic propellants such as liquid hydrogen (LH2) and liquid oxygen (LO2). The use of densified cryogenic propellants can reduce the gross
payload weight of a launch vehicle by up to 20%, or increase payload capability. Helium compressors are critical components for cryogenic propellant storage and distribution
systems, whether used in cryocoolers for densification or to compress gaseous helium for propellant pressurization. Regenerative compressor technology can serve high
head, low flow helium pressurization applications in a compact form with high reliability. Pressure ratios on the order of 3:1 per impeller-stage are commercially available. Nonlubricated gas-bearing supported prototypes have been successfully demonstrated. However, even state-of-the-art prototype regenerative compressors are limited to
efficiencies of about 55%. This was achieved using aerodynamic rotor blades rather than the straight radial blades previously used. Commercially available regenerative
compressors with straight vaned rotors operate at much lower efficiency. An innovation is proposed that promises to improve the efficiency of regenerative compressors well
beyond the current state of the art.
The Nitrous Oxide Paraffin Hybrid engine (N2OP) is a proposed technology designed to provide small launch vehicles with high specific impulse, indefinitely storable
propulsion. In the N2OP engine, the combination of liquid nitrous oxide on solid paraffin as a rocket propellant allows for the development of compact lightweight high
performance stages using densely packed propellant tankage. This is because N2O/paraffin hybrids have a very high oxidizer/fuel mixture ratio and because paraffin has a
much higher regression rate than typical hybrid hydrocarbon fuels. Propellant slumping can be prevented by molding the paraffin into a 3% by volume graphite sponge matrix.
Currently, space launch missions require cryogenic or extremely toxic propellants which are limited in their storage times, reducing their capability for rapid response launch.
The much more storable solid propellants have higher cost, and lower performance while still being a large explosive hazard. The N2OP propulsion system also is compatible
with ocean temperatures, allowing launch by floating in water. The achievable Isp for this propellant combination using autogenous pressurization is about 235 seconds at sea
level and over 310 s in vacuum, making its performance fully adequate to support operation of a safe, fully storable, highly-responsive multi-stage launch vehicle.
Nanosat Launch Vehicle Technologies
Propulsion – Chemical
Company Name
CFD Research Corporation
Sierra Engineering Inc.
Title
HIGH ENERGY, LOW
TEMPERATURE GELLED BIPROPELLANT FORMULATION
FOR LONG-DURATION IN-SPACE
PROPULSION
TRIAXIAL SWIRLER LIQUID
INJECTOR DEVELOPMENT
Field Center
MSFC
MSFC
NASA SBIR Phase II
Quad Chart
The use of gelled propellants for deep space planetary missions may enable adoption of high performance (Isp-vac>360 sec) propellant combinations that
do not require power-intensive heating and stirring cycles before firings, and whose handling and safety characteristics are close to stated goals of "green"
propellants. Phase II will culminate in a hot-fire demonstration of a GLP/GMON-30 rocket chamber, to be performed at AMRDEC facilities.
The triaxial swirl injector is ideally suited to a wide range of liquid oxidizers and fuels, including hydrogen and a wide range of hydrocarbons. It holds the
potential of excellent high-frequency combustion stability characteristics and combustion chamber thermal compatibility. The design is well suited for both
main injector and pre-burner applications.
Nanosat Launch Vehicle Technologies
Propulsion – Chemical
Company Name
Orbital Research Inc.
Title
AN ACTIVE THRUST VECTORING (ATV)
CONTROL SYSTEM FOR TACTICAL
MISSILE STEERING
SpaceDev
SMALL SHUTTLE-COMPATIBLE
PROPULSION MODULE
DOD SBIR Phase II
Quad Chart
The program will focus on the application of low-power, light-weight, mechanical control actuators to achieve thrust vector control with limited thrust
losses. Phase II will focus on the development of an optimal Active Thrust Vectoring control system for 3D nozzles to fit inside the 7” diameter missile
hardware to ensure compatibility with the current air-to-air missiles.
Phase II will pursue the development of a full-scale prototype MTV Propulsion Module incorporating a hybrid motor that will demonstrate the
functional characteristics and technology of an MTV hybrid motor. Our proposed Propulsion Module is a scalable, affordable and modular design that
utilizes safe, storable propellants.
Nanosat Launch Vehicle Technologies
Propulsion – Chemical
USPTO
Patent Number
Title
US6360993
Expendable
Launch Vehicle
Assignee
Space Systems/ Loral,
Inc.
Abstract
A system for supplying a space station or satellite comprises a low cost expendable single-stage to orbit launch vehicle having a moderate reliability in the
range of about 0.5 to 0.8 for launching a payload of consumable items to low earth orbit. For propulsion, the launch vehicle includes a pressure-fed rocket
engine, preferably with a nozzle for aiming products of combustion in a direction away from the launch vehicle and having a fixed orientation relative to the
launch vehicle. A pressurant tank contains an inert pressurant gas under pressure. A conduit system serves to introduce the inert pressurant gas from the
pressurant tank to a fuel tank and thence to the rocket engine and also to introduce the inert pressurant gas from the pressurant tank to an oxidizer tank and
thence to the rocket engine for combustion with the fuel. The initial pressure in the pressurant tank is maintained at a level in excess of about 100 bar; the
initial pressure in the fuel tank and in the oxidizer tank is maintained at a level in the range of about 8.5 to 20.0 bar; and the initial pressure in the rocket
engine is maintained at a level in the range of about 5 to 10 bar. Additional conduits are provided for adding inert pressurant gas to the pressurant tank or for
removing pressurant gas therefrom, for adding fuel to the fuel tank or for removing fuel therefrom, and for adding oxidizer to the oxidizer tank or for
removing oxidizer therefrom.
Nanosat Launch Vehicle Technologies
Propulsion – Chemical
U.S. Patent Applications
Patent Number
Title
US2004231766A1
Propulsion Device, Flying Object
Comprising the Same and
Propulsion Device Igniting Method
US2007044450A1
Powder Propellant-Based Space
Propulsion Device
None
Disclosed is a powder propellant-based space propulsion device using a powder propellant having high density and excellent handleability. The powder
propellant-based space propulsion device comprises a powder-propellant storage container for storing a powder propellant, a powder-propellant attracting
surface for attracting the powder propellant thereto through an opening of the powder-propellant storage container and attractively holding the attracted
powder propellant thereon, powder-propellant transfer means for transferring the held powder propellant to a release position for releasing the powder
propellant, and propulsive-energy supply means for energizing the transferred powder propellant to release the powder propellant from the powderpropellant attracting surface, toward a downstream side thereof as a propulsive jet, while accelerating the powder propellant in a direction approximately
perpendicular to the powder-propellant attracting surface at said release position. The powder-propellant transfer means is designed to move the powderpropellant attracting surface in such a manner that a powder-propellant holding area of the powder-propellant attracting surface is returned to a position
adjacent to the opening of the powder-propellant storage container in a repetitive manner.
US2007075190A1
Solid Propellant-Based Space
Propulsion Device
None
Disclosed is a space propulsion device capable of continuous operation using a solid propellant having high density and excellent handleability. The solid
propellant-based space propulsion device comprises solid-propellant support means for supporting a solid propellant, a solid-propellant attachment surface
for permitting the solid propellant to be attached thereon, solid-propellant feed means for feeding the solid propellant to an attachment position on the
solid-propellant attachment surface, adhesion-energy supply means for heatingly melting or sublimating the solid propellant in the attachment position to
adherently attach the solid propellant onto the solid-propellant attachment surface, solid-propellant transfer means for transferring the solid propellant to a
release position for releasing the solid propellant, and propulsive-energy supply means for energizing the transferred solid propellant to release the solid
propellant from the solid-propellant attachment surface at the release position, as a propulsive jet, while accelerating the solid propellant in a direction
approximately perpendicular to the solid-propellant attachment surface. The solid-propellant transfer means is designed to move the solid-propellant
attachment surface in such a manner that a area of the solid-propellant attachment surface for adherently holding the solid propellant is returned to a
position adjacent to an end of the solid propellant in a repetitive manner.
US2007144140A1
High Propellant Mass Fraction
Hybrid Rocket Propulsion
None
A chemical hybrid propulsion motor can be comparable in performance to a solid motor or liquid fueled engine if it uses cryogenic nitrous oxide as its
oxidizer and a pump such as a turbopump to transfer the oxidizer into the motor case. Cryogenic nitrous oxide (at about -100° F.) has high density which
will reduce oxidizer tank volume and weight and this oxidizer combusts at a high oxidizer to fuel (O/F) ratio which minimizes motor case size and weight.
The high O/F would also reduce unburnt fuel sliver and hence residual propellant weight. The pump not only transfers the oxidizer into the motor at
increased chamber pressure that in turn increases specific impulse, but it also significantly reduces oxidizer tank pressure and weight as compared to a
pressure fed motor.
Assignee Abstract
None
A propulsion device operated in a bi-propellant mode uses a safe hydrogen peroxide as oxidizer and yet has high specific impulse and high response
performance. A preheated net 18 is provided in a combustion chamber 14. Both of an oxidizer supply means 10 and a fuel supply means 12 open toward
the net 18. Oxidizer 30 and fuel 32 are atomized on the net 18 to thereby increase the surface area. At the same time, the oxidizer 30 and fuel 32 are heated
on the net 18 and their decomposition is accelerated. By quickly effecting collision and mixing of the oxidizer 30 with the fuel 32, an instantaneous
ignitability can be obtained.
Nanosat Launch Vehicle Technologies
Propulsion – Chemical
FC
MSFC
Title
Auxiliary Propulsion Activities in
Support of NASA's Exploration
Initiative
MSFC
Demonstration of a Non-Toxic
Reaction Control Engine
MSFC
Ignition Characterization Tests
of the LOX Ethanol Propellant
Combination
Marshall
Space
Flight
Center
In-Space Chemical Propulsion
System Model
NTRS
Abstract
The Space Launch Initiative (SLI) procurement mechanism NRA8-30 initiated the Auxiliary Propulsion System Main Propulsion System (APS MPS) Project in 2001 to address technology gaps and
development risks for non-toxic and cryogenic propellants for auxiliary propulsion applications. These applications include reaction control and orbital maneuvering engines, and storage, pressure
control, and transfer technologies associated with on-orbit maintenance of cryogens. The project has successfully evolved over several years in response to changing requirements for re-usable
launch vehicle technologies, general launch technology improvements, and, most recently, exploration technologies. Lessons learned based on actual hardware performance have also played a
part in the project evolution to focus now on those technologies deemed specifically relevant to the Exploration Initiative. Formal relevance reviews held in the spring of 2004 resulted in authority for
continuation of the Auxiliary Propulsion Project through Fiscal Year 2005 (FY05), and provided for a direct reporting path to the Exploration Systems Mission Directorate. The tasks determined to be
relevant under the project were continuation of the development, fabrication, and delivery of three 870 lbf thrust prototype LOX ethanol reaction control engines the fabrication, assembly, engine
integration and testing of the Auxiliary Propulsion Test Bed at White Sands Test Facility and the completion of FY04 cryogenic fluid management component and subsystem development tasks
(mass gauging, pressure control, and liquid acquisition elements). This paper presents an overview of those tasks, their scope, expectations, and results to-date as carried forward into the
Exploration Initiative.
Three non-toxic demonstration reaction control engines (RCE) were successfully tested at the Aerojet Sacramento facility under a technology contract sponsored by the National Aeronautics and
Space Administration's (NASA) Marshall Space Flight Center (MSFC). The goals of the NASA MSFC contract (NAS8-01109) were to develop and expand the technical maturity of a non-toxic, onorbit auxiliary propulsion system (APS) thruster under the auspices of the Exploration Systems Mission Directorate. The demonstration engine utilized Liquid Oxygen (LOX) and Ethanol as
propellants to produce 870 lbf thrust. The Aerojet RCE's were successfully acceptance tested over a broad range of operating conditions. Steady state tests evaluated engine response to varying
chamber pressures and mixture ratios. In addition to the steady state tests, a variety of pulsing tests were conducted over a wide range of electrical pulse widths (EPW). Each EPW condition was
also tested over a range of percent duty cycles (DC), and bit impulse and pulsing specific impulse were determined for each of these conditions. Subsequent to acceptance testing at Aerojet, these
three engines were delivered to the NASA White Sands Test Facility (WSTF) in April 2005 for incorporation into a cryogenic Auxiliary Propulsion System Test Bed (APSTB). The APSTB is a test
article that will be utilized in an altitude test cell to simulate anticipated mission applications. The objectives of this APSTB testing included evaluation of engine performance over an extended duty
cycle map of propellant pressure and temperature, as well as engine and system performance at typical mission duty cycles over extended periods of time. This paper provides acceptance test
results and a status of the engine performance as part of the system level testing.
A series of contracts have been issued by the Marshall Space Flight Center (MSFC) of the National Aeronautics and Space Administration (NASA) to explore candidate technologies considered to
be important for the Next Generation Launch Technology (NGLT) effort. One aspect of the NGLT effort is to explore the potential of incorporating non-toxic propellants for Reaction Control
Subsystems (RCS). Contract NAS8-01109 has been issued to Aerojet to develop a dual thrust Reaction Control Engine (RCE) that utilizes liquid oxygen and ethanol as the propellants. The dual
thrust RCE incorporates a primary thrust level of 870 lbf, and a vernier thrust level of 10 - 30 lbf. Aerojet has designed and tested a workhorse LOX igniter to determine LOX Ethanol ignition
characteristics as part of a risk mitigation effort for the dual thrust RCE design. The objective of the ignition testing was to demonstrate successfid ignition from GOX to LOX, encompassing potential
two-phase flow conditions. The workhorse igniter was designed to accommodate the full LOX design flowrate, as well as a reduced GOX flowrate. It was reasoned that the initial LOX flow through
the igniter would flash to GOX due to the inherent heat stored in the hardware, causing a reduced oxygen flowrate because of a choked, or sonic, flow condition through the injection elements. As
LOX flow continued, the inherent heat of the test hardware would be removed and the hardware would chill-in, with the injected oxygen flow transitioning from cold GOX through two-phase flow to
subcooled LOX. Pressure and temperature instrumentation permitted oxygen state points to be determined, and gas-side igniter chamber thermocouples provided chamber thermal profile
characteristics. The cold flow chamber pressure (P(sub c)) for each test was determined and coupled with the igniter chamber diameter (D(sub c)) to calculate the characteristic quench parameter
(P(sub c) x D(sub c)), which was plotted as a function of core mixture ratio, MR(sub c). Ignition limits were determined over a broad range of valve inlet conditions, and ignition was demonstrated
with oxygen inlet conditions that ranged from subcooled 173 R LOX to 480 R GQX. Once ignited at cold GOX conditions, combustion was continuous as the hardware chilled in and the core mixture
ratio transitioned from values near 1.0 to over 12.5.
Multiple, new technologies for chemical systems are becoming avai lable and include high temperature rockets, very light propellant tanks and structures, new bipropellant and monopropellant
options, lower mass propellant control components, and zero boil off subsystems. Such technologies offer promise of increasing the performance of in-space chemical propulsion for energetic
space missions. A mass model for pressure-fed, Earth and space-storable, advanced chem cal propulsion systems (ACPS) was developed in support of the NASA MSFC In-Space Propulsion
Program. Data from flight systems and studies defined baseline system architectures and subsystems and analyses were formulated for parametric scaling relationships for all ACPS subsystems.
The paper will first provide summary descriptions of the approaches used for the systems and the subsystems and then present selected analyses to illustrate use of the model for missions with
characteristics of current interest.
Nanosat Launch Vehicle Technologies
Propulsion – Chemical
FC
Marshall
Space
Flight
Center
Title
Reciprocating Feed System
Development Status
MSFC
Test Results for a Non-toxic,
Dual Thrust Reaction Control
Engine
NTRS
Abstract
The reciprocating feed system (RFS) is an alternative means of providing high pressure propellant flow at low cost and system mass, with high fail-operational reliability. The RFS functions by
storing the liquid propellants in large, low-pressure tanks and then expelling each propellant through two or three small, high-pressure tanks. Each RFS tank is sequentially filled, pressurized,
expelled, vented, and refilled so as to provide a constant, or variable, mass flow rate to the engine. This type of system is much lighter than a conventional pressure fed system in part due to the
greatly reduced amount of inert tank weight. The delivered payload for an RFS is superior to that of conventional pressure fed systems for conditions of high total impulse and it is competitive with
turbopump systems, up to approximately 2000 psi. An advanced version of the RFS uses autogenous pressurization and thrust augmentation to achieve higher performance. In this version, the
pressurization gases are combusted in a small engine, thus making the pressurization system, in effect, part of the propulsion system. The RFS appears to be much less expensive than a
turbopump system, due to reduced research and development cost and hardware cost, since it is basically composed of small high- pressure tanks, a pressurization system, and control valves. A
major benefit is the high reliability fail-operational mode in the event of a failure in one of the three tank-systems, it can operate on the two remaining tanks. Other benefits include variable pressure
and flow rates, ease of engine restart in micro-gravity, and enhanced propellant acquisition and control under adverse acceleration conditions. We present a system mass analysis tool that accepts
user inputs for various design and mission parameters and calculates such output values payload and vehicle weights for the conventional pressure fed system, the RFS, the Autogenous
Pressurization Thrust Augmentation (APTA) RFS, and turbopump systems. Using this tool, a preliminary design of a representative crew exploration vehicle (CEV) has been considered. The design
parameters selected for a representative system were modeled after the orbital maneuvering system (OMS) on the Shuttle Orbiter, with an increase of roughly a factor of ten in the delta- V
capability and a greater thrust (30,000 lbs, vs. 12,000 lbs). Both storable and cryogenic propellants were considered. Results show that a RFS is a low mass alternative to conventional pressure fed
systems, with a substantial increase in payload capability and that it is weight-competitive with turbopump systems at low engine pressure (a few hundred psi) at high engine pressures, the APTA
RFS appears to offer the highest payload. We also present the status of the RFS test bed fabrication, assembly, and checkout. This test bed is designed to provide flow rates appropriate for engines
in the roughly 10,000 to 30,000 lb thrust range.
A non-toxic, dual thrust reaction control engine (RCE) was successfully tested over a broad range of operating conditions at the Aerojet Sacramento facility. The RCE utilized LOX Ethanol
propellants and was tested in steady state and pulsing modes at 25-lbf thrust (vernier) and at 870-lbf thrust (primary). Steady state vernier tests vaned chamber pressure (Pc) from 0.78 to 5.96 psia,
and mixture ratio (MR) from 0.73 to 1.82, while primary steady state tests vaned Pc from 103 to 179 psia and MR from 1.33 to 1.76. Pulsing tests explored EPW from 0.080 to 10 seconds and DC
from 5 to 50 percent at both thrust levels. Vernier testing accumulated a total of 6,670 seconds of firing time, and 7,215 pulses, and primary testing accumulated a total of 2,060 seconds of firing
time and 3,646 pulses.
Nanosat Launch Vehicle Technologies
Propulsion – Electromagnetic Thrusters
Company Name
Phoenix Nuclear Labs
Title
Non-ambipolar Electron Source
Field Center
GRC
NASA SBIR Phase I
Quad Chart
A device to produce electron beams from magnetized plasma created with rf fields combined with electron extraction by electron sheaths is proposed. The source
can provide electrons for neutralizing positive ion beams emerging from ion thrusters or as a generic electron source. With hollow cathode sources currently
employed to provide neutralizing electrons, operation is limited in time and/or current density by cathode deterioration. RF electron sources provide an alternative
approach that does not consume electrode material. The current from this Non-ambipolar Electron Source (NES) exceeds the current normally extracted from
conventional rf plasma sources by a factor of (mi/me)1/2 where mi and me are the ion and electron mass. Ions are lost to a negatively biased conducting cylinder
with area Ai chosen to be Ai ≥ (mi/me)1/2 *Ae where Ae is the electron extraction area. Slots in the conducting cylinder allow the cylinder to serve as a Faraday
shield to reduce capacitive coupling from the antenna to the plasma. Proposed phase 1 design improvements should result in electron currents comparable to
hollow cathode sources with lower neutral gas flow in the inductive discharge phase and higher currents with helicon operation. Phase 2 will develop prototype
sources suitable for spacecraft testing.
Nanosat Launch Vehicle Technologies
Propulsion – Electromagnetic Thrusters
Company Name
Busek Co. Inc.
Title
SHARED MAGNETICS HALL THRUSTER
Field Center
MSFC
NASA SBIR Phase II
Quad Chart
To fulfill the full range of Hall thruster power requirements (100 kW/1 MW) envisioned by NASA for orbit insertion,
planetary transfers and manned exploration, rather than developing a >500kW system with one or two large thrusters there
are clear advantages to reach the very high power by clustering multiple thrusters of lower power.
Nanosat Launch Vehicle Technologies
Propulsion – Electromagnetic Thrusters
Company Name
Busek Company, Inc. and
Worcester Polytechnic
Institute
Title
COMPACT INDUCED CURRENT
HALL THRUSTER
DOD SBIR Phase II
Quad Chart
This Phase II STTR program will investigate a new kind of electric thruster, an inductively-driven Hall thruster. In contrast to conventional Hall thrusters, this device needs no
cathode. We expect it to be more efficient and smaller. It is also quite different from existing inductive thrusters, because it has a magnetic core and a bias field, has much
longer–duration pulses, runs at lower ionization and can be made much smaller.
Busek Company, Inc.
HIGH CURRENT CATHODE
DEVELOPMENT
High power dual mode Hall effect thruster system is an enabling technology for the orbital transport of DoD and commercial space assets. A dual mode Hall thruster is a
propulsion system that is capable of efficient operation in a high thrust to power mode as well as a high specific impulse mode.
Busek Company, Inc.
EXTENDED LIFETIME LOW
POWER HALL THRUSTERS
The focus of the Phase II effort is to extend the lifetime of the BHT-200 to 3000 hours making it attractive to a broader range of users. In the Phase II, Busek will continue with
design modifications, primarily to the neutral flow distribution and magnetic field profiles that minimize the impingement of ions onto the insulator surfaces.
Exquadrum, Inc.
DUAL MODE PROPULSION
MODULE TECHNOLOGY FOR
SPACE CONTROL
The Dual Mode Propulsion Module Technology for Space Control concept utilizes an innovative approach to rocketry in which a single thruster can function as a solid rocket
motor or as an electric thruster. The resulting system is highly flexible and hence able to perform a wide variety of space control missions. In this project, a full-scale thruster will
be fabricated and experimentally demonstrated.
Nanosat Launch Vehicle Technologies
Propulsion – Electromagnetic Thrusters
U.S. Patent Applications
Patent Number
Title
US2004149861A1
S & H Cycle Engine
Assignee
None
Abstract
The S \A H Cycle Engine will provide very high specific impulse propulsion at a reasonable thrust for fast human or cargo or scientific spacecraft travel to the
moon or other planets and/or moons of other planets. Significantly shorter flight times will reduce astronaut sickness from zero gravity effects and space radiation.
The same system will also provide enough electrical energy to potentially power an electromagnetic shield to protect human passengers from the energetic
charged particle component of solar radiation throughout the flight. The S \A H Cycle Engine differs from other electric propulsion devices in that it flows the
thruster cryogenic propellant through a heat exchanger before it reaches the thruster as a bottom beat sink for the thermal process that generates the electricity for
the electric thrusters. In this manner the thermal efficiency increases significantly, thus increasing the total propulsion system efficiency and hence increase the
velocity or payload mass of the spacecraft.
Nanosat Launch Vehicle Technologies
Propulsion – Electromagnetic Thrusters
USPTO
Patent Number
Title
US6523338
Plasma Accelerator
Arrangement
Assignee
Abstract
Thales Electron Devices Gmbh For a plasma accelerator arrangement in particular for use as an ion thruster in a spacecraft, a structure is proposed in connection with which an
accelerated electron beam is admitted into an ionization chamber with fuel gas, and is guided through the ionization chamber in the form of a
focused beam against an electric deceleration field, said electric deceleration field acting at the same time as an acceleration field for the fuel ions
produced by ionization. The arrangement generates a focused beam of a largely neutral plasma with a high degree of efficiency. Configurations for
electric and magnetic fields for guiding and focusing the beams are given by way of example.
Nanosat Launch Vehicle Technologies
Propulsion – Feed System Components
USPTO
Patent Number
Title
US6378291
Reduced toxicity fuel
satellite propulsion
system including
catalytic decomposing
element with hydrogen
peroxide
US6546714
Reduced toxicity fuel
satellite propulsion
system including
plasmatron
Assignee
The United States of America as
represented by the Administrator
of the National Aeronatics and
Space Administration
Abstract
A reduced toxicity fuel satellite propulsion system including a reduced toxicity propellant supply (10) for consumption in an axial class
thruster (14) and an ACS class thruster (16). The system includes suitable valves and conduits (22) for supplying the reduced toxicity
propellant to the ACS decomposing element (26) of an ACS thruster. The ACS decomposing element is operative to decompose the
reduced toxicity propellant into hot propulsive gases. In addition the system includes suitable valves and conduits (18) for supplying the
reduced toxicity propellant to an axial decomposing element (24) of the axial thruster. The axial decomposing element is operative to
decompose the reduced toxicity propellant into hot gases. The system further includes suitable valves and conduits (20) for supplying a
second propellant (12) to a combustion chamber (28) of the axial thruster, whereby the hot gases and the second propellant auto-ignite and
begin the combustion process for producing thrust.
The United States of America, as
represented by the Administrator
of the National Aeronautics and
Space Administration
A reduced toxicity fuel satellite propulsion system including a reduced toxicity propellant supply (10) for consumption in an axial class
thruster (14) and an ACS class thruster (16). The system includes suitable valves and conduits (22) for supplying the reduced toxicity
propellant to the ACS decomposing element (26) of an ACS thruster. The ACS decomposing element is operative to decompose the
reduced toxicity propellant into hot propulsive gases. In addition the system includes suitable valves and conduits (18) for supplying the
reduced toxicity propellant to an axial decomposing element (24) of the axial thruster. The axial decomposing element is operative to
decompose the reduced toxicity propellant into hot gases. The system further includes suitable valves and conduits (20) for supplying a
second propellant (12) to a combustion chamber (28) of the axial thruster, whereby the hot gases and the second propellant auto-ignite and
begin the combustion process for producing thrust.
Nanosat Launch Vehicle Technologies
Propulsion – Feed System Components
U.S. Patent Applications
Patent Number
Title
US2002139901A1
X33 Aeroshell and Bell Nozzle
Rocket Engine Launch Vehicle
Assignee
The Aerospace Corporation
Abstract
Various launch vehicles configurations each include an X33 aeroshell of a booster or orbiter both of which use bell nozzle engines,
and a feeding stage for supplying liquid propellant to the engines for providing primary lifting thrust for lifting a payload into
orbit. The feeding stage can be an external tank without engines or a core vehicle also with bell nozzle engines. The orbiter or
booster use three, two-two or five bell nozzle engines configurations and the feeding stage uses a zero or two bell nozzle engines.
The combination of orbiters, boosters, external tanks and core vehicles offer a variety of configuration to meet particular mission
requirements.
Nanosat Launch Vehicle Technologies
Propulsion – Feed System Components
WIPO
Patent Number
Title
WO05005255A1
Aerospace Vehicle with
Separate Propellant
Reservoir or Energy
Source
Assignee
Ab Technologies (Ab),S.R.L.,
Ancarola, Biagio
Abstract
This invention proposes an aerospace transportation system and operation method that applies to launch vehicles, sounding rockets,
spacecraft, aeroplanes and helicopters. In this system the mass of propellant, or the energy source necessary to feed the propulsion
system in order to generate thrust, is kept separated from the vehicle body. In this way the thrust generated by the propulsion
system of the vehicle body can be fully used to accelerate the vehicle body and not the propellant, or energy source. The propellant,
or energy transfer to the propulsion system of the vehicle body occurs through: -An umbilical tube (2a) that connects the propellant
tank (3a) the the vehicle body (1a), in the case of the propellant or energy for the generation of thrust being fluid. - Otherwise
through air or vacuum, if the energy is transmittable via air or vacuum. - Otherwise through electric-magnetic connectors (2d, 4d)
that link the vehicle body to the energy source necessary to generate thrust.
Nanosat Launch Vehicle Technologies
Propulsion – High Energy Propellants
U.S. Patent Apps
Patent Number
Title
US2002036038A1
High Regression Rate Hybrid
Rocket Propellants and Method
of Selecting
US2003019204A1
US2003098397A1
Assignee
None
Abstract
This invention comprises a new process for developing high regression rate propellants for application to hybrid rockets and solid
fuel ramjets. The process involves the use of a criterion to identify propellants which form an unstable liquid layer on the melting
surface of the propellant. Entrainment of droplets from the unstable liquid-gas interface can substantially increase propellant mass
transfer leading to much higher surface regression rates over those that can be achieved with conventional hybrid propellants. The
main reason is that entrainment is not limited by heat transfer to the propellant from the combustion zone. The process has been used
to identify a new class of non-cryogenic hybrid fuels whose regression rate characteristics can be tailored for a given mission. The
fuel can be used as the basis for a simpler hybrid rocket design with reduced cost, reduced complexity and increased performance.
Propulsion System
Electron Power Systems, Inc.
The propulsion system of the present invention includes charged particles held in a containment structure at high temperatures.
Propellant is injected into the containment system where it collides elastically with the charged particles. The propellant is heated
until it achieves a high velocity. It can then be ejected to produce thrust. The propulsion system can include thrusters to propel, for
example, space launch vehicles and jet aircraft.
Method for Placing Payload in
Orbit By Multifunctional Launch
Vehicle Of Combined Scheme
with Cruise Liquid Rocket
Engine System (Lres),
Multifunctional Launch Vehicle
Of Combined Scheme with
Cruise Lres and Method for
Refining It
None
[From equivalent US6712319] The invention relates to a space-rocket technology and can be used for placing both a pilotcontrolled and unpiloted space craft in the earth's orbit. The inventive method consists in igniting-up cruising liquid-fuel rocket
engines (LRE) of all assembly units, when the propulsion power of LRE of the side units is set to maximum and the LRE of the
central unit is set relatively low.
Nanosat Launch Vehicle Technologies
Propulsion – High Energy Propellants
WIPO
Patent Number
Title
WO28010180A2
Launching a Flight Vehicle
Assignee
Spacego Technologies (Proprietary)
Limited
Abstract
This invention relates to launching of a flight vehicle, such as a rocket, an aircraft and the like, from a generally upwardly
directed subterranean launch shaft of at least 300 m length, preferably of a length of 500 m or more. Preferably, the launch
shaft is an amortized mine shaft. Launching is by means of one of, or a combination of, several possible propulsion systems.
For each system, at least some components are external of the flight vehicle and associated with or mounted in the launch
shaft. One kind of propulsion system may use expansion of gas in the launch shaft behind the flight vehicle. Other propulsion
systems use electromagnetism or Lorentz-type actuators, such as a coil gun or gauss gun, a linear induction motor or linear
synchronous motor, or a rail gun. The launch shaft mounts guiding means for guiding the flight vehicle along the launch
shaft.
Nanosat Launch Vehicle Technologies
Propulsion – Launch Assist
U.S. Patent Applications
Patent Number
Title
US20050247230A1
Artillery Shell, Satellite
Launcher, & Global Reach
Missile
Assignee
None
Abstract
The present invention relates to artillery shell range, cannon fired satellite, missiles, and communication satellite launcher. Cannon shell
range can be increased by decreasing the drag through making the flat base shape into inverted cone or by fastening a wood or
lightweight plastic cone. Fastening a booster filled with gun powder to the base of the shell with a delay primer fuse that explodes in
flight get the recoil push the trajectory higher in the air and faster in speed to reach a longer range. Adding multiple boosters in tandem
causes successive explosions in flight that the successive recoils speeds up the shell in-flight and increases the range. Fastening a round
thin steel smooth barrel to the base of the shell and stacking multiple flat sides or cone shaped boosters separated by felt or plastic
separators with center holes filled with slow burning primer adds more blank-type recoil to enchance shell speed and range. Fastening a
long round steel barrel to the base of the shell or trajectory body to fired from a cannon or from a silos can give the trajectory enough
speed and enough fuel to become a long range missile with a bomb tip or can become a satellite body that can reach orbiting level. This
starting or in-flight booster and boosters system can power a long range rocket, a satellite launcher, or a fast-travel vehicle that is
starting at ground or fired from a high flying plane.
Nanosat Launch Vehicle Technologies
Propulsion – Micro Thrusters
Company Name
Busek Co. Inc.
Title
Field-Effect Modulated Electro-Osmotic Pumps for High
Precision Colloid Thrusters
Field Center
JPL
NASA SBIR Phase I
Quad Chart
The ability to precisely control the position of satellites is a critical enabling technology for space missions involving
interferometric arrays. One proposed mission, LISA (Laser Interferometer Space Antenna), would use an array of 3 satellites
whose relative position is monitored and controlled to an accuracy of 10 nm. Precise station-keeping such as this demands
precise, high stability thrusters supplied with propellant flows on the order of microliters/min and producing micro-newtons of
thrust. These requirements are difficult or impossible to meet with traditional thrusters and feed systems such as cold-gas
thrusters or monopropellants. The proposed program will evaluate the use of electro-osmosis to supply and control the flow of
ionic liquid propellants to micronewton colloid thrusters. In addition, the use of a gate electrode to control the surface charge and
therefore the magnitude and direction of flow will be examined as will the use of AC fields to limit electrolysis effects. Phase I will
provide basic information on the electro-osmotic behavior of ionic liquids using simple test devices and electrospray emitters.
Phase II will involve detailed design work to fabricate a practical propellant feed system using electro-osmotic pumps.
Nanosat Launch Vehicle Technologies
Propulsion – Micro Thrusters
Company Name
Busek Co. Inc.
Title
RADIO FREQUENCY MICRO ION THRUSTER FOR PRECISION
PROPULSION
Tethers Unlimited
TECHNOLOGIES FOR MOMENTUMEXCHANGE/ELECTRODYNAMIC-REBOOST
TETHER FACILITIES
NASA SBIR Phase II
Field Center
Quad Chart
JPL
Busek proposes to continue development of an engineering model radio frequency discharge, gridded micro
ion thruster that produces sub-mN to mN thrust precisely adjustable over a wide dynamic thrust range.
MSFC
The MXER Tether Boost Facility will serve as a fully-reusable in-space upper-stage that will provide
propellantless propulsion for orbital transfer and Earth-to-Orbit launch assist. By eliminating the need to launch
an upper stage along with each payload, the MXER Tether can reduce the size of the launch vehicle needed
to deploy the payloads, achieving dramatic reductions in total launch costs.
Nanosat Launch Vehicle Technologies
Propulsion – Micro Thrusters
USPTO
Patent Number
Title
US6378292
Mems Microthruster
Array
US6483368
Assignee
Honeywell International Inc.
Abstract
A microelectrical mechanical system (MEMS) microthruster array is disclosed. The MEMS microthruster array of the present invention can be
used for maintaining inter satellite distance in small satellites. One microthruster array includes numerous microthruster propulsion cells, each
having a vacuum enclosed explosive igniter disposed on one side by a breakable diaphragm and having a propellant-filled chamber on the
opposite side of the diaphragm. Upon explosion of the explosive igniter, the first diaphragm breaks, which, together with the explosion of the
explosive igniter, causes the propellant to expand rapidly, thereby providing exhaust gases which are ejected from an exterior face of the
microthruster propulsion array, thereby providing a small unit of thrust.
Addressable diode
isolated thin film cell
array
The Aerospace Corporation
An address element, including a polysilicon resistor functioning as a heating element and blocking diode preventing sneak current to
unaddressed elements, is selectively addressed using row and column address lines in a thin film structure having a minimum number of address
lines and a minimum number of layers. The resistor heater element is well suited for igniting a fuel cell such as a fuel cell in an array of fuel
cells disposed in a thin film microthruster.
US6487844
Aerospike
augmentation of
microthruster impulse
TRW Inc.
A microthruster (11) provides a small unit force. The microthruster is useful, for example, as a propellant for a microsatellite. The microthruster
includes an aerospike (12) extending outwardly beyond the face of an outer wall of the chamber (14) of the microthruster in the vicinity of an
outlet nozzle thereof formed by two diaphragms or burst disks (15, 16), closing the chamber on respective sides of the aerospike. The aerospike
is preferably formed integrally with the chamber by batch microelectronic fabrication methods. Higher thrust efficiency and more controllable
and uniform impulse characteristics are attainable with the microthruster and with an array (17) comprising a plurality of the microthrusters.
US6494402
Lateral exhaust
microthruster
The Aerospace Corporation
A microthruster having an inverted exhaust system traps burst diaphragm fragments providing a clean exhaust while an exhaust port provides
increased back pressure for efficient combustion of a propellant charge in a fuel cell. A converging diverging micronozzle provides a predictable
exhaust vector for improved microthrusting well suited for propulsion system on small spacecraft.
Nanosat Launch Vehicle Technologies
Propulsion – Micro Thrusters
U.S. Patent Applications
Patent Number
Title
US2002023427A1
Micro-Colloid Thruster
System
US20040245406A1
Micropump-Based
Microthruster
Assignee
None
Abstract
A micro-colloid thruster system may be fabricated using micro electromechanical system (MEMS) fabrication techniques. A beam of
charged droplets may be extracted from an emitter tip in an emitter array by an extractor electrode and accelerated by an accelerator
electrode to produce thrust. The micro-colloid thruster system may be used as the main propulsion system for microspacecraft and for
precision maneuvers in larger spacecraft.
None
A thruster for providing thrust for spacecraft positioning, which has a propellant reservoir for storing propellant, a reaction chamber for
discharging a vapor for providing thrust, a pump module comprising one or more micropumps for drawing propellant from the reservoir and
for systematically metering propellant to the reaction chamber in a controlled manner, and a controller for actuating the pump module.
Nanosat Launch Vehicle Technologies
Propulsion – Solar
FC
Marshall Space Flight Center
Title
Application of Solar-Electric
Propulsion to Robotic Missions in
Near-Earth Space
NTRS
Abstract
Interest in applications of solar electric propulsion (SEP) is increasing. Application of SEP technology is favored when (1) the mission is compatible with low-thrust
propulsion, (2) the mission needs high total delta V such that chemical propulsion is disadvantaged and (3) performance enhancement is needed. If all such
opportunities for future missions are considered, many uses of SEP are likely. Representative missions are surveyed and several SEP applications selected for
analysis, including orbit raising, lunar science and robotic exploration, and planetary science. These missions span SEP power range from 10 kWe to about 100 kWe.
A SEP design compatible with small inexpensive launch vehicles, and capable of lunar science missions, is presented. Modes of use and benefits are described, and
potential SEP evolution is discussed.
Nanosat Launch Vehicle Technologies
Sensors and Sources – Optical
Company Name
Physical Sciences, Inc.
Title
DEMONSTRATION OF INFLATABLE
REFLECTOR TECHNOLOGY FOR
PICOSATELLITE APPLICATIONS
DOD SBIR Phase II
Quad Chart
PSI proposes to use a masked inflatable optic to image a well characterized light source onto a CCD array. Analysis of these images enables the
measurement of surface accuracy to within 20 microns, enabling us to verify static and dynamic models of the systems thermal and inflation properties.
Nanosat Launch Vehicle Technologies
Structures – Kinematic-Deployable
Company Name
Planning Systems, Inc. (Part of
QinetiQ)
Title
A REDUNDANT DEPLOYMENT
CONTROLLER AND EXPERIMENT
COMPUTER SYSTEM FOR SMALL
SATELLITE EXPERIMENTS
DOD SBIR Phase II
Quad Chart
The Phase II program will develop the deployment controller to a prototype experiment computer and deployment control system applicable to the Deployed
Structures Experiment (DSX). The control systems will serve as the sequencer for deployment control as well as handling command, control and data
acquisition functions for DSX.
Nanosat Launch Vehicle Technologies
Structures – Kinematic-Deployable
FC
MSFC
Title
Development of a Deployable
Nonmetallic Boom for Reconfigurable
Systems of Small Modular Spacecraft
MSFC
Development of a Deployable
Nonmetallic Boom for Reconfigurable
Systems of Small Spacecraft
NTRS
Abstract
Launch vehicle payload capacity and the launch environment represent two of the most operationally limiting constraints on space system mass, volume, and configuration. Large-scale
space science and power platforms as well as transit vehicles have been proposed that greatly exceed single-launch capabilities. Reconfigurable systems launched as multiple small
modular spacecraft with the ability to rendezvous, approach, mate, and conduct coordinated operations have the potential to make these designs feasible. A key characteristic of these
proposed systems is their ability to assemble into desired geometric (spatial) configurations. While flexible and sparse formations may be realized by groups of spacecraft flying in close
proximity, flyers physically connected by active structural elements could continuously exchange power, fluids, and heat (via fluids). Configurations of small modular spacecraft
temporarily linked together could be sustained as long as needed with minimal propellant use and reconfigured as often as needed over extended missions with changing requirements.
For example, these vehicles could operate in extremely compact configurations during boost phases of a mission and then redeploy to generate power or communicate while coasting
and upon reaching orbit. In 2005, NASA funded Phase 1 of a program called Modular Reconfigurable High-Energy Technology Demonstrator Assembly Testbed (MRHE) to investigate
reconfigurable systems of small spacecraft. The MRHE team was led by NASA's Marshall Space Flight Center and included Lockheed Martin's Advanced Technology Center (ATC) in
Palo Alto and its subcontractor, ATK. One of the goals of Phase 1 was to develop an MRHE concept demonstration in a relevant 1-g environment to highlight a number of requisite
technologies. In Phase 1 of the MRHE program, Lockheed Martin devised and conducted an automated space system assembly demonstration featuring multipurpose free-floating
robots representing Spacecraft in the newly built Controls and Automation Laboratory (CAL) at the ATC. The CAL lab features a 12' x 24' granite air-bearing table and an overhead
simulated starfield. Among the technologies needed for the concept demo were mating interfaces allowing the spacecraft to dock and deployable structures allowing for adjustable
separation between spacecraft after a rigid connection had been established. The decision to use a nonmetallic deployable boom for this purpose was driven by the MRHE concept
demo requirements reproduced in Table 1.
In 2005, NASA commenced Phase 1 of the Modular Reconfigurable High Energy Technology Demonstrator (MRHE) program to investigate reconfigurable systems of small spacecraft.
During that year, Lockheed Martin's Advanced Technology Center (ATC) led an accelerated effort to develop a 1-g MRHE concept demonstration featuring robotic spacecraft simulators
equipped with docking mechanisms and deployable booms. The deployable boom built for MRHE was the result of a joint effort in which ATK was primarily responsible for developing
and fabricating the Collapsible Rollable Tube (CRT patent pending) boom while Lockheed Martin designed and built the motorized Boom Deployment Mechanism (BDM) under a
concurrent but separate IR and amp;D program. Tight coordination was necessary to meet testbed integration and functionality requirements. This paper provides an overview of the
CRT boom and BDM designs and presents preliminary results of integration and testing to support the MRHE demonstration.
Nanosat Launch Vehicle Technologies
Structures – Launch and Flight Vehicle
USPTO
Patent
Number
US6360994
Title
Assignee
Abstract
Configurable Space
Launch System
Don A. Hart &
Associates, Inc.
A configurable space launch system of multiple different vehicle configurations that use a common reusable spaceplane and cost effective external tanks is
presented. Each vehicle configuration in the system incorporates one or more reusable spaceplanes and most or all ascent propellant in multiple releasable
external tanks. The flight trajectory of each vehicle has multiple in-flight staging points to increase vehicle performance efficiency. The system is
governed and configured by a unique set of eight prescripts that together minimize launch costs.
US6409124
High-Energy to LowEnergy Orbit Transfer
Vehicle
None
US6430920
Nozzleless Rocket
Motor
Technanogy, Llc
US6508435
Method for
Controlling an
Aerospace System to
Put a Payload into an
Orbit
Karpov; Anatoly
Stepanovich, Rachuk;
Vladimir Sergeevich,
Ivanov; Robert
Konstantinovich
US6508437
Launch Lock for
Spacecraft Payloads
Honeywell
International Inc.
The excess space and weight capacity of a conventional launch vehicle for a high-energy orbit, such as GEO, is used to deploy satellites to a low-energy
orbit, such as LEO. In a preferred embodiment, an orbit-transfer vehicle provides the navigation, propulsion, and control systems required to transport a
payload satellite from a high-energy-transfer orbit, such as GTO, to a predetermined low-energy orbit. Upon entering the low-energy orbit, the payload
satellite is released from the orbit-transfer vehicle. To reduce the fuel requirements for this deployment via the orbit-transfer vehicle, a preferred
embodiment includes aerobraking to bring the satellite into a low-earth orbit. In a preferred embodiment of this method of deployment, the provider of the
orbit-transfer vehicle identifies and secures available excess capacity on launch vehicles, and allocates the excess capacity to the satellites requiring lowearth orbit deployment, thereby providing a deployment means that is virtually transparent to the purchaser of this deployment service.
A solid rocket motor for accelerating a payload comprises a motor casing and a solid propellant matrix, utilizing a high burn-rate fuel. The use of a high
burn rate fuel allows the rocket motor to operate in an end-burning fashion without the use of a constricting aperture to increase the back-pressure upon the
burn-front of the fuel matrix. The exhaust gas produced from combustion of the propellant matrix exits directly to the ambient environment through a
simple aperture without the use of an expansion nozzle. By eliminating the mass of the nozzle and allowing the use of lighter, less structurally robust
motor casings, the needed acceleration of the vehicle can be achieved while using less propellant and a lighter launch vehicle.
Aerospace technology, in particular, methods for orbital injection of payloads (communication satellites, monitoring satellites, etc.) into low and medium
earth orbits with the aid of aerospace systems comprising a carrier aircraft (CA) and a launch vehicle (LV) with a payload (PL). In the present method,
after the CA 1 takeoff from the base aerodrome, its flight in the maximum cruising speed mode to the LV 2 launch area, pitchdown 7 of the CA 1 is
effected to gain the maximum permissible horizontal flight speed, and at the moment that speed is attained, CA 1 pitchup 8 with the maximum allowable
angle of attack is executed, culminating in transition to an angle of attack with a near-zero g-load (zero-gravity condition), with the pitchup parameters
chosen so that at the LV 2 with PL 3 point of separation from the CA 1, the CA 1 has attained a speed V.sub.D, flight altitude H.sub.D and a trajectory
pitch angle O.sub.d ensuring a maximum PL 3 and a near-zero normal g-load 9. Separation of the LV 2 from the CA 1 is executed, imparting the LV 2
with a CA 1 related speed assuring that the LV 2 lags behind the CA 1 at a safe distance 10, then the LV 2 sustainers are fired, and, either prior to the
sustainers' ignition, or after the ignition (by sustainers), the launch vehicle with the PL 3 is turned into a position differing form the vertical by an angle of
10-30 .degree. in the vertical plane in the launch direction 11. At the LV 2-PL 3 separation point 22, the CA 1 position is stabilized in an inertial
coordinate system. The present invention makes it possible to safely, efficiently and with maximum payload capacity, place payloads into designated
orbits and deliver payloads to given terrestrial and oceanic areas.
A spacecraft payload, suspended by an isolator, is locked down during launch by using a compliant member to pull on a pin, clamping the payload against
the launch lock. The payload is unlocked, after launch, by heating a shape memory alloy to move the pin to overcome the compliant member, releasing the
payload from launch lock. The payload may be re-locked for reentry by heating another shape memory alloy to push the pin against the other shaped
memory alloy.
Nanosat Launch Vehicle Technologies
Structures – Launch and Flight Vehicle
USPTO
Patent
Number
US6543715
Title
Assignee
Abstract
Aerospace System
Karpov; Anatoly
Stepanovich, Rachuk;
Vladimir Sergeevich,
Ivanov; Robert
Konstantinovich
US6547476
Universal Spacecraft
Separation Node
Lockheed Martin
Corporation
US6612522
Flyback Booster with
Removable Rocket
Propulsion Module
Starcraft Boosters, Inc.
An aerospace system, which comprises a carrier aircraft, a launch vehicle and a payload. The launch vehicle with liquid fuel propulsion units is placed
inside a fuselage of the carrier aircraft within a transport and launch container with the aid of support units in at least in two areas, an airtight pneumatic
chamber is formed between the blind end of the transport and launch container and the end of the launch vehicle, which chamber incorporates airborne
elements of units for supplying the launch vehicle with propellant and working mediums, elements of a draining unit, elements of a unit for replenishing
the liquid fuel propulsion units with propellant, electrical connections, wherein all the units, electrical connections and airborne elements are connected to
the end of the launch vehicle by pull-off couplings, wherein the transport and launch container is provided with a pneumatic ejection unit, made in the
form of a high pressure source connected by stop valves to the pneumatic chamber and positioned in the transport and launch container, provided with
thermal insulation, wherein the diameter of the free end the container, provided with a frangible membrane, is hermetically connected to the perimeter of
the port in the fuselage of the carrier aircraft. The proposed system makes it possible to enhance payload capacity and to reduce the specific cost of putting
a payload in orbit, to provide a high degree of safety for the carrier aircraft and its crew, and also to ensure ecological safety of the system.
An integrated separation system and interface structure is disclosed for a variety of deployment applications. In one embodiment, a Universal Spacecraft
Separation Node (100) includes a separation nut assembly (102), a separation spring assembly (104) an LV node fitting (106) and an SC node fitting (108).
The LV node fitting (106) is connected to a launch vehicle (210) and the SC node fitting (108) is connected to a spacecraft (216). The separation nut
assembly (102) holds the fittings (106 and 108) together until separation is desired. Upon separation, the separation spring assembly (104) provides a force
to urge the launch vehicle (210) and spacecraft (216) apart. Prior to separation, an annular tongue (224) of fitting (108) mates with an annular groove (226)
of fitting (106) to resist shear forces.
A flyback booster (200) comprising an aircraft (203) housing a launch vehicle stage as a removable rocket propulsion module (502) and several space
launch vehicles using variations of the flyback booster (200) are disclosed. This flyback booster (200) functions as the first stage of a multistage space
launch vehicle. The stage used in the flyback booster (200) and the upper stages of the multistage space launch vehicle (213) are selected to optimize the
launch cost for a specific payload.
US6685141
X33 aeroshell and bell
nozzle rocket engine
launch vehicle
The Aerospace
Corporation
US6845949
System and Methods
The Boeing Company
for Integrating a
Payload with a Launch
Vehicle
US6905097
Launch Vehicle
Payload Carrier and
Related Methods
The Boeing Company
Various launch vehicles configurations each include an X33 aeroshell of a booster or orbiter both of which use bell nozzle engines, and a feeding stage for
supplying liquid propellant to the engines for providing primary lifting thrust for lifting a payload into orbit. The feeding stage can be an external tank
without engines or a core vehicle also with bell nozzle engines. The orbiter or booster use three, two-two or five bell nozzle engines configurations and the
feeding stage uses a zero or two bell nozzle engines. The combination of orbiters, boosters, external tanks and core vehicles offer a variety of
configuration to meet particular mission requirements.
A system for providing an interface between a launch vehicle and a payload of the launch vehicle includes payload integration points of the launch vehicle
that remain unchanged relative to different payload configurations, such that a payload is configurable apart from the launch vehicle. The system makes it
possible to de-couple payload installation and launch vehicle turnaround operations. Payloads can be configured off-line from launch vehicle processing,
thus making it possible to accommodate unique payloads while reducing wear and tear on the launch vehicle.
A modular payload carrier for use in a launch vehicle includes at least one module configured to fit in a payload bay of the vehicle and attachable to at
least one other module configured to fit in the bay. The module includes an outer wall contoured generally to fit a bottom surface contour of the bay. The
carrier can be used for the manifesting of both deployable and non-deployable payloads. Processing of payloads with the carrier can be performed, in large
part, separately from launch vehicle processing. Thus launch costs and turnaround times can be reduced.
Nanosat Launch Vehicle Technologies
Structures – Launch and Flight Vehicle
USPTO
Patent
Number
US6945498
US7036773
US7090171
US7281682
Title
Assignee
Abstract
Commercial
Experiment System in
Orbit
Kistler Aerospace
Corporation
An orbital experiment system with different internal experiment locations within a reusable launch vehicle making daylong delivery trips to space. The
experiments get access to the attributes of low earth orbit, the reusable launch vehicle's power and other subsystems. The experimenter utilizes uniform
experiment trays having a uniform connector for connection to an experiment management unit mounted on the orbital vehicle. The experiment
management unit provides power and data from the orbital vehicle related to the operation of the orbital vehicle and permits an experiment check
simulation prior to integration into the launch vehicle. The uniform size and connectivity requirements provide low cost options for the delivery of an
experiment into space and the return of the experiment from space. A commercial transportation system to and from orbit delivers a primary payload and
provides a 24-hour return cycle for the internal secondary experiments, which provide a quick confirmation of technical experiment exposure to space and
quick re-flight opportunities.
Compact External
Launcher for Small
Space Payloads
Momentum stabilized
launch vehicle upper
stage
Ecliptic Enterprises
Corporation
The present invention allows a payload to be launched from the exterior of a conventional launch vehicle with little integration expense. In one example,
the invention includes an aerodynamic fairing having an internal cavity to contain a payload, a mounting adapter to attach the fairing to an exterior surface
of a launch vehicle, and a release mechanism to separate the payload from the launch vehicle out of the fairing cavity during the launch vehicle's flight.
A spacecraft system including a spacecraft assembly or "stack" having an upper stage of a rocket-powered launch vehicle providing a final boost phase
during launch. The stack also includes a payload structure rotatably interconnected with the upper stage. The upper stage and the payload structure
together define a central axis that is generally coincident with the thrust axis during launch. The stack has an axis of maximum moment of inertia that is
not parallel to the central axis. The stack has internal damping such that unstable nutation occurs if the upper stage and the payload structure rotate
together about the central axis at the same rotational rate and in the same direction. The system includes a controller that rotates the payload structure
relative to the upper stage during the final boost phase to alleviate coning motion of the stack.
Spacecraft and launch
system
DBI/Century Fuels &
Aerospace Services
Honeywell
International, Inc.
A spacecraft system for achieving LEO and beyond includes a lifting body spacecraft, and an acceleration bed unit for accelerating the spacecraft
horizontally on a runway to liftoff. The rocket engines of the spacecraft then power the spacecraft into LEO, and the spacecraft glides back to earth, where
it is refurbished for reiterative use. A belly assembly of the spacecraft is removable and replaceable. The hydrogen fuel tanks of the spacecraft are modular
units, and each include a bladder that expands to fill the tank when empty to prevent explosion hazards.
Nanosat Launch Vehicle Technologies
Structures – Launch and Flight Vehicle
U.S. Patent Applications
Patent Number
Title
US2005211828A1
Aerodynamic orbit
inclination control
US2005230517A1
Assignee
AeroAstro, Inc.
Abstract
A method and system for deploying a spacecraft into a target orbit includes the use of controllable aerodynamic surfaces that can be deployed to facilitate a
change in the inclination angle of the trajectory of the spacecraft. In a typical embodiment, the spacecraft includes an orbit transfer vehicle containing the
aerodynamic structure, and a satellite that is to be placed into the target orbit. The spacecraft is launched to a higher-energy orbit than the target orbit, and the
energy released by traveling to the target orbit is used to change the inclination angle. After entering a transfer orbit that includes a passage through the upper
limits of the earth's atmosphere, the orbit transfer vehicle deploys the aerodynamic structure, and controls the aerodynamic surfaces of the structure to induce lift
forces that alter its inclination angle each time the vehicle enters the atmosphere.
Payload Delivery
Vehicle and Method
Teledyne
Solutions, Inc.
An un-manned multi-stage payload delivery vehicle and methods of deployment therefor are disclosed. According to various embodiments, the payload delivery
vehicle includes first stage and a second stage. At least one of the first and second stages comprises a jet engine. According to various embodiments, methods of
deploying the payload delivery vehicle include the steps of launching the payload delivery vehicle from a launch site, controlling the flight of the payload delivery
vehicle in accordance with one or more deployment parameters, deploying a payload attached to the payload delivery vehicle, and controlling the flight of the
payload delivery vehicle subsequent to payload deployment such that the payload delivery vehicle is flown to a pre-determined location for recovery and reuse.
US2002190160A1
Aerobraking Orbit
Transfer Vehicle
None
The excess space and weight capacity that is typical of a launch of large satellites to high-energy orbits, such as a geosynchronous orbit, is used to deploy small
satellites at a substantially lower-energy orbit, such as a low-earth orbit. An orbit-transfer vehicle provides the navigation, propulsion, and control systems
required to transport a payload satellite from a geosynchronous-transfer orbit (GTO) to a predetermined low-earth orbit (LEO). Depending upon the particular
configuration, upon achieving the low-earth orbit, the orbit transfer vehicle either releases the payload satellite, or remains attached to the payload satellite to
provide support services, such as power, communications, and navigation, to the payload satellite. To reduce the fuel requirements for this deployment via the
orbit-transfer vehicle, the orbit-transfer vehicle employs aerobraking to bring the satellite into a low-earth orbit. The aerobraking is preferably performed at a
nominal altitude of 150 km above the earth, where the atmosphere is dense enough to allow for a reasonably sized drogue device, yet rare enough to avoid the
need for special purpose heat-shielding materials. In a preferred operation, the provider of the orbit-transfer vehicle identifies and secures available excess
capacity on geosynchronous-transfer launch vehicles, and allocates the excess capacity to the satellites requiring low-earth orbit deployment, thereby providing a
deployment means that is virtually transparent to the purchaser of this deployment service.
US2003010868A1
Orbit transfer
vehicle with support
services
None
The excess space and weight capacity of a conventional launch vehicle for a high-energy orbit, such as GEO, is used to deploy satellites to a low-energy orbit,
such as LEO. In a preferred embodiment, an orbit-transfer vehicle provides the navigation, propulsion, and control systems required to transport a payload
satellite from a high-energy-transfer orbit, such as GTO, to a predetermined low-energy orbit. Upon entering the low-energy orbit, the payload satellite is released
from the orbit-transfer vehicle. To reduce the fuel requirements for this deployment via the orbit-transfer vehicle, a preferred embodiment includes aerobraking to
bring the satellite into a low-earth orbit. In a preferred embodiment of this method of deployment, the provider of the orbit-transfer vehicle identifies and secures
available excess capacity on launch vehicles, and allocates the excess capacity to the satellites requiring low-earth orbit deployment, thereby providing a
deployment means that is virtually transparent to the purchaser of this deployment service.
Nanosat Launch Vehicle Technologies
Structures – Launch and Flight Vehicle
U.S. Patent Applications
Patent Number
Title
US2003132350A1
Orbit Transfer
Vehicle with
Support Services
US2004031885A1
In Orbit Space
Transportation &
Recovery System
Assignee
None
Abstract
The excess space and weight capacity of a conventional launch vehicle for a high-energy orbit, such as GEO, is used to deploy satellites to a low-energy orbit,
such as LEO. In a preferred embodiment, an orbit-transfer vehicle provides the navigation, propulsion, and control systems required to transport a payload
satellite from a high-energy-transfer orbit, such as GTO, to a predetermined low-energy orbit. Upon entering the low-energy orbit, the payload satellite is released
from the orbit-transfer vehicle. To reduce the fuel requirements for this deployment via the orbit-transfer vehicle, a preferred embodiment includes aerobraking to
bring the satellite into a low-earth orbit. In a preferred embodiment of this method of deployment, the provider of the orbit-transfer vehicle identifies and secures
available excess capacity on launch vehicles, and allocates the excess capacity to the satellites requiring low-earth orbit deployment, thereby providing a
deployment means that is virtually transparent to the purchaser of this deployment service.
None
An In Orbit Transportation & Recovery System (IOSTAR(TM)) (10) is disclosed. One preferred embodiment of the present invention comprises a space tug
powered by a nuclear reactor (19). The IOSTAR(TM) includes a collapsible boom (11) connected at one end to a propellant tank (13) which stores fuel for an
electric propulsion system (12). This end of the boom (11) is equipped with docking hardware (14) that is able to grasp and hold a satellite (15) and as a means to
refill the tank (13). Radiator panels (16) mounted on the boom (11) dissipate heat from the reactor (19). A radiation shield (20) is situated next to the reactor (19)
to protect the satellite payload (15) at the far end of the boom (11). The IOSTAR(TM) (10) will be capable of accomplishing rendezvous and docking maneuvers
which will enable it to move spacecraft between a low Earth parking orbit and positions in higher orbits or to other locations in our Solar System.
US2004135035A1
Momentum
None
Stabilized Launch
Vehicle Upper Stage
A spacecraft system including a spacecraft assembly or "stack" having an upper stage of a rocket-powered launch vehicle providing a final boost phase during
launch. The stack also includes a payload structure rotatably interconnected with the upper stage. The upper stage and the payload structure together define a
central axis that is generally coincident with the thrust axis during launch. The stack has an axis of maximum moment of inertia that is not parallel to the central
axis. The stack has internal damping such that unstable nutation occurs if the upper stage and the payload structure rotate together about the central axis at the
same rotational rate and in the same direction. The system includes a controller that rotates the payload structure relative to the upper stage during the final boost
phase to alleviate coning motion of the stack.
US2004217231A1
Spacecraft and
launch system
None
A spacecraft system for achieving LEO and beyond includes a lifting body spacecraft, and an acceleration bed unit for accelerating the spacecraft horizontally on
a runway to liftoff. The rocket engines of the spacecraft then power the spacecraft into LEO, and the spacecraft glides back to earth, where it is refurbished for
reiterative use. A belly assembly of the spacecraft is removable and replaceable. The hydrogen fuel tanks of the spacecraft are modular units, and each include a
bladder that expands to fill the tank when empty to prevent explosion hazards.
US2004245407A1
In Orbit Space
Transportation &
Recovery System
None
An In Orbit Transportation \A Recovery System (IOSTAR(TM)) (10) is disclosed. One preferred embodiment of the present invention comprises a space tug
powered by a nuclear reactor (19). The IOSTAR(TM) includes a collapsible boom (11) connected at one end to a propellant tank (13) which stores fuel for an
electric propulsion system (12). This end of the boom (11) is equipped with docking hardware (14) that is able to grasp and hold a satellite (15) and as a means to
refill the tank (13). Radiator panels (16) mounted on the boom (11) dissipate heat from the reactor (19). A radiation shield (20) is situated next to the reactor (19)
to protect the satellite payload (15) at the far end of the boom (11). The IOSTAR(TM) (10) will be capable of accomplishing rendezvous and docking maneuvers
which will enable it to move spacecraft between a low Earth parking orbit and positions in higher orbits or to other locations in our Solar System.
Nanosat Launch Vehicle Technologies
Structures – Launch and Flight Vehicle
U.S. Patent Applications
Patent Number
Title
US2007108349A1
In Orbit Space
Transportation and
Recovery System
US2007120020A1
Small Reusable
Payload Delivery
Vehicle
Assignee
None
Abstract
An In Orbit Transportation & Recovery System (IOSTAR(TM)) ( 10 ) is disclosed. One preferred embodiment of the present invention comprises a space
tug powered by a nuclear reactor ( 19 ). The IOSTAR(TM) includes a collapsible boom ( 11 ) connected at one end to a propellant tank ( 13 ) which stores fuel for
an electric propulsion system ( 12 ). This end of the boom ( 11 ) is equipped with docking hardware ( 14 ) that is able to grasp and hold a satellite ( 15 ) and as a
means to refill the tank ( 13 ). Radiator panels ( 16 ) mounted on the boom ( 11 ) dissipate heat from the reactor ( 19 ). A radiation shield ( 20 ) is situated next to
the reactor ( 19 ) to protect the satellite payload ( 15 ) at the far end of the boom ( 11 ). The IOSTAR(TM) ( 10 ) will be capable of accomplishing rendezvous and
docking maneuvers which will enable it to move spacecraft between a low Earth parking orbit and positions in higher orbits or to other locations in our Solar
System.
Mei
Technologies,
Inc.
The present invention provides a small unmanned payload delivery vehicle that can deploy one or more payloads into space and then bring the payloads back to
earth. The delivery vehicle can be sent into space by an expendable launch vehicle, a space shuttle, or be launched from the space station. The delivery vehicle
together with the payload contained therein can be left in space for a variable period of time, and the attitude of the delivery vehicle can be adjusted from time to
time to maintain the vehicle in the desired orbit. When it is time to return the payload to earth, the delivery vehicle is de-orbited and re-enters the earth's
atmosphere. The descent of the delivery vehicle is controlled by a parachute system packed within the vehicle. The delivery vehicle together with the payload
contained therein can finally be retrieved based on signals emitted from a beacon.
Nanosat Launch Vehicle Technologies
Structures – Launch and Flight Vehicle
WIPO
Patent Number
Title
WO04085252A2
Unitized Hybrid
Rocket System
WO05015116A2
Assignee
Mojave Aerospace
Ventures, Llc
Abstract
The hybrid rocket system of this invention is characterized by use of an oxidizer tank having a cylindrical midsection surrounded by a skirt and
bonded thereto by a layer of elastomeric adhesive. The skirt outer surface is in turn adhesively secured to a spacecraft inner surface. An elongated
solid-fuel motor case is mechanically rigidly secured to a central rear surface of the tank, and the case terminates in a throat and nozzle. The
elastomeric-adhesive bonding of tank to skirt, and rigid adhesion of skirt to spacecraft forms the sole support for the rocket system, and separate
support for the motor case is not required.
Air-Based Vertical
Launch Ballistic
Missile Defense
Bae Systems Applied
Technologies, Inc.
An air-based vertical launch system is described by means of which ballistic missile defense can be achieved effectively from a large aircraft. A
method for ensuring safe missile egress is proposed. A method for ensuring that the missile strikes the ballistic missile payload section is also
proposed. Together, the air basing method employing vertical (or near-vertical) launch and semi-active laser guidance yield an affordable and
operationally effective missile defense against both tactical and long-range ballistic missiles. The affordability of missile defense is enhanced by the
ability of an aircraft equipped with a vertical launcher to simultaneously carry out several defensive and offensive missions and to provide other
capabilities such as satellite launch at other times. Methods for employing an aircraft equipped with a vertical (or near-vertical launcher) and one or
more of the proposed egress assurance mechanisms in offensive ground attack missions, mine laying, and satellite launch missions are also proposed.
WO06119056A2
Lighter Than Air
Supersonic Vehicle
General Orbital
Corporation
The present invention comprises a lighter than air space vehicle which is optimized for supersonic speeds and altitudes from 30,000 feet to space.
WO07126526A1
TWO Part Spacecraft
Servicing Vehicle
System with
Universal Docking
Adaptor
The Boeing Company
An in-space spacecraft servicing system (10) includes a servicing spacecraft (22) and a propellant module (24). The servicing spacecraft includes a
client servicing system (136), as well as navigation avionics (108) for independent flight operation and a servicing propellant tank (170). The
propellant module moves the servicing spacecraft from an upper stage drop off location and releases it in proximity to a client spacecraft (16) for a
servicing mission. It has a propellant tank (172) with capacity for multiple missions and is used to refill the servicing spacecraft's propellant tanks
between missions. Either or both the servicing spacecraft and the propellant module may have navigation avionics. The servicing spacecraft also has
a universal docking adaptor (70) for different client spacecraft, and can convert a client spacecraft from non-cooperative to cooperative.
Nanosat Launch Vehicle Technologies
Structures – Launch and Flight Vehicle
FC
GSFC
Title
A NASA Stratem for Leveranin2 Emerainn
Launch Vehicles for Routine, Small Pavload
Missions
GRC
Affordable Flight Demonstration of the GTX
Air-Breathing SSTO Vehicle Concept
GRC
High Altitude Launch for a Practical SSTO
NTRS
Abstract
Orbital flight opportunities for small payloads have always been few and far between, and then on February 1, 2002, the situation got worse. In the wake of the loss of the Columbia
during STS- 107, changing NASA missions and priorities led to the termination of the Shuttle Small Payloads Projects, including Get-Away Special, Hitcbker, and Space Experiment
Module. In spite of the limited opportunities, long queue, and restrictions associated with flying experiments on a man-rated transportation system the carriers provided a sustained,
high quality experiment services for education, science, and technology payloads, and was one of the few games in town. Attempts to establish routine opportunities aboard existing
ELVs have been unsuccessful, as the cost-per-pound on small ELVs and conflicts with primary spacecraft on larger vehicles have proven prohibitive. Ths has led to a backlog of
existing NASA-sponsored payloads and no prospects or plans for fbture opportunities within the NASA community. The prospects for breaking out of this paradigm appear promising
as a result of NASA s partnership with DARPA in pursuit of low-cost, responsive small ELVs under the Falcon Program. Through this partnership several new small ELVs, providing
1000 lbs. to LEO will be demonstrated in less than two years that promise costs that are reasonable enough that NASA, DoD, and other sponsors can once again invest in small
payload opportunities. Within NASA, planning has already begun. NASA will be populating one or more of the Falcon demonstration flights with small payloads that are already under
development. To accommodate these experiments, Goddard s Wallops Flight Facility has been tasked to develop a multi-payload ejector (MPE) to accommodate the needs of these
payloads. The MPE capabilities and design is described in detail in a separately submitted abstract. Beyond use of the demonstration flights however, Goddard has already begun
developing strategies to leverage these new ELVs as elements of a larger system designed to provide routine, low-cost end-to-end services for small science, Exploration, and
education payloads. The plan leverages the management approaches of the successful Sounding Rocket Program and Shuttle Small Payloads Projects. The strategy consists of
using a systems implementation approach of elements, including 1) Falcon ELVs, 2) advanced launch site technologies and processes, 3) suite of experiment carriers accommodating
different mission requirements, 4) streamlined integration and test operations, 5 ) experiment brokering and management, and 6) standardized, distributed payload operations. The
envisioned suite of carriers includes the MPE, a standard interface experiment carrier, and potentially a reentry fieeflyer experiment carrier. Key to the success of this strategy is
standard experiment interfaces within the carriers to limit mission- unique tasks, establishmg and managing a program of scheduled reoccurring flights rather than discrete missions,
and streamlined, centralized implementation of the elements. These individual elements are each under development and Goddard will demonstrate the overall system strategy lowcost small payload missions on the initial Falcon demonstration launches from Wallops. goal is to show that this model should be converted to a sustained NASA program supporting
science, technology, and education, with annual flight opportunities. The paper will define in detail the various elements of the overall program, as well as provide status, philosophy,
and strategy for the program that will hopefully once-and-for-all provide low-cost, routine access to space for the small payloads community.
The rocket based combined cycle (RBCC) powered single-stage-to-orbit (SSTO) reusable launch vehicle has the potential to significantly reduce the total cost per pound for orbital
payload missions. To validate overall system performance, a flight demonstration must be performed. This paper presents an overview of the first phase of a flight demonstration
program for the GTX SSTO vehicle concept. Phase 1 will validate the propulsion performance of the vehicle configuration over the supersonic and hypersonic airbreathing portions of
the trajectory. The focus and goal of Phase 1 is to demonstrate the integration and performance of the propulsion system flowpath with the vehicle aerodynamics over the airbreathing trajectory. This demonstrator vehicle will have dual mode ramjet scramjets, which include the inlet, combustor, and nozzle with geometrically scaled aerodynamic surface
outer mold lines (OML) defining the forebody, boundary layer diverter, wings, and tail. The primary objective of this study is to demonstrate propulsion system performance and
operability including the ram to scram transition, as well as to validate vehicle aerodynamics and propulsion airframe integration. To minimize overall risk and development cost the
effort will incorporate proven materials, use existing turbomachinery in the propellant delivery systems, launch from an existing unmanned remote launch facility, and use basic vehicle
recovery techniques to minimize control and landing requirements. A second phase would demonstrate propulsion performance across all critical portions of a space launch trajectory
(lift off through transition to all-rocket) integrated with flight-like vehicle systems.
Existing engineering materials allow the construction of towers to heights of many kilometers. Orbital launch from a high altitude has significant advantages over sea-level launch due
to the reduced atmospheric pressure, resulting in lower atmospheric drag on the vehicle and allowing higher rocket engine performance. High-altitude launch sites are particularly
advantageous for single-stage to orbit (SSTO) vehicles, where the payload is typically 2 of the initial launch mass. An earlier paper enumerated some of the advantages of high
altitude launch of SSTO vehicles. In this paper, we calculate launch trajectories for a candidate SSTO vehicle, and calculate the advantage of launch at launch altitudes 5 to 25
kilometer altitudes above sea level. The performance increase can be directly translated into increased payload capability to orbit, ranging from 5 to 20 increase in the mass to orbit.
For a candidate vehicle with an initial payload fraction of 2 of gross lift-off weight, this corresponds to 31 increase in payload (for 5-km launch altitude) to 122 additional payload (for
25-km launch altitude).
Nanosat Launch Vehicle Technologies
Structures – Launch and Flight Vehicle
NTRS
Abstract
The Defense Advanced Research Projects Agency has proposed a two-stage system to deliver a small payload to orbit. The proposal calls for an airplane to perform an exoatmospheric zoom climb maneuver, from which a second-stage rocket is launched carrying the payload into orbit. The NASA Dryden Flight Research Center has conducted an inhouse generic simulation study to determine how accurately a human-piloted airplane can deliver a second-stage rocket to a desired exo-atmospheric launch condition. A highperformance, fighter-type, fixed-base, real-time, pilot-in-the-loop airplane simulation has been modified to perform exo-atmospheric zoom climb maneuvers. Four research pilots
tracked a reference trajectory in the presence of winds, initial offsets, and degraded engine thrust to a second-stage launch condition. These launch conditions have been compared to
the reference launch condition to characterize the expected deviation. At each launch condition, a speed change was applied to the second-stage rocket to insert the payload onto a
transfer orbit to the desired operational orbit. The most sensitive of the test cases was the degraded thrust case, yielding second-stage launch energies that were too low to achieve
the radius of the desired operational orbit. The handling qualities of the airplane, as a first-stage vehicle, have also been investigated.
FC
DFRC
Title
Launch Condition Deviations of Reusable
Launch Vehicle Simulations in ExoAtmospheric Zoom Climbs
MSFC
Small Satellites and the DARPA Air Force
Falcon Program
GRC
SRM-Assisted Trajectory for the GTX
Reference Vehicle
MSFC
The DARPA USAF Falcon Program Small
Launch Vehicles
Earlier in this decade, the U.S. Air Force Space Command and the Defense Advanced Research Projects Agency (DARPA), in recognizing the need for low-cost responsive small
launch vehicles, decided to partner in addressing this national shortcoming. Later, the National Aeronautics and Space Administration (NASA) joined in supporting this effort, dubbed
the Falcon Program. The objectives of the Small Launch Vehicle (SLV) element of the DARPA USAF Falcon Program include the development of a low-cost small launch vehicle(s)
that demonstrates responsive launch and has the potential for achieving a per mission cost of less than 5M when based on 20 launches per year for 10 years. This vehicle class can
lift 1000 to 2000 lbm payloads to a reference low earth orbit. Responsive operations include launching the rocket within 48 hours of call up. A history of the program and the current
status will be discussed with an emphasis on the potential impact on small satellites.
GSFC
Multiple Payload Ejector for Education,
Science and Technology Experiments
The education research community no longer has a means of being manifested on Space Shuttle flights, and small orbital payload carriers must be flown as secondary payloads on
ELV flights, as their launch schedule, secondary payload volume and mass permits. This has resulted in a backlog of small payloads, schedule and cost problems, and an inability for
the small payloads community to achieve routine, low-cost access to orbit. This paper will discuss Goddard's Wallops Flight Facility funded effort to leverage its core competencies in
small payloads, sounding rockets, balloons and range services to develop a low cost, multiple payload ejector (MPE) carrier for orbital experiments. The goal of the MPE is to provide
a low-cost carrier intended primarily for educational flight research experiments. MPE can also be used by academia and industry for science, technology development and
Exploration experiments. The MPE carrier will take advantage of the DARPAI NASA partnership to perform flight testing of DARPA s Falcon small, demonstration launch vehicle. The
Falcon is similar to MPE fiom the standpoint of focusing on a low-cost, responsive system. Therefore, MPE and Falcon complement each other for the desired long-term goal of
providing the small payloads community with a low-cost ride to orbit. The readiness dates of Falcon and MPE are complementary, also. MPE is being developed and readied for flight
within 18 months by a small design team. Currently, MPE is preparing for Critical Design Review in fall 2005, payloads are being manifested on the first mission, and the carrier will be
ready for flight on the first Falcon demonstration flight in summer, 2006. The MPE and attached experiments can weigh up to 900 lb. to be compatible with Falcon demonstration
vehicle lift capabilities fiom Wallops, and will be delivered to the Falcon demonstration orbit - 100 nautical mile circular altitude.
The FALCON ((Force Application and Launch from CONUS) program is a technology demonstration effort with three major components a Small Launch Vehicle (SLV), a Common
Aero Vehicle (CAV), and a Hypersonic Cruise Vehicle (HCV). Sponsored by DARPA and executed jointly by the United States Air Force and DARPA with NASA participation, the
objectives are to develop and demonstrate technologies that will enable both near-term and far-term capability to execute time-critical, global reach missions. The focus of this paper is
on the SLV as it relates to small satellites and the implications of lower cost to orbit for small satellites. The target recurring cost for placing 1000 pounds payloads into a circular
reference orbit of 28.5 degrees at 100 nautical miles is 5,000,000 per launch. This includes range costs but not the payload or payload integration costs. In addition to the nominal
1000 pounds to LEO, FALCON is seeking delivery of a range of orbital payloads from 220 pounds to 2200 pounds to the reference orbit. Once placed on alert status, the SLV must be
capable of launch within 24 hours.
A goal of the GTX effort has been to demonstrate the feasibility of a single stage-to-orbit (SSTO) vehicle that delivers a small payload to low earth orbit. The small payload class was
chosen in order to minimize the risk and cost of development of this revolutionary system. A preliminary design study by the GTX team has resulted in the current configuration that
offers considerable promise for meeting the stated goal. The size and gross lift-off weight resulting from scaling the current design to closure however may be considered impractical
for the small payload. In lieu of evolving the project' reference vehicle to a large-payload class, this paper offers the alternative of using solid-rocket motors in order to close the vehicle
at a practical scale. This approach offers a near-term, quasi-reusable system that easily evolves to reusable SSTO following subsequent development and optimization. This paper
presents an overview of the impact of the addition of SRM's to the GTX reference vehicle#s performance and trajectory. The overall methods of vehicle modeling and trajectory
optimization will also be presented. A key element in the trajectory optimization is the use of the program OTIS 3.10 that provides rapid convergence and a great deal of flexibility to
the user. This paper will also present the methods used to implement GTX requirements into OTIS modeling.
Nanosat Launch Vehicle Technologies
Structures – Modular Interconnects
USPTO
Patent Number
Title
US7093805
System and Methods for
Integrating a Payload with a
Launch Vehicle
Assignee
The Boeing Company
Abstract
A system for providing an interface between a launch vehicle and a payload of the launch vehicle includes payload integration points of the
launch vehicle that remain unchanged relative to different payload configurations, such that a payload is configurable apart from the launch
vehicle. The system makes it possible to de-couple payload installation and launch vehicle turnaround operations. Payloads can be
configured off-line from launch vehicle processing, thus making it possible to accommodate unique payloads while reducing wear and tear
on the launch vehicle.
Nanosat Launch Vehicle Technologies
Thermal – Control Instrumentation
FC
JPL
Title
Development of MEMS microchannel heat sinks for
micro nano spacecraft thermal control
NTRS
Abstract
MEMS-based microchannel heat sinks are being investigated at the Jet Propulsion Laboratory for use in micro nano spacecraft thermal control.
Nanosat Launch Vehicle Technologies
Thermal – Control Instrumentation
Company Name
Allcomp, Inc.
Title
AN ULTRA-LIGHTWEIGHT, HIGH PERFORMANCE
CARBON-CARBON SPACE RADIATOR
NASA SBIR Phase II
Field Center
Quad Chart
GRC
C-C composite material is an ideal candidate to solve the challenge in new generation space radiators because
of its very low density, high thermal conductivity, good mechanical properties, and tailor-ability of its thermal and
structural properties.
Nanosat Launch Vehicle Technologies
Thermal – Control Instrumentation
FC
GSFC
Title
Development of the variable emittance
thermal suite for the space technology 5
microsatellite
GSFC
Electrochromic Variable Emittance
Devices on Silicon Wafer for Spacecraft
Thermal Control
NTRS
Abstract
The advent of very small satellites, such as nano and microsatellites, logically leads to a requirement for smaller thermal control subsystems. In addition, the thermal control needs of the
smaller spacecraft instrument may well be different from more traditional situations. For example, power for traditional heaters may be very limited or unavailable, mass allocations may
be severely limited, and fleets of nano microsatellites will require a generic thermal design as the cost of unique designs will be prohibitive. Some applications may require significantly
increased power levels while others may require extremely low heat loss for extended periods. Small spacecraft will have low thermal capacitance thus subjecting them to large
temperature swings when either the heat generation rate changes or the thermal sink temperature changes. This situation, combined with the need for tighter temperature control, will
present a challenging situation during transient operation. The use of 'off-the-shelf' commercial spacecraft buses for science instruments will also present challenges. Older thermal
technology, such as heaters, thermostats, and heat pipes, will almost certainly not be sufficient to meet the requirements of these new spacecraft instruments. They are generally too
heavy, not scalable to very small sizes, and may consume inordinate amounts of power. Hence there is a strong driver to develop new technology to meet these emerging needs.
Variable emittance coatings offer an exciting alternative to traditional control methodologies and are one of the technologies that will be flown on Space Technology 5, a mission of three
microsatellites designed to validate 'enabling' technologies. Several studies have identified variable emittance coatings as applicable to a wide range of spacecraft, and to potentially
offer substantial savings in mass and or power over traditional approaches. This paper discusses the development of the variable emittance thermal suite for ST-5. More specifically, it
provides a description of and the infusion and validation plans for the variable emittance coatings.
Small light-weight satellites and space vehicles under development for future NASA missions have reduced thermal mass and are strongly affected by changes in orbital conditions,
resulting in large temperature variations. Restrictions on payload weight and volume limit the usefulness of many thermal control technologies. One thermal control approach, being
considered by NASA in both nano- and micro- spacecraft applications, involves the use of electrochromic (EC) variable emittance devices (VEDs). VEDs operating in the harsh space
environment (UV radiation, atomic oxygen) must be properly protected if they are to reach their design operational life. In this paper, we discuss the design of an all-solid-state EC VED
built on a silicon wafer. The silicon wafer serves as a window for IR radiation and protects EC layers from the space environment. This paper also discusses the expected limits of
emittance modulation of the EC VED on the silicon substrate as well as possible impact of an antireflective coating on IR emittance modulation. copyright 2004 American Institute of
Physics
Nanosat Launch Vehicle Technologies
Thermal – Thermal Insulating Materials
Company Name
Aspen Aerogels
Title
AEROGEL INSULATION FOR
INTEGRATED CRYOTANKS AND TPS
Field Center
JSC
Quest Product Development
Corporation
INTEGRATED MLI: ADVANCED
THERMAL INSULATION USING
MICROMOLDING TECHNOLOGY
GRC
NASA SBIR Phase II
Quad Chart
The proposed Phase II project will develop aerogel insulation materials for composite cryotanks and thermal protection systems (TPS).
Introduction of aerogel materials to reusable launch vehicles will result in significant reductions in the weight and volume of cryogenic
insulation and high temperature TPS. Aspen Aerogels’ materials typically demonstrate 2-4x improvement in thermal conductivity over
traditional insulation materials.
Integrated Multi-Layer Insulation (IMLI) is an innovative new technology using a micro-molded polymer substructure integrated with radiation barriers
to provide an ultra-high performance thermal insulation system. Tasks include next generation design, material selection, fabrication methods for
seams and corners including interleaving and layer thermal matching, and building and testing prototypes in realistic environments.
Nanosat Launch Vehicle Technologies
Verification and Validation – Operations Concepts and Requirements
FC
DFRC
Title
Ground Support for the Space-Based Range Flight
Demonstration 2
DFRC/KSC
Space-Based Range Safety and Future Space
Range Applications
NTRS
Abstract
The primary objective of the NASA Space-Based Range Demonstration and Certification program was to develop and demonstrate space-based range capabilities.
The Flight Demonstration 2 flights at NASA Dryden Flight Research Center were conducted to support Range Safety (commanding and position reporting) and
high-rate (5 Mbps) Range User (video and data) requirements. Required ground support infrastructure included a flight termination system computer, the grounddata distribution network to send range safety commands and receive range safety and range user telemetry data and video, and the ground processing systems at
the Dryden Mission Control Center to process range safety and range user telemetry data and video.
The National Aeronautics and Space Administration (NASA) Space-Based Telemetry and Range Safety (STARS) study is a multiphase project to demonstrate the
performance, flexibility and cost savings that can be realized by using space-based assets for the Range Safety global positioning system (GPS) metric tracking
data, flight termination command and range safety data relay and Range User (telemetry) functions during vehicle launches and landings. Phase 1 included flight
testing S-band Range Safety and Range User hardware in 2003 onboard a high-dynamic aircraft platform at Dryden Flight Research Center (Edwards, California,
USA) using the NASA Tracking and Data Relay Satellite System (TDRSS) as the communications link. The current effort, Phase 2, includes hardware and
packaging upgrades to the S-band Range Safety system and development of a high data rate Ku-band Range User system. The enhanced Phase 2 Range Safety
Unit (RSU) provided real-time video for three days during the historic Global Flyer (Scaled Composites, Mojave, California, USA) flight in March, 2005. Additional
Phase 2 testing will include a sounding rocket test of the Range Safety system and aircraft flight testing of both systems. Future testing will include a flight test on a
launch vehicle platform. This paper discusses both Range Safety and Range User developments and testing with emphasis on the Range Safety system. The
operational concept of a future space-based range is also discussed.
Nanosat Launch Vehicle Technologies
Verification and Validation – Simulation Modeling Environment
FC
ARC
Title
Modeling and Simulation of Shuttle Launch and
Range Operations
NTRS
Abstract
The simulation and modeling test bed is based on a mockup of a space flight operations control suitable to experiment physical, procedural, software, hardware
and psychological aspects of space flight operations. The test bed consists of a weather expert system to advise on the effect of weather to the launch
operations. It also simulates toxic gas dispersion model, impact of human health risk, debris dispersion model in 3D visualization. Since all modeling and
simulation is based on the internet, it could reduce the cost of operations of launch and range safety by conducting extensive research before a particular launch.
Each model has an independent decision making module to derive the best decision for launch.
Nanosat Launch Vehicle Technologies
Verification and Validation – Testing Facilities
FC
JSC
Title
A Facility for Testing High-Power
Electric Propulsion Systems in
Space A Design Study
GRC/JSC/MSFC
A High-power Electric Propulsion
Test Platform in Space
NTRS
Abstract
This paper will describe the results of the preliminary phase of a NASA design study for a facility to test high-power electric propulsion systems in space. The results of this design study
are intended to provide a firm foundation for a subsequent detailed design and development activities leading to the deployment of a valuable space facility supporting the new vision of
space exploration. The objectives for human and robotic exploration of space can be accomplished affordably, safely and effectively with high-power electric propulsion systems. But, as
thruster power levels rise to the hundreds of kilowatts and up to megawatts, their testing will pose stringent and expensive demands on existing Earth-based vacuum facilities. These
considerations and the access to near-Earth space provided by the International Space Station (ISS) have led to a renewed interest in space testing. The ISS could provide an excellent
platform for a space-based test facility with the continuous vacuum conditions of the natural space environment and no chamber walls to modify the open boundary conditions of the
propulsion system exhaust. The platform would be designed to accommodate the side-by-side testing of multiple types of electric thrusters currently under development and thus provide a
strong basis for comparing their relative performance. The utility of testing on the station is further enhanced by the human presence, enabling close interaction with and modification of
the test hardware in a true laboratory environment. These conditions facilitate rapid development and flight certification at potentially lower cost than with conventional Earth-bound
facilities. As an added benefit, the propulsive effect of these tests could provide some drag compensation for the station, reducing the re-boost cost for the orbital facility. While it is
expected that the ISS will not be capable of generating continuous levels of high power, the utilization of state-of-the-art energy storage media would be sufficient to achieve very high
power levels over intervals short enough to be feasible and long enough to provide ample demonstration of steady-state operation. This paper will outline the results of the preliminary
phase of the design study with emphasis on the requirements that will dictate the system design.
This paper will describe the results of the preliminary phase of a NASA design study for a facility to test high-power electric propulsion systems in space. The results of this design study
are intended to provide a firm foundation for subsequent detailed design and development activities leading to the deployment of a valuable space facility. The NASA Exploration Systems
Mission Directorate is sponsoring this design project. A team from the NASA Johnson Space Center, Glenn Research Center, the Marshall Space Flight Center and the International
Space Station Program Office is conducting the project. The test facility is intended for a broad range of users including government, industry and universities. International participation is
encouraged. The objectives for human and robotic exploration of space can be accomplished affordably, safely and effectively with high-power electric propulsion systems. But, as thruster
power levels rise to the hundreds of kilowatts and up to megawatts, their testing will pose stringent and expensive demands on existing Earth-based vacuum facilities. These
considerations and the human access to near-Earth space provided by the International Space Station (ISS) have led to a renewed interest in space testing. The ISS could provide an
excellent platform for a space-based test facility with the continuous vacuum conditions of the natural space environment and no chamber walls to modify the open boundary conditions of
the propulsion system exhaust. The test platform could take advantage of the continuous vacuum conditions of the natural space environment. Space testing would provide open
boundary conditions without walls, micro-gravity and a realistic thermal environment. Testing on the ISS would allow for direct observation of the test unit, exhaust plume and spaceplasma interactions. When necessary, intervention by on-board personnel and post-test inspection would be possible. The ISS can provide electrical power, a location for diagnostic
instruments, data handling and thermal control. The platform will be designed to accommodate the side-by-side testing of multiple types of electric thrusters. It is intended to be a
permanent facility in which different thrusters can be tested over time. ISS crews can provide maintenance for the platform and change out thruster test units as needed. The primary
objective of this platform is to provide a test facility for electric propulsion devices of interest for future exploration missions. These thrusters are expected to operate in the range of
hundreds of kilowatts and above. However, a platform with this capability could also accommodate testing of thrusters that require much lower power levels. Testing at the higher power
levels would be accomplished by using power fiom storage devices on the platform, which would be gradually recharged by the ISS power generation system. This paper will summarize
the results of the preliminary phase of the study with an explanation of the user requirements and the initial conceptual design. The concept for test operations will also be described. The
NASA project team is defining the requirements but they will also reflect the inputs of the broader electric propulsion community including those at universities, commercial enterprises and
other government laboratories. As a facility on the International Space Station, the design requirements are also intended to encompass the needs of international users. Testing of
electric propulsion systems on the space station will help advance the development of systems needed for exploration and could also serve the needs of other customers. Propulsion
systems being developed for commercial and military applications could be tested and certification testing of mature thrusters could be accomplished in the space environment.
Secondary and Tertiary Launch Technologies
Avionics and Astionics – Telemetry, Tracking, and Control
U.S. Patent Applications
Patent Number
Title
US2007235592A1
Minimum time or thrust
separation trajectory for
spacecraft emergency separation
Assignee
None
Abstract
A method and apparatus for separating a first spacecraft from a rotating second spacecraft within a separation corridor having a
trailing boundary surface is disclosed. The method and apparatus use a non-stationkeeping trajectory that uses a greater portion of
the separation corridor to reduce thrust requirements and reduce the time necessary for separation.
Secondary and Tertiary Launch Technologies
Electronics – Highly-Recongifurable
WIPO
Patent Number
Title
WO06093715A2
Spacecraft Adapter Having
Embedded Resources, and
Methods of FormingSame
Assignee
The Boeing Company
Abstract
A spacecraft adapter having embedded resources for supporting a non-primary payload on a launch vehicle. The spacecraft adapter includes a
battery, a power distribution and control system, and an interface circuit for interfacing with the non- primary payload. Other
modules/subsystems such as data storage, sensor and data interface and communications may be included to suit the needs of a particular nonprimary payload and/or particular mission of the non-primary payload. The adaptor does not require any interfacing with the bus of the
primary payload and can be scaled/modified as needed to provide only that degree of functionality needed for a given non-primary payload
being carried by the launch vehicle.
Secondary and Tertiary Launch Technologies
Power and Energy – Energy Storage
USPTO
Patent Number
Title
US7216834
Orbit Space Transportation and
Recovery System
Assignee
Iostar Corporation
Abstract
An In Orbit Transportation \A Recovery System (IOSTAR(TM)) ( 10 ). One preferred embodiment of the present invention comprises a
space tug powered by a nuclear reactor ( 19 ). The IOSTAR(TM) includes a collapsible boom ( 11 ) connected at one end to a propellant
tank ( 13 ) which stores fuel for an electric propulsion system ( 12 ). This end of the boom ( 11 ) is equipped with docking hardware ( 14
) that is able to grasp and hold a satellite ( 15 ) and as a means to refill the tank ( 13 ). Radiator panels ( 16 ) mounted on the boom ( 11 )
dissipate heat from the reactor ( 19 ). A radiation shield ( 20 ) is situated next to the reactor ( 19 ) to protect the satellite payload ( 15 ) at
the far end of the boom ( 11 ). The IOSTAR(TM) ( 10 ) will be capable of accomplishing rendezvous and docking maneuvers which will
enable it to move spacecraft between a low Earth parking orbit and positions in higher orbits or to other locations in our Solar System.
Secondary and Tertiary Launch Technologies
Propulsion --Tethers
FC
MSFC
Title
Propulsive Small Expendable Deployer System
(ProSEDS)
MSFC
Review of the ProSEDS Electrodynamic Tether
Mission Development
NASA (non
Center
Specific)
The muTORQUE momentum-exchange tether
experiment
NTRS
Abstract
The Propulsive Small Expendable Deployer System (ProSEDS) space experiment will demonstrate the use of an electrodynamic tether propulsion system to generate thrust in
space by decreasing the orbital altitude of a Delta 11 Expendable Launch Vehicle second stage. ProSEDS, which is planned on an Air Force GPS Satellite replacement mission
in June 2002, will use the flight proven Small Expendable Deployer System (SEDS) to deploy a tether (5 km bare wire plus 10 km non-conducting Dyneema) from a Delta 11
second stage to achieve approx. 0.4N drag thrust. ProSEDS will utilize the tether-generated current to provide limited spacecraft power. The ProSEDS instrumentation includes
Langmuir probes and Differential Ion Flux Probes, which will determine the characteristics of the ambient ionospheric plasma. Two Global Positioning System (GPS) receivers
will be used (one on the Delta and one on the endmass) to help determine tether dynamics and to limit transmitter operations to occasions when the spacecraft is over selected
ground stations. The flight experiment is a precursor to the more ambitious electrodynamic tether upper stage demonstration mission, which will be capable of orbit raising,
lowering and inclination changes-all using electrodynamic thrust. An immediate application of ProSEDS technology is for the removal of spent satellites for orbital debris
mitigation. In addition to the use of this technology to provide orbit transfer and debris mitigation it may also be an attractive option for future missions to Jupiter and any other
planetary body with a magnetosphere.
The Propulsive Small Expendable Deployer System (ProSEDS) space experiment was ready to fly as a secondary payload on a Delta-II expendable launch vehicle in late
March 2003. Concerns raised in February 2003 by the International Space Station resulted in the delay of the launch of ProSEDS. Issues associated with the delayed launch
date and a change in starting altitude resulted in the cancellation of the mission. ProSEDS was intended to deploy a tether (5 km bare wire plus 10 km non-conducting
Dyneema) from a Delta I1 second stage to achieve adequate drag thrust that would lower the orbit of the system over days as opposed to months due to atmospheric drag. It
was also designed to utilize the tether-generated current to provide limited spacecraft power. Considerable effort and testing went in to developing the ProSEDS system by a
dedicated team. Through this effort, important technological issues were identified and addressed and this presentation will discuss some of the important technical issues and
hurdles that had to be addressed to successfully prepare for flight. It is intended that this information will be of use for future tether mission and experiment designers.
Long, high-strength tethers can provide a mechanism for transferring orbital momentum and energy from one space object to another without the consumption of propellant. By
providing a highly-reusable transportation architecture, systems built upon such "momentum-exchange" tethers may be able to achieve significant cost reductions for a number
of in-space propulsion missions. Before such systems could be placed into operation, however, a number of technical challenges must be met, including flight demonstration of
high-strength, highly survivable tethers, demonstration of the ability to control the dynamics of a rotating tether system, and the ability for a tether system to rendezvous with,
capture, and then toss a payload. In this paper, we discuss a concept design for a small momentum exchange tether experiment that is intended to serve as the first step in
demonstrating these key technologies. The "Microsatellite Tethered Orbit Raising Qualification Experiment" (muTORQUE) will be designed to fly as a secondary payload on an
upper stage of a rocket used to deliver a satellite to GEO. The muTORQUE experiment will remain on the upper stage left in a GTO trajectory. After the primary satellite has
been deployed into GEO, the muTORQUE experiment will deploy a microsatellite at the end of a 20 km long tether. Utilizing tether reeling and or electrodynamic propulsion, the
muTORQUE system will set the tether in rotation around the upper stage, accelerating the rotation until the tip velocity is approximately 400 m s. The experiment will then
release the microsatellite when the system is at its perigee, tossing the payload into a near-minimum-energy transfer to the Moon. The microsatellite can then utilize a Belbruno
weak-boundary trajectory to transfer into a lunar orbit using only a few m s of delta-V. Preliminary analyses indicate that the tether system could be mass-competitive with a
chemical propellant system for the same mission. copyright 2002 American Institute of Physics.
Secondary and Tertiary Launch Technologies
Structures – Launch and Flight Vehicle
Company Name
ATA Engineering, Inc.
Title
DEVELOPMENT OF A PAYLOAD ADAPTER FOR
SATELLITE MISSIONS LAUNCHED USING ICBMDERIVED LAUNCH VEHICLES
DOD SBIR Phase II
Quad Chart
Lightweight payload adapter able to accommodate a wide range of satellite geometries. This innovation reduces cost and part count
and greatly facilitates the manufacturing process while meeting challenging structural requirements without resorting to unproven
and difficult-to-obtain materials.
Secondary and Tertiary Launch Technologies
Structures – Launch and Flight Vehicle
USPTO
Patent Number
Title
US6561461
Orbit Transfer
Vehicle with Support
Services
US6789767
US7114683
Assignee
Aero Astro, Inc.
Abstract
An orbit-transfer vehicle provides the navigation, propulsion, and control systems required to transport a payload satellite from a geosynchronous-transfer
orbit (GTO) to a predetermined low-earth orbit (LEO). Upon entering low-earth orbit, the payload satellite is deployed from the orbit-transfer vehicle. To
reduce the cost and complexity of the payload satellite, the orbit-transfer vehicle is configured to provide common functional services, such as
communications and power regulation, to the payload satellite during the transport, and/or after deployment. To reduce the fuel requirements for this
deployment via the orbit-transfer vehicle, a preferred embodiment includes aerobraking to bring the satellite into a low-earth orbit. In a preferred
embodiment of this method of deployment, the provider of the orbit-transfer vehicle identifies and secures available excess capacity on launch vehicles, and
allocates the excess capacity to the satellites requiring deployment, thereby providing a deployment means that is virtually transparent to the purchaser of
this deployment service.
Active Satellite
Dispenser for
Reusable Launch
Vehicle
Kistler Aerospace
Corporation
An active satellite dispenser is preferably attachable to a reusable launch vehicle for deployment of one or more satellites into one or more desired orbits.
The active satellite dispenser includes a center mast that releasably receives the satellite(s), a liquid propellant rocket, and an orbital control system on an
avionics pallet. In the preferred embodiment, a pressurized gas selectively pressurizes the propellant tanks (which may include fuel and oxidizer tanks), to
provide propellant to the rocket. In operation, the launch vehicle releases the satellite dispenser in a first deployment orbit. The active dispenser rocket and
orbital control system then transport the active dispenser and satellite(s) into the final deployment orbit. In the preferred embodiment the active dispenser
can operate multiple times to place individual satellites in different orbits.
Device and Method
for a Spacecraft
SAAB Ericsson
Space AB
The invention presented here relates to a device in a spacecraft comprising a carrier rocket, a satellite and a detachable adapter ( 4 ) for connecting a satellite
to the carrier rocket, where the adapter ( 4 ) comprises a wall structure ( 14, 15 ) for supporting the satellite on the carrier rocket, and is fixed at its lower rim
to the carrier rocket and is fixed at its upper rim to the satellite. The device is characterised in that the adapter ( 4 ) is divided essentially across the
longitudinal direction of the spacecraft and that both of the adapter parts ( 12, 13 ) are detachably connected to each other. The invention comprises the said
adapter design and a method for separating the invented device.
Secondary and Tertiary Launch Technologies
Structures – Launch and Flight Vehicle
WIPO
Patent Number
Title
WO03059741A1
Launch Lock for
Spacecraft Payloads
WO03082674A1
Assignee
Honeywell
International Inc.
Abstract
A spacecraft payload (12), suspended by an isolator (14), is locked down during launch by using a compliant member (32) to pull on a pin (26),
clamping the payload (12) against the launch lock (10). The payload (12) is unlocked, after launch, by heating a first shape memory alloy element (20)
to move the pin (26) to overcome the compliant member (32), releasing the payload (12) from launch lock (10). The payload (12) may be re-locked for
reentry by heating a second shape memory alloy element (18) to push the pin (26) against the first shape memory alloy element (20).
Spacecraft, Method for
Building Such a
Spacecraft, and
Adaptor to BeUSed in
Such a Spacecraft
Dutch Space B.V.
WO04113170A1
Satellite Launch
Apparatus
Ahad, Mohammad,
Anayet, Rubby
The invention relates to a method of assembling a spacecraft (1), comprising a carrier rocket (2) and at least a first payload (3), such as a satellite or the
like, wherein the first payload is placed on the carrier rocket employing an adapter (4) between the carrier rocket and the first payload, wherein a twopart adapter is used, wherein a first relatively light part (5) of the adapter is mounted under the first payload, and a second relatively heavy part (6) of
the adapter is placed on the carrier rocket, after which the first part of the adapter together with the first payload is placed on the second part of the
adapter and the first part of the adapter is secured is a permanent connection on the second part of the adapter so as to bring the same into the
completed form.
A satellite container (10) for packaging a satellite during transportation thereof to a launch vehicle, comprising a body (12) defining a chamber (20)
for housing a satellite (25), wherein the body (12) is configured to be coupled to or mounted on a launch vehicle for carrying the body (12) into space
with the satellite housed therein. A satellite transportation module (50) for use in a multi-stage launch vehicle is also provided, the transportation
module (50) being configured to be coupled to or to house the satellite container (10).
WO06106500A1
Structure Coupling and
Coupler Therefore
Israel Aircraft
Industries Ltd.
WO06131818A2
Human-Propelled
Method and Apparatus
for Launching a
Satellite
Method for Placing in
Operational Orbit an
Artificial Satellite and
Associated Propulsion
Device
Icon Orbital
Technologies
Corporation
WO28001002A1
AstriumSas
A coupler (36) comprising a lower lug (40) and an upper lug (42) each formed with a locking recess (54, 58) engageable with a corresponding locking
piece (52) slindingly received within a cavity (50) extending between lugs. The locking piece is displaceable between a locked position in which it
engages the locking recess of the lugs and the lugs are tightly coupled to one another, and an open position in which it disengages from the recesses to
allow detaching of the lugs from one another. At least one of the lugs and the locking piece are formed with at leas.t one ejecting surfaces to facilitate
disengagement and separation of the lugs from one another and sliding displacement of the locking piece, further comprising a power activated
separation system (130) for displacing the locking piece into its open position.
A manual method of launching a satellite from space borne vehicle includes fixing a launching platform onto a space vehicle and fixing a satellite
holder on the launching platform. A Space suited person is fastened to the launching platform and a satellite is engaged with the satellite holder
adjacent to the person. The person throws or strikes the satellite to dislodge it from the holder and to direct it in a path away from the vehicle.
The method for placing in operational orbit (9) a satellite (3) provided with a specific system of propulsion from a transfer orbit (7) obtained by means
of a launch vehicle (6) is such that a propulsion device (1) controlled by a satellite (3) is attached in a separable manner to said latter, and the assembly
(1-3) thus constituted is mounted on the launch vehicle (6) before the injection of said assembly (1-3) by the launch vehicle (6) on the transfer orbit
(7), then said assembly (1-3) is brought, via the propulsion device (1), on an intermediate orbit (8) between the transfer orbit (7) and the operational
orbit (9), the intermediate orbit (8) neighbouring the operational orbit (9) but being far enough away to prevent possible interferences, then the satellite
(3) is separated from the propulsion device (1), which remains on the intermediate orbit (8), and the satellite (3) rejoins, by means of its specific system
of propulsion, the operational orbit (9) from the intermediate orbit (8).
Secondary and Tertiary Launch Technologies
Structures – Launch and Flight Vehicle
FC
GSFC
Title
Deployment Mechanism for the Space Technology 5
Micro Satellite
NTRS
Abstract
Space Technology 5 (ST5) is a technology mission that will send three spin-stabilized, 25-kg satellites into a highly elliptical Earth orbit. Each of these satellites
must be deployed separately from the same launch vehicle with a spin rate of 3.4 rads (32.4 rpm). Because of the satellite's small size and the requirement to
achieve its mission spin rate on deploy, typical spin table, pyrotechnic deployment devices or spin up thrusters could not be used. Instead, this new
mechanism design employs a "Frisbee" spin up strategy with a shape memory alloy actuated Pinpuller to deploy each satellite. The mechanism has
undergone several design and test iterations and has been successfully qualified for flight.
Secondary and Tertiary Launch Technologies
Structures – Launch and Flight Vehicle
WIPO
Patent Number
Title
WO04012995A1
A Secondary Payload Satellite
Assignee
Astrium Limited
Abstract
An improved satellite design (10), carrying an onboard propulsion system, has a central box structure compartment (1), and a
plurality of outer box structure compartments (2, 3) containing propellant tanks of the propulsion system which are in turn
positioned above a plurality of launch vehicle attachment points/locations (4, 5). Significantly, the proposed arrangement of the
invention bridges mechanically the various separate attachment points/locations, permitting the available mass and volume for
the spacecraft to be effectively increased.
Low-Cost, Rapid Spacecraft Design and Multi-Subsystem Functionality
Avionics and Astrionics –Attitude Determination and Control
Company Name
Milli Sensor Systems and
Actuators, Inc.
Title
Stable Tactical-Grade MEMS IMU for SpinStabilized Rockets
Field Center
GSFC
NASA SBIR Phase I
Quad Chart
An Integrated MEMS IMU is proposed that will operate effectively in a spinning rocket up to 7 revs/sec. The IMU contains three
gyroscopes and nine accelerometers on the same chip. The instrument designs have the low cross-axis sensitivity that permit the
orthogonal gyros to sense the relatively smaller pitch and yaw rates in the presence of the much larger rate about the spin axis. An
algorithm is proposed to combine the signals from the instruments to co-operatively obtain spatial reference. A lab experiment is
planned during Phase I that will use available equipment and MSSA IMUs to prove the concept.
Low-Cost, Rapid Spacecraft Design and Multi-Subsystem Functionality
Avionics and Astrionics –Guidance, Navigation, and Control
WIPO
Patent Number
Title
WO02077660A2
A System for the Delivery and
Orbital Maintenance of Micro
Satellites and Small SpaceBased Instruments
Assignee
Space Launch Corporation
Abstract
A low cost, on demand, dedicated launch system is provided for placing micro satellites or space-based instruments at orbital and suborbital altitudes and velocities. The invention describes a space launch vehicle (SLV) that incorporates a single, integrated guidance,
navigation, and control unit (GNCU) that performs all guidance and control for the SLV from main stage ignition to orbital insertion. The
GNCU can remain with the payload after orbital insertion to provide satellite station keeping and orbital maneuvering capability. The use
of a single integrated avionics unit for all guidance, navigation, and control simplifies the SLV, reducing weight and significantly
reducing cost. In addition, this architecture allows for a combined launch and satellite bus system as the GNCU can also be used as a
satellite bus. This further reduces cost and increases the payload capacity to orbit by optimizing the use of launch vehicle and satellite bus
subsystems and reducing non-instrument mass delivered to orbit. All support functions are provided by the IDMV. This approach
represents a significant improvement over conventional systems, especially with respect to the orbital launch of payloads less than about
100 kg.
Low-Cost, Rapid Spacecraft Design and Multi-Subsystem Functionality
Avionics and Astrionics –Guidance, Navigation, and Control
FC
Goddard Space Flight Center
Title
Integrated Orbit and Attitude Control for a
Nanosatellite with Power Constraints
NTRS
Abstract
Small satellites tend to be power-limited, so that actuators used to control the orbit and attitude must compete with each other as well as with other
subsystems for limited electrical power. The Virginia Tech nanosatellite project, HokieSat, must use its limited power resources to operate pulsedplasma thrusters for orbit control and magnetic torque coils for attitude control, while also providing power to a GPS receiver, a crosslink transceiver,
and other subsystems. The orbit and attitude control strategies were developed independently. The attitude control system is based on an application of
LQR to an averaged system of equations, whereas the orbit control is based on orbit element feedback. In this paper we describe the strategy for
integrating these two control systems and present simulation results to verify the strategy.
Low-Cost, Rapid Spacecraft Design and Multi-Subsystem Functionality
Avionics and Astrionics –On-Board Computing and Data Management
FC
Jet Propulsion Laboratory
Title
Towards a distributed information architecture
for avionics data
NTRS
Abstract
Avionics data at the National Aeronautics and Space Administration's (NASA) Jet Propulsion Laboratory (JPL consists of distributed, unmanaged, and heterogeneous
information that is hard for flight system design engineers to find and use on new NASA JPL missions. The development of a systematic approach for capturing,
accessing and sharing avionics data critical to the support of NASA JPL missions and projects is required. We propose a general information architecture for
managing the existing distributed avionics data sources and a method for querying and retrieving avionics data using the Object Oriented Data Technology (OODT)
framework. OODT uses XML messaging infrastructure that profiles data products and their locations using the ISO-11179 data model for describing data products.
Queries against a common data dictionary (which implements the ISO model) are translated to domain dependent source data models, and distributed data products
are returned asynchronously through the OODT middleware. Further work will include the ability to 'plug and play' new manufacturer data sources, which are
distributed at avionics component manufacturer locations throughout the United States.
Low-Cost, Rapid Spacecraft Design and Multi-Subsystem Functionality
Communications – Architectures and Networks
Company Name
Planning Systems, Inc.
Title
Plug and Play Compatibility Enhancements for the Network Data
Acquisition System (NDAS)
DOD SBIR Phase I
Quad Chart
The Air Force has an operational need to launch satellites on demand for a variety of missions as well as to provide last minute tailoring
for dynamic mission requirements. One of the critical capabilities required to achieve this goal is the ability to rapidly integrate all the
necessary satellite subsystems, including mission specific payloads, in a short time period before launch. The capability to rapidly
integrate components prior to launch also supports the ability to rapidly reconfigure a spacecraft to accommodate changing mission
requirements. In support of this operational need, the AF has established a roadmap for the development of satellite components that
can perform these functions, with the initial phase concentrating on establishment of a "Plug and Play" (PnP) interface similar to those
implemented on standard personal computers via Universal Serial Bus (USB). In support of this Phase I SBIR, PSI is modifying their
existing high-speed, dual redundant, multi-drop network for space (Network Data Acquisition System - NDAS) by adding USB
compatibility. We will provide an integrated PnP design that that can plug into an existing USB device as well as host numerous USB
components along the network. The Phase II effort completes the design work for an ASIC implementation.
Low-Cost, Rapid Spacecraft Design and Multi-Subsystem Functionality
Communications – Architectures and Networks
Company Name
SI2 Technologies
Title
INTEGRATION OF MULTIFUNCTIONAL
CONFORMAL ELECTRONIC STRUCTURES FOR
ADVANCED AIRCRAFT (1000-019)
DOD SBIR Phase II
Quad Chart
The technology is applicable to a wide variety of electrical systems, including communication and phased array antennas, and structural
locations, such as wings and fuselage. The Phase II program will build upon Phase I success and further develop the technology to demonstrate
an advanced conformal antenna system for a long endurance UAV.
Low-Cost, Rapid Spacecraft Design and Multi-Subsystem Functionality
Communications – Architectures and Networks
U.S. Patent Applications
Patent Number
Title
US2007299864A1
Object Storage Subsystem
Computer Program
Assignee
Strachan, Mark
Abstract
An object storage subsystem program with federated object storage on multiple computing nodes, which may be added as a
component to existing open source platforms. The subsystem program increases programming efficiency by leveraging existing
open source solutions, directly integrating with an application development framework, increasing the efficiency of the framework,
and allowing other mechanisms to be introduced that ease implementation for large scale enterprise software development. The
program also provides an object storage subsystem with multiple modes of operation to provide high availability and fault tolerant
object storage, as well as the capability to manage a massive amount of data across multiple computing nodes with features that
enable it to store data on hard drives, clean up unused data, isolate and manage transactions, and provide communication between
storage nodes.
Low-Cost, Rapid Spacecraft Design and Multi-Subsystem Functionality
Communications – Architectures and Networks
FC
Goddard Space Flight Center
Title
Center for Space Telemetering and
Telecommunications Systems, New Mexico State
University
NTRS
Abstract
This viewgraph presentation gives an overview of the Center for Space Telemetering and Telecommunications Systems activities at New Mexico
State University. Presentations cover the following topics (1) small satellite communications, including nanosatellite radio and virtual satellite
development (2) modulation and detection studies, including details on smooth phase interpolated keying (SPIK) spectra and highlights of an
adaptive turbo multiuser detector (3) decoupled approaches to nonlinear ISI compensation (4) space internet testing (4) optical communication
(5) Linux-based receiver for lightweight optical communications without a laser in space, including software design, performance analysis, and
the receiver algorithm (6) carrier tracking hardware and (7) subband transforms for adaptive direct sequence spread spectrum receivers.
Low-Cost, Rapid Spacecraft Design and Multi-Subsystem Functionality
Communications – Autonomous Control and Monitoring
FC
GSFC
Title
A Survey of Formal Methods for
Intelligent Swarms
GSFC
ANTS Applying A New Paradigm
for Lunar and Planetary
Exploration
GSFC
Asteroid Exploration with
Autonomic Systems
NTRS
Abstract
Swarms of intelligent autonomous spacecraft, involving complex behaviors and interactions, are being proposed for future space exploration missions. Such missions provide greater
flexibility and offer the possibility of gathering more science data than traditional single spacecraft missions. The emergent properties of swarms make these missions powerful, but
simultaneously far more difficult to design, and to assure that the proper behaviors will emerge. These missions are also considerably more complex than previous types of missions, and
NASA, like other organizations, has little experience in developing or in verifying and validating these types of missions. A significant challenge when verifying and validating swarms of
intelligent interacting agents is how to determine that the possible exponential interactions and emergent behaviors are producing the desired results. Assuring correct behavior and
interactions of swarms will be critical to mission success. The Autonomous Nano Technology Swarm (ANTS) mission is an example of one of the swarm types of missions NASA is
considering. The ANTS mission will use a swarm of picospacecraft that will fly from Earth orbit to the Asteroid Belt. Using an insect colony analogy, ANTS will be composed of specialized
workers for asteroid exploration. Exploration would consist of cataloguing the mass, density, morphology, and chemical composition of the asteroids, including any anomalous
concentrations of specific minerals. To perform this task, ANTS would carry miniaturized instruments, such as imagers, spectrometers, and detectors. Since ANTS and other similar
missions are going to consist of autonomous spacecraft that may be out of contact with the earth for extended periods of time, and have low bandwidths due to weight constraints, it will be
difficult to observe improper behavior and to correct any errors after launch. Providing V and ampV (verification and validation) for this type of mission is new to NASA, and represents the
cutting edge in system correctness, and requires higher levels of assurance than other (traditional) missions that use a single or small number of spacecraft that are deterministic in nature
and have near continuous communication access. One of the highest possible levels of assurance comes from the application of formal methods. Formal methods are mathematics-based
tools and techniques for specifying and verifying (software and hardware) systems. They are particularly useful for specifying complex parallel systems, such as exemplified by the ANTS
mission, where the entire system is difficult for a single person to fully understand, a problem that is multiplied with multiple developers. Once written, a formal specification can be used to
prove properties of a system (e.g., the underlying system will go from one state to another or not into a specific state) and check for particular types of errors (e.g., race or livelock
conditions). A formal specification can also be used as input to a model checker for further validation. This report gives the results of a survey of formal methods techniques for verification
and validation of space missions that use swarm technology. Multiple formal methods were evaluated to determine their effectiveness in modeling and assuring the behavior of swarms of
spacecraft using the ANTS mission as an example system. This report is the first result of the project to determine formal approaches that are promising for formally specifying swarm-based
systems. From this survey, the most promising approaches were selected and are discussed relative to their possible application to the ANTS mission. Future work will include the
application of an integrated approach, based on the selected approaches identified in this report, to the formal specification of the ANTS mission.
ANTS (Autonomous Nano- Technology Swarm), a mission architecture consisting of a large (1000 member) swarm of picoclass (1 kg) totally autonomous spacecraft with both adaptable
and evolvable heuristic systems, is being developed as a NASA advanced mission concept, and is here examined as a paradigm for lunar surface exploration. As the capacity and
complexity of hardware and software, demands for bandwidth, and the sophistication of goals for lunar and planetary exploration have increased, greater cost constraints have led to fewer
resources and thus, the need to operate spacecraft with less frequent human contact. At present, autonomous operation of spacecraft systems allows great capability of spacecraft to 'safe'
themselves and survive when conditions threaten spacecraft safety. To further develop spacecraft capability, NASA is at the forefront of development of new mission architectures which
involve the use of Intelligent Software Agents (ISAs), performing experiments in space and on the ground to advance deliberative and collaborative autonomous control techniques. Selected
missions in current planning stages require small groups of spacecraft weighing tens, instead of hundreds, of kilograms to cooperate at a tactical level to select and schedule measurements
to be made by appropriate instruments onboard. Such missions will be characterizing rapidly unfolding real-time events on a routine basis. The next level of development, which we are
considering here, is in the use of autonomous systems at the strategic level, to explore the remote terranes, potentially involving large surveys or detailed reconnaissance.
NASA is studying advanced technologies for a future robotic exploration mission to the asteroid belt. The prospective ANTS (Autonomous Nano Technology Swarm) mission comprises
autonomous agents including worker agents (small spacecra3) designed to cooperate in asteroid exploration under the overall authoriq of at least one ruler agent (a larger spacecraft)
whose goal is to cause science data to be returned to Earth. The ANTS team (ruler plus workers and messenger agents), but not necessarily any individual on the team, will exhibit
behaviors that qualify it as an autonomic system, where an autonomic system is defined as a system that self-reconfigures, self-optimizes, self-heals, and self-protects. Autonomic system
concepts lead naturally to realistic, scalable architectures rich in capabilities and behaviors. In-depth consideration of a major mission like ANTS in terms of autonomic systems brings new
insights into alternative definitions of autonomic behavior. This paper gives an overview of the ANTS mission and discusses the autonomic properties of the mission.
Low-Cost, Rapid Spacecraft Design and Multi-Subsystem Functionality
Communications – Autonomous Control and Monitoring
NTRS
Abstract
NASA is working on new mission concepts for exploration of the solar system. The concepts for these missions include swarms of hundreds of cooperating intelligent spacecraft which will
be able to work in teams and gather more data than current single spacecraft missions. These spacecraft will not only have to operate independently for long periods of time on their own
and in teams, but will also need to have autonomic properties of self healing, self configuring, self optimizing and self protecting for them to survive in the harsh space environment. Software
for these types of missions has never been developed before and represents some of the challenges of software development in the new millennia. The Autonomous Nano Technology
Swarm (ANTS) mission is an example of one of the swarm missions NASA is considering. The ANTS mission will use a swarm of one thousand pico-spacecraft that weigh less than five
pounds. Using an insect colony analog, ANTS will explore the asteroid belt and catalog the mass, density, morphology, and chemical composition of the asteroids. Due to the size of the
spacecraft, each will only carry a single miniaturized science instrument which will require them to cooperate in searching for asteroids that are of scientific interest. This article also
discusses the ANTS mission, the properties the spacecraft will need and how that will effect future software development.
FC
GSFC
Title
Developing Software for NASA
Missions in the New Millennia
GSFC
Exploring with PAM Prospecting
ANTS Missions for Solar System
Surveys
ANTS (Autonomous Nano-Technology Swarm), a large (1000 member) swarm of nano to picoclass (10 to 1 kg) totally autonomous spacecraft, are being developed as a NASA advanced
mission concept. ANTS, based on a hierarchical insect social order, use an evolvable, self-similar, hierarchical neural system in which individual spacecraft represent the highest level
nodes. ANTS uses swarm intelligence attained through collective, cooperative interactions of the nodes at all levels of the system. At the highest levels this can take the form of cooperative,
collective behavior among the individual spacecraft in a very large constellation. The ANTS neural architecture is designed for totally autonomous operation of complex systems including
spacecraft constellations. The ANTS (Autonomous Nano Technology Swarm) concept has a number of possible applications. A version of ANTS designed for surveying and determining the
resource potential of the asteroid belt, called PAM (Prospecting ANTS Mission), is examined here.
GSFC
From Present Surveying to Future
Prospecting of the Asteroid Belt
We have applied a future mission architecture, the Autonomous Nano-Technology Swarm (ANTS), to a proposed mission for in situ survey, or prospecting, of the asteroid belt, the
Prospecting Asteroid Mission (PAM) as part of a NASA 2003 Revolutionary Aerospace Concept (RASC) study. ANTS architecture builds on and advances recent trends in robotics, artificial
intelligence, and materials processing to minimize costs and maximize effectiveness of space operations. PAM and other applications have been proposed for the survey of inaccessible,
high surface area populations of great interest from the standpoint of resources and or solar system origin. The ANTS architecture is inspired by the success of social insect colonies, a
success based on the division of labor within the colonies in two key ways 1) within their specialties, individual specialists generally outperform generalists, and 2) with sufficiently efficient
social interaction and coordination, the group of specialists generally outperforms the group of generalists. Thus systems designed as ANTS are built from potentially very large numbers of
highly autonomous, yet socially interactive, elements. The architecture is self-similar in that elements and sub-elements of the system may also be recursively structured as ANTS on scales
ranging from microscopic to interplanetary distances. Here, we analyze requirements for the mission application at the low gravity target end of the spectrum, the Prospecting Asteroid
Mission (PAM), and for specialized autonomous operations which would support this mission. ANTS as applied to PAM involves the activities of hundreds of individual specialist
'sciencecraft'. Most of them, called Workers, carry and operate eight to nine different scientific instruments, as listed in the table, including spectrometers, ranging and radio science devices,
and imagers. The remaining specialists, Messenger Rulers, provide communication and coordination functions among specialists operating autonomously as individuals, team members,
and subswarms.
Low-Cost, Rapid Spacecraft Design and Multi-Subsystem Functionality
Communications – RF
FC
Glenn Research Center
Title
A Compact, Broadband Antenna for
Planetary Surface-to-Surface Wireless
Communications
NTRS
Abstract
The Compact Microstrip Monopole Antenna (CMMA) is a novel antenna design that combines a microstrip patch antenna with a three-dimensional structure to attain a highly
directive, broadband, compact antenna. A Tri-Lobed Patch (TLP) was designed to minimize the patch's area while reducing the antenna's operating frequency. A Grounding
Wall (GW) connects the patch to the ground plane and a Vertical Enclosure Wall (VEW) extends up away from portions of the patch's perimeter. This VEW supplies the
antenna with a higher directivity in the radial direction as well as reduces the operating frequency. The CMMA was designed to operate at 2.23 GHz, but experimental
results have shown this antenna resonates at 2.05 GHz which is on the order of approximately Lambda(sub o) 11.6 with respect to the antenna's largest dimension, with a
directivity and bandwidth of 6.0 dBi, and 130 MHz (6.3 percent), respectively. This miniature, radially emitting antenna makes the CMMA attractive for planetary-based
surface-to-surface communications.
Low-Cost, Rapid Spacecraft Design and Multi-Subsystem Functionality
Electronics – Highly-Reconfigurable
FC
Goddard Space Flight
Center
Title
SpaceWire Plug and Play
NTRS
Abstract
The ability to rapidly deploy inexpensive satellites to meet tactical goals has become an important goal for military space systems. In fact, Operationally Responsive Space (ORS) has
been in the spotlight at the highest levels. The Office of the Secretary of Defense (OSD) has identified that the critical next step is developing the bus standards and modular interfaces.
Historically, satellite components have been constructed based on bus standards and standardized interfaces. However, this has not been done to a degree, which would allow the rapid
deployment of a satellite. Advancements in plug-and-play (PnP) technologies for terrestrial applications can serve as a baseline model for a PnP approach for satellite applications. Since
SpaceWire (SpW) has become a de facto standard for satellite high-speed (greater than 200Mbp) on-board communications, it has become important for SpW to adapt to this Plug and
Play (PnP) environment. Because SpW is simply a bulk transport protocol and lacks built-in PnP features, several changes are required to facilitate PnP with SpW. The first is for Host(s)
to figure out what the network looks like, i.e., how pieces of the network, routers and nodes, are connected together network mapping, and to receive notice of changes to the network.
The second is for the components connected to the network to be understood so that they can communicate. The first element, network topology mapping and amp change of status
indication, is being defined (topic of this paper). The second element describing how components are to communicate has been defined by ARFL with the electronic data sheets known as
XTEDS. The first element, network mapping, is recent activities performed by Air Force Research Lab (ARFL), Naval Research Lab (NRL), NASA and US industry (Honeywell, Clearwater,
FL, and others). This work has resulted in the development of a protocol that will perform the lower level functions of network mapping and Change Of Status (COS) indication required by
Plug 'n' Play over SpaceWire. This work will be presented to the SpaceWire working group for standardization under European Cooperation for Space Standardization (ECSS) and to
obtain a permanent Protocol ID (see SpaceWire Protocol ID What Does it Mean to You IEEE Aerospace Conference 2006). The portion of the Plug 'n' Play protocol that will be described
in this paper is how the Host(s) of a SpaceWire network map the network and detect additions and deletions of devices on a SpaceWire network.
Low-Cost, Rapid Spacecraft Design and Multi-Subsystem Functionality
Electronics – Photonics
Company Name
Opticomp Corporation and California
Institute of Technology
Title
CHIP-TO-CHIP OPTICAL INTERCONNECTS
USING INTEGRATED PHOTONIC CRYSTAL
LIGHTWAVE CIRCUITS
DOD SBIR Phase II
Quad Chart
This Phase II is to develop photonic crystal technologies which can enable photonic lightwave circuits, providing a new technology platform for the
heterogeneous integration of photonic and optoelectronic functions. Coherent PC-based devices can perform novel optical functions such as selfmodulation, optical clock generation and distribution, optical logic and optical storage.
Low-Cost, Rapid Spacecraft Design and Multi-Subsystem Functionality
Electronics – Ultra-High Density/Low Power
FC
GSFC/JPL
Title
Advanced Instruments and Their Impact on Earth
Science Missions (I)
NTRS
Abstract
A past paper (IAA-B4-1004) analyzed the costs associated with developing, launching and operating a constellation of small satellites for earth
observations. That study provided examples of measurements that could be made with such a system, per-unit cost goals and an overview of
technologies that might be applied to spacecraft (s c) subsystems to minimize power, mass and volume. However, that paper largely ignored
instruments and the impacts of instrument size, mass, and power reductions on future mission feasibility. This paper reviews instruments that have
been or are being funded by NASA to reduce power consumption, mass, volume to lower downlink demands and or lower unit costs. The
instruments in question could be used in single spacecraft missions or implemented in spacecraft constellations to increase temporal or spatial
coverage. The instrument descriptions describe measurements enabled and or enhanced, concluding with descriptions of the instruments.
Instruments chosen consist primarily of those with low s c system resource demands, e.g., power or communications bandwidth.
Low-Cost, Rapid Spacecraft Design and Multi-Subsystem Functionality
Information – Autonomous Reasoning/Artificial Intelligence
FC
Langley Research Center
Title
Integrated System-Level Optimization for Concurrent
Engineering With Parametric Subsystem Modeling
NTRS
Abstract
The introduction of concurrent design practices to the aerospace industry has greatly increased the productivity of engineers and teams during
design sessions as demonstrated by JPL's Team X. Simultaneously, advances in computing power have given rise to a host of potent
numerical optimization methods capable of solving complex multidisciplinary optimization problems containing hundreds of variables,
constraints, and governing equations. Unfortunately, such methods are tedious to set up and require significant amounts of time and processor
power to execute, thus making them unsuitable for rapid concurrent engineering use. This paper proposes a framework for Integration of
System-Level Optimization with Concurrent Engineering (ISLOCE). It uses parametric neural-network approximations of the subsystem models.
These approximations are then linked to a system-level optimizer that is capable of reaching a solution quickly due to the reduced complexity of
the approximations. The integration structure is described in detail and applied to the multiobjective design of a simplified Space Shuttle
external fuel tank model. Further, a comparison is made between the new framework and traditional concurrent engineering (without system
optimization) through an experimental trial with two groups of engineers. Each method is evaluated in terms of optimizer accuracy, time to
solution, and ease of use. The results suggest that system-level optimization, running as a background process during integrated concurrent
engineering sessions, is potentially advantageous as long as it is judiciously implemented.
Low-Cost, Rapid Spacecraft Design and Multi-Subsystem Functionality
Information – Computer System Architectures
Company Name
sci_zone
Title
OpenSAT, An Open Source
Based Satellite Design Data
Architecture with API Design and
Management Plugins
Field Center
ARC
NASA SBIR Phase I
Quad Chart
Satellite design encompasses a multitude of steps from concept to flight. Mission specification to flight can take several years, depending on
the scope, requirements and budget of the mission. The process also requires a wide range of design and management tools, with limited
consistency data interchange capability, and a lack of coherency. Detailing the relationships between the satellite configuration, inventory
control systems, life cycle management, design, analysis and test data is difficult at best. No tool exists that meets these needs for the general
satellite design, system engineering and integration process. Sci_Zone is proposing our innovative Satellite Design Automation architecture
SatBuilder Designer, in conjunction with the OpenSAT open database architecture to meet this need. OpenSAT seamlessly integrates existing
detail design tools with SatBuilder Designer, as well as databases tracking requirements, components and inventory, with the final configuration
of the satellite. SatBuilder Designer, an AI based toolset, provides for rapid design via design wizards and integration to existing design tools; it
provides coherency between a range of applications and data sets. OpenSAT stores and distributes supporting satellite design, configuration,
mission and test data from a centralized database server and can distribute the data across multiple platforms
and via the internet.
Low-Cost, Rapid Spacecraft Design and Multi-Subsystem Functionality
Information – Computer System Architectures
Company Name
Research South, Inc.
Title
UNSTRUCTURED MESH MOVEMENT AND VISCOUS
MESH GENERATION
FOR CFD-BASED DESIGN OPTIMIZATION
Field Center
DFRC
NASA SBIR Phase II
Quad Chart
Develop two types of user interfaces: graphical access (for the end-user) and programming access for integration with flow
solvers. Assemble all of the methods developed in Phase II into a single, coherent, design-oriented, product-version code with
extensive focus on incorporating a parallel processing capability into the software.
Low-Cost, Rapid Spacecraft Design and Multi-Subsystem Functionality
Information – Computer System Architectures
Company Name
Seakr Engineering Inc.
Title
LOW COST, TAILOR ABLE AVIONICS FOR
RAPID RESPONSE SATELLITES
DOD SBIR Phase II
Quad Chart
The RCC architecture reduces risk, costs, and schedule for satellite missions by providing a reconfigurable space based platform that could be used for a
number of different spacecraft avionic applications. It provides a platform that can not only support rapid, low cost plug-and-play integration into a spacecraft
but because it is reconfigurable, it may also be used for many different subsystems.
Low-Cost, Rapid Spacecraft Design and Multi-Subsystem Functionality
Information – Data Acquisition and End-to-End Management
Company Name
Microcosm, Inc.
Title
A Sustainable Spacecraft Component
Database Solution
Field Center
ARC
NASA SBIR Phase I
Quad Chart
Numerous spacecraft component databases have been developed to support NASA, DoD, and contractor design centers and
design tools. Despite the clear utility of component databases for improving and accelerating the spacecraft design process, they
are observably uneconomic for individual organizations to build and maintain –otherwise there would be several that were
actively maintained and growing. The problem is not the challenge of architecting the database or identifying the component types
and data elements. In fact, many organizations have succeeded in developing excellent solutions that have languished due to the
cost of populating and maintaining them. Microcosm is focused on developing an economically viable solution to creating a
sustainable component database. The approach is four-fold: 1. Reduce the cost of gathering/maintaining data by leveraging webbased data with automated gathering and monitoring tools; 2. Drive sufficient value to component suppliers such that they are
motivated to supply and maintain their own data; 3. Supply data to a broad range of users and organizations; and 4. Enable
value-added contributions from users via venues like wikis, discussion rooms, links to relevant websites. This solution will stand
on its own as a viable business, and provide increasing value to customers (suppliers and users) over
time.
Low-Cost, Rapid Spacecraft Design and Multi-Subsystem Functionality
Information – Data Acquisition and End-to-End Management
Company Name
RED CANYON SOFTWARE
Title
OpenSAT, a framework for satellite design
automation for responsive space
DOD SBIR Phase I
Quad Chart
Satellite design encompasses a multitude of steps from concept to flight. Mission specifications to flight can take several years, depending on the scope,
requirements and budget of the mission. A key requirement of the Air Force’s responsive space program is the capability to assemble, test, and launch a
satellite within days, or even hours, versus the years of a battlefield Commander’s request. An operationally responsive space capability, such as this,
will provide our forces with an advantage in future conflicts. The Commander will be able to rapidly deploy tailored space assets to strategic orbits.
Currently, this capability does not exist. Red Canyon Software is proposing our innovative Satellite Design Automation (SDA) architecture SatBuilder™,
in conjunction with our OpenSAT open architecture. This suite of products seamlessly integrates existing detail design tools with SatBuilder™, as well as
databases tracking requirements, PnP components and payloads in inventory, with the final configuration of the satellite. SatBuilder™, an AI based
toolset, provides for not only rapid design, via design wizards and integration to existing design tools, but will also generate the code for PnP components
based on the final design, generate the interface for a ground control station, and track satellite test and certification.
Low-Cost, Rapid Spacecraft Design and Multi-Subsystem Functionality
Information – Database Development and Interfacing
U.S. Patent Applications
Patent Number
Title
US2005055328A1
Method and apparatus for
data integration
Assignee
Hitachi, Ltd.
Abstract
Aspects of the present invention provide integration of geographically distributed data. The data can be integrated in a single
database. An illustrative embodiment of the invention comprises a tight combination between conventional ETL (extraction,
translation, and loading) tools and conventional remote copy functionality used for data backup and recovery.
Low-Cost, Rapid Spacecraft Design and Multi-Subsystem Functionality
Information – Database Development and Interfacing
FC
Marshall Space Flight
Center
Stennis Space Center
Title
Intelligent Software for
System Design and
Documentation
Goddard Space Flight
Center
ISAIA Interoperable Systems
for Archival Information
Access
NTRS
Abstract
In an effort to develop a real-time, on-line database system that tracks documentation changes in NASA's propulsion test facilities, engineers at Stennis Space Center teamed with
ECT International of Brookfield, WI, through the NASA Dual-Use Development Program to create the External Data Program and Hyperlink Add-on Modules for the promise
software. Promise is ECT's top-of-the-line intelligent software for control system design and documentation. With promise the user can make use of the automated design process
to quickly generate control system schematics, panel layouts, bills of material, wire lists, terminal plans and more. NASA and its testing contractors currently use promise to create
the drawings and schematics at the E2 Cell 2 test stand located at Stennis Space Center.
The ISAIA project was originally proposed in 1999 as a successor to the informal AstroBrowse project. AstroBrowse, which provided a data location service for astronomical
archives and catalogs, was a first step toward data system integration and interoperability. The goals of ISAIA were ambitious '...To develop an interdisciplinary data location and
integration service for space science. Building upon existing data services and communications protocols, this service will allow users to transparently query hundreds or
thousands of WWW-based resources (catalogs, data, computational resources, bibliographic references, etc.) from a single interface. The service will collect responses from
various resources and integrate them in a seamless fashion for display and manipulation by the user.' Funding was approved only for a one-year pilot study, a decision that in
retrospect was wise given the rapid changes in information technology in the past few years and the emergence of the Virtual Observatory initiatives in the US and worldwide.
Indeed, the ISAIA pilot study was influential in shaping the science goals, system design, metadata standards, and technology choices for the virtual observatory. The ISAIA pilot
project also helped to cement working relationships among the NASA data centers, US ground-based observatories, and international data centers. The ISAIA project was formed
as a collaborative effort between thirteen institutions that provided data to astronomers, space physicists, and planetary scientists. Among the fruits we ultimately hoped would
come from this project would be a central site on the Web that any space scientist could use to efficiently locate existing data relevant to a particular scientific question.
Furthermore, we hoped that the needed technology would be general enough to allow smaller, more-focused community within space science could use the same technologies
and standards to provide more specialized services. A major challenge to searching for data across a broad community is that information that describe some data products are
either not relevant to other data or not applicable in the same way. Some previous metadata standard development efforts (e.g., in the earth science and library communities)
have produced standards that are very large and difficult to support. To address this problem, we studied how a standard may be divided into separable pieces. Data providers
that wish to participate in interoperable searches can support only those parts of the standard that are relevant to them. We prototyped a top-level metadata standard that was
small and applicable to all space science data.
Low-Cost, Rapid Spacecraft Design and Multi-Subsystem Functionality
Information – Expert System
FC
Glenn Research Center
Title
A Management Model for
International Participation in Space
Exploration Missions
Glenn Research Center
A Practical Approach to Starting
Fission Surface Power Development
Ames Research Center
Design Methods and Practices for
Fault Prevention and Management in
Spacecraft
NTRS
Abstract
This paper proposes an engineering management model for NASA's future space exploration missions based on past experiences working with the International Partners of
the International Space Station. The authors have over 25 years of combined experience working with the European Space Agency, Japan Aerospace Exploration Agency,
Canadian Space Agency, Italian Space Agency, Russian Space Agency, and their respective contractors in the design, manufacturing, verification, and integration of their
elements electric power system into the United States on-orbit segment. The perspective presented is one from a specific sub-system integration role and is offered so that the
lessons learned from solving issues of technical and cultural nature may be taken into account during the formulation of international partnerships. Descriptions of the types of
unique problems encountered relative to interactions between international partnerships are reviewed. Solutions to the problems are offered, taking into consideration the
technical implications. Through the process of investigating each solution, the important and significant issues associated with working with international engineers and
managers are outlined. Potential solutions are then characterized by proposing a set of specific methodologies to jointly develop spacecraft configurations that benefits all
international participants, maximizes mission success and vehicle interoperability while minimizing cost.
The Prometheus Power and Propulsion Program has been reformulated to address NASA needs relative to lunar and Mars exploration. Emphasis has switched from the
Jupiter Icy Moons Orbiter (JIMO) flight system development to more generalized technology development addressing Fission Surface Power (FSP) and Nuclear Thermal
Propulsion (NTP). Current NASA budget priorities and the deferred mission need date for nuclear systems prohibit a fully funded reactor Flight Development Program.
However, a modestly funded Advanced Technology Program can and should be conducted to reduce the risk and cost of future flight systems. A potential roadmap for FSP
technology development leading to possible flight applications could include three elements 1) Conceptual Design Studies, 2) Advanced Component Technology, and 3) NonNuclear System Testing. The Conceptual Design Studies would expand on recent NASA and DOE analyses while increasing the depth of study in areas of greatest
uncertainty such as reactor integration and human-rated shielding. The Advanced Component Technology element would address the major technology risks through
development and testing of reactor fuels, structural materials, primary loop components, shielding, power conversion, heat rejection, and power management and distribution
(PMAD). The Non-Nuclear System Testing would provide a modular, technology testbed to investigate and resolve system integration issues.
Integrated Systems Health Management (ISHM) is intended to become a critical capability for all space, lunar and planetary exploration vehicles and systems at NASA.
Monitoring and managing the health state of diverse components, subsystems, and systems is a difficult task that will become more challenging when implemented for longterm, evolving deployments. A key technical challenge will be to ensure that the ISHM technologies are reliable, effective, and low cost, resulting in turn in safe, reliable, and
affordable missions. To ensure safety and reliability, ISHM functionality, decisions and knowledge have to be incorporated into the product lifecycle as early as possible, and
ISHM must be considered as an essential element of models developed and used in various stages during system design. During early stage design, many decisions and
tasks are still open, including sensor and measurement point selection, modeling and model-checking, diagnosis, signature and data fusion schemes, presenting the best
opportunity to catch and prevent potential failures and anomalies in a cost-effective way. Using appropriate formal methods during early design, the design teams can
systematically explore risks without committing to design decisions too early. However, the nature of ISHM knowledge and data is detailed, relying on high-fidelity, detailed
models, whereas the earlier stages of the product lifecycle utilize low-fidelity, high-level models of systems and their functionality. We currently lack the tools and processes
necessary for integrating ISHM into the vehicle system subsystem design. As a result, most existing ISHM-like technologies are retrofits that were done after the system
design was completed. It is very expensive, and sometimes futile, to retrofit a system health management capability into existing systems. Last-minute retrofits result in
unreliable systems, ineffective solutions, and excessive costs (e.g., Space Shuttle TPS monitoring which was considered only after 110 flights and the Columbia disaster).
High false alarm or false negative rates due to substandard implementations hurt the credibility of the ISHM discipline. This paper presents an overview of the current state of
ISHM design,and a review of formal design methods to make recommendations about possible approaches to enable the ISHM capabilities to be designed-in at the systemlevel, from the very beginning of the vehicle design process.
Low-Cost, Rapid Spacecraft Design and Multi-Subsystem Functionality
Information – Expert System
FC
Goddard Space Flight
Center
Title
Modular, Adaptive, Reconfigurable
Systems Technology for Sustainable,
Reliable, Effective, and Affordable
Space Exploration
Jet Propulsion
Laboratory
Repeat Ground Track Lunar Orbits in
the Full-Potential Plus Third-Body
Problem
Jet Propulsion
Laboratory
The NASA Exploration Design Team;
Blueprint for a New Design Paradigm
NTRS
Abstract
In order to execute the Vision for Space Exploration, we must find ways to reduce cost, system complexity, design, build, and test times, and at the same time increase
flexibility to satisfy multiple functions. Modular, Adaptive, Reconfigurable System (MARS) technologies promise to set the stage for the delivery of system elements that form
the building blocks of increasingly ambitious missions involving humans and robots. Today, space systems are largely specialized and built on a case-by-case basis. The
notion of modularity however, is nothing new to NASA. The 1970's saw the development of the Multi-Mission Modular spacecraft (MMS). From 1980 to 1992 at least six
satellites were built under this paradigm, and included such Goddard Space Flight Center missions as SSM, EUVE, UARS, and Landsat 4 and 5. Earlier versions consisted of
standard subsystem "module" or "box" components that could be replaced within a structure based on predefined form factors. Although the primary motivation for MMS was
faster cheaper integration and test, standardization of interfaces, and ease of incorporating new subsystem technology, it lacked the technology maturity and programmatic
"upgrade infrastructure" needed to satisfy varied mission requirements, and ultimately it lacked user buy-in. Consequently, it never evolved and was phased out. Such
concepts as the Rapid Spacecraft Development Office (RSDO) with its regularly updated catalogue of pre-qualified busses became the preferred method for acquiring
satellites. Notwithstanding, over the past 30 years since MMS inception, technology has advanced considerably and now modularity can be extended beyond the traditional
MMS module or box to cover levels of integration, from the chip, card, box, subsystem, to the space system and to the system-of-systems. This paper will present the MARS
architecture, cast within the historical context of MMS. Its application will be highlighted by comparing a state-of-the-art point design vs. a MARS-enabled lunar mission, as a
representative robotic case design. copyright 2005 American Institute of Physics.
A high degree and order Lunar gravitational field is superimposed on the Earth-Moon Restricted Three Body model to capture the dominating forces on a spacecraft in the
vicinity of the Moon. For the synchronously rotating Moon, periodic orbits in this model map repeat ground tracks and represent higher order solutions to the frozen orbit
problem. The near-circular, stable or near-stable solutions are found over a wide range of defining characteristics making them suitable for long-lifetime parking applications
such as science orbits, crew exploration vehicle parking orbits, and global coverage constellation orbits. A full ephemeris is considered for selected orbits to evaluate the
validity of the time-invariant, simplified model. Of the most promising results are the low-altitude families of near-circular, inclined orbits that maintain long-term stability despite
the highly non-spherical Lunar gravity. The method is systematic and enables rapid design and analysis of long-life orbits around any tidally-locked celestial body with an
arbitrarily high degree and order spherical harmonic gravity field.
NASA has chosen JPL to deliver a NASA-wide rapid-response real-time collaborative design team to perform rapid execution of program, system, mission, and technology
trade studies. This team will draw on the expertise of all NASA centers and external partners necessary. The NASA Exploration Design Team (NEDT) will be led by NASA
Headquarters, with field centers and partners added according to the needs of each study. Through real-time distributed collaboration we will effectively bring all NASA field
centers directly inside Headquarters. JPL's Team X pioneered the technique of real time collaborative design 8 years ago. Since its inception, Team X has performed over 600
mission studies and has reduced per-study cost by a factor of 5 and per-study duration by a factor of 10 compared to conventional design processes. The Team X concept
has spread to other NASA centers, industry, academia, and international partners. In this paper, we discuss the extension of the JPL Team X process to the NASA-wide
collaborative design team. We discuss the architecture for such a process and elaborate on the implementation challenges of this process. We further discuss our current
ideas on how to address these challenges.
Low-Cost, Rapid Spacecraft Design and Multi-Subsystem Functionality
Information – Software Tools for Distributed Analysis and Simulation
Company Name
Continuum Dynamics,
Inc.
Title
Serial In-Line
Instrumentation Bus for
ROV Engineering Research
Field Center
LARC
NASA SBIR Phase I
Quad Chart
Advanced microcontrollers having digital signal processing features have enabled the capability to distribute on-board computation for remotely operated vehicles
(ROVs). Distributed processing can result in a lighter weight avionics suite with improved performance, by locating data conversion units adjacent to the sensors and
control actuators, and reducing EMI through minimization of the amount of interconnection wiring. The proposed work will leverage CDI's and AMDI's substantial prior
experience in the development and operation of flight control avionics for ROVs in the design of a new system for supporting advanced research using these systems.
The avionics suite to be developed consists of serially interconnected distributed nodes that may be programmed through a Matlab graphical interface to perform control
and sensing functions in support of custom requirements from the research community. The flexibility of custom-configured distributed computing nodes for use in a
research context ensures that "just enough" instrumentation and control is provided for the specific test requirements at hand. Phase I will provide risk reduction by
demonstrating the operation of the subcomponent technologies, culminating in a simplified flight test of the avionics system. Phase II continuation will develop the
complete system to support testing
activities at a NASA research center of interest.
Low-Cost, Rapid Spacecraft Design and Multi-Subsystem Functionality
Information – Software Tools for Distributed Analysis and Simulation
Company Name
Design Net Engineering LLC
Title
LOW COST, TAILOR ABLE AVIONICS FOR
RAPID RESPONSE SATELLITES
Calabazas Creek Research,
Inc. and Old Dominion
University
FIELD MARSHALL SIMULATION
ENVIRONMENT
DOD SBIR Phase II
Quad Chart
Addressing the needs of the Responsive Space initiative requires a seamless integrated tool capability, particularly when applied to meet the most
challenging case of the “six day satellite” concept, where mission objectives are reduced to a satellite configuration and the spacecraft is assembled, tested
and flown within a week.
Simulation codes need to have user interface consistency for common functions and to use the research community’s vocabulary and units in specific
technical areas. Access to source code is of enormous benefit to research groups but not possible within the present technical software market. Along with
these “usability” issues, the codes must provide state-of-the-art simulation power.
Low-Cost, Rapid Spacecraft Design and Multi-Subsystem Functionality
Information – Software Tools for Distributed Analysis and Simulation
USPTO
Patent
Number
US6629065
Title
Assignee
Abstract
Methods and apparata
for rapid computeraided design of objects
in virtual reality and
other environments
Wisconsin Alumni
Research Foundation
Apparata and methods for rapid design of objects/shapes in Computer-Aided Design (CAD) tools and in Virtual Reality (VR) environments are
described. The underlying geometric representation of the objects within the design tool is optimized so that design activities such as modeling,
editing, rendering, etc. can be processed extremely rapidly, thereby enhancing the response time of the design tool. The representation is preferably
provided in two parts, which may be referred to as a "design intent model" and a "shape model". The design intent model is a higher-level
representation wherein elements are arranged in hierarchical parent-child relationships which record the elements' assembly sequence. The shape
model is a lower-level representation storing more detailed information about the elements and their relationships. During editing of the design, the
user acts on the design intent model, and the design intent model is mapped to the shape design model so that it is updated to reflect the changes
therein. The design intent model is in many cases sufficient by itself to allow basic editing of the model and rendering of the edited model, but where
editing operations grow sufficiently complex that the design intent model lacks sufficient information to allow the operation to be performed, the shape
model can be relied upon for the information necessary to complete the operation.
Low-Cost, Rapid Spacecraft Design and Multi-Subsystem Functionality
Information – Software Tools for Distributed Analysis and Simulation
U.S. Patent Applications
Patent Number
Title
US2006015299A1
Network architecture and
protocol for spacecraft systems
Assignee
None
Abstract
A network architecture and protocol provides "plug and play" spacecraft capabilities. A distributed-control architecture is used,
wherein each component operates semi-autonomously, and interacts with other components on a task/resource level. Each
component announces its requirement for system resources as a request to the network, and components that can provide some or all
of the requested resources respond to the request. An arbitration device centralizes and coordinates requests for critical and/or
singular resources, such as requests for a specific orientation of the spacecraft. To facilitate such distributed control, requests are
made in advance of the requirement for the resource, and include a time interval during which the resource is required. A
configuration and test system is provided to process the mission requirements and provide a set of components that can be
configured to satisfy the requirements.
Low-Cost, Rapid Spacecraft Design and Multi-Subsystem Functionality
Information – Software Tools for Distributed Analysis and Simulation
NTRS
Abstract
One aspect of collaborative engineering is facilitated through the use of parameter-exchange software. Collaborative design efforts usually require the definition of a set of shared
parameters that summarize the current state of design or that affect the overall design of the system. The Global Integrated Design Environment (GLIDE) facilitates easy passing and
sharing of parameters between engineers from remote locations. GLIDE uses a common Web-based MySQL server (MySQL AB, Uppsala, Sweden) that is accessed using PHP scripts
(the PHP Group). GLIDE s Web-served database allows secure and controlled access to design data by using firewall-friendly securesockets- layer- (SSL-) based user authentication.
Data can be queried and published to the GLIDE database via a Microsoft Excel (Microsoft Corporation, Redmond, WA) interface. The training of engineers on GLIDE goes quickly
because most users are familiar with Excel. Another benefit of using Excel as an interface is that optimization and other legacy tool integration developments like XLerator (ref. 1) can be
used. The following diagram shows the data flow and interfaces that enable parameter exchange within GLIDE.
FC
Glenn Research
Center
Title
Global Integrated Design
Environment (GLIDE) v.1.1
Software Released
Jet Propulsion
Laboratory
How to Quickly Import CAD
Geometry into Thermal Desktop
There are several groups at JPL (Jet Propulsion Laboratory) that are committed to concurrent design efforts, two are featured here. Center for Space Mission Architecture and Design
(CSMAD) enables the practical application of advanced process technologies in JPL's mission architecture process. Team I functions as an incubator for projects that are in the Discovery,
and even pre-Discovery proposal stages. JPL's concurrent design environment is to a large extent centered on the CAD (Computer Aided Design) file. During concurrent design sessions
CAD geometry is ported to other more specialized engineering design packages.
Goddard Space
Flight Center
Model-Based Systems
NASA (non Center
Specific)
System Level Uncertainty
Assessment for Collaborative
RLV Design
Engineers, who design systems using text specification documents, focus their work upon the completed system to meet Performance, time and budget goals. Consistency and integrity is
difficult to maintain within text documents for a single complex system and more difficult to maintain as several systems are combined into higher-level systems, are maintained over
decades, and evolve technically and in performance through updates. This system design approach frequently results in major changes during the system integration and test phase, and
in time and budget overruns. Engineers who build system specification documents within a model-based systems environment go a step further and aggregate all of the data. They
interrelate all of the data to insure consistency and integrity. After the model is constructed, the various system specification documents are prepared, all from the same database. The
consistency and integrity of the model is assured, therefore the consistency and integrity of the various specification documents is insured. This article attempts to define model-based
systems relative to such an environment. The intent is to expose the complexity of the enabling problem by outlining what is needed, why it is needed and how needs are being addressed
by international standards writing teams.
A collaborative design process utilizing Probabilistic Data Assessment (PDA) is showcased. Given the limitation of financial resources by both the government and industry, strategic
decision makers need more than just traditional point designs, they need to be aware of the likelihood of these future designs to meet their objectives. This uncertainty, an ever-present
character in the design process, can be embraced through a probabilistic design environment. A conceptual design process is presented that encapsulates the major engineering
disciplines for a Third Generation Reusable Launch Vehicle (RLV). Toolsets consist of aerospace industry standard tools in disciplines such as trajectory, propulsion, mass properties,
cost, operations, safety, and economics. Variations of the design process are presented that use different fidelities of tools. The disciplinary engineering models are used in a collaborative
engineering framework utilizing Phoenix Integration's ModelCenter and AnalysisServer environment. These tools allow the designer to join disparate models and simulations together in a
unified environment wherein each discipline can interact with any other discipline. The design process also uses probabilistic methods to generate the system level output metrics of
interest for a RLV conceptual design. The specific system being examined is the Advanced Concept Rocket Engine 92 (ACRE-92) RLV. Previous experience and knowledge (in terms of
input uncertainty distributions from experts and modeling and simulation codes) can be coupled with Monte Carlo processes to best predict the chances of program success.
Low-Cost, Rapid Spacecraft Design and Multi-Subsystem Functionality
Power and Energy – Power Management and Distribution
Company Name
AeroAstro, Inc.
Title
AUTONOMOUS SPACECRAFT POWER
SCHEDULING
Field Center
ARC
NASA SBIR Phase II
Quad Chart
Designed in collaboration with the Univ. of Colorado, the Autonomous Power Scheduler (APS) will dynamically monitor, schedule and
distribute power on microsatellites. The APS represents an innovative paradigm shift from conventional designs, controlling primary
power in an intelligent, semi- or fully autonomous way to maximize mission capability and performance. With APS, appropriate power
will be distributed when and where it is
most needed.
Low-Cost, Rapid Spacecraft Design and Multi-Subsystem Functionality
Power and Energy – Power Management and Distribution
Company Name
Design Net Engineering LLC
Title
Power Management for PnP
Spacecraft
DOD SBIR Phase I
Quad Chart
This research work will produce an innovative Power Management and Distribution (PMAD) system design that can support both heritage satellite components and
the new power compatible plug-and-play (PnP) methodology. The proposed PMAD system design will support a significant range of power requirements for
satellites with fewer power switches, fewer current monitors and simpler hierarchical management. In addition, a sophisticated simulation tool will be developed to
assist design engineers at the AFRL Testbed with component selection, confirm power system is sized correctly and mission requirements are met. Anticipated
benefits of PMAD system are reduced design time, simpler integration, rapid mission configuration, more autonomous operation, and reduced mission operations as
well as decreased cost and schedule. This study will define standards for both PMAD and PnP power components to assure compatibility during integration,
enhance mission and systems design tools, implement PnP power standards for xTEDS, develop a PMAD “smart” architecture design that supports the PnP
principle and the power management tools that allow autonomous on-orbit operation, as well as demonstrate utility and efficacy of those tools with “day in the life”
simulations and provide a mature design for a prototype flight-like device for Phase II efforts.
Low-Cost, Rapid Spacecraft Design and Multi-Subsystem Functionality
Propulsion – Chemical
FC
Marshall Space Flight Center
Title
NASA's In-Space Propulsion Technology
Program Overview and Status
NTRS
Abstract
NASA's In-Space Propulsion Technology Program is investing in technologies that have the potential to revolutionize the robotic exploration of deep space.
For robotic exploration and science missions, increased efficiencies of future propulsion systems are critical to reduce overall life-cycle costs and, in some
cases, enable missions previously considered impossible. Continued reliance on conventional chemical propulsion alone will not enable the robust
exploration of deep space - the maximum theoretical efficiencies have almost been reached and they are insufficient to meet needs for many ambitious
science missions currently being considered. The In-Space Propulsion Technology Program s technology portfolio includes many advanced propulsion
systems. From the next generation ion propulsion system operating in the 5 - 10 kW range, to advanced cryogenic propulsion, substantial advances in
spacecraft propulsion performance are anticipated. Some of the most promising technologies for achieving these goals use the environment of space itself
for energy and propulsion and are generically called, 'propellantless' because they do not require onboard fuel to achieve thrust. Propellantless propulsion
technologies include scientific innovations such as solar sails, electrodynamic and momentum transfer tethers, aeroassist, and aerocapture. This paper will
provide an overview of both propellantless and propellant-based advanced propulsion technologies, and NASA s plans for advancing them as part of the
60M per year In-Space Propulsion Technology Program.
Low-Cost, Rapid Spacecraft Design and Multi-Subsystem Functionality
Sensors and Sources – Sensor Webs/Distributed Sensors
Company Name
AeroAstro Corporation
Title
Wireless Data and Power Transfer on
Small Spacecraft
Field Center
ARC
NASA SBIR Phase I
Quad Chart
Achieving low-cost space missions implies lowering all phases of mission development, including spacecraft design, assembly, integration and
test. The concept of the wireless spacecraft bus is something most technical people are at least half familiar with - the half which includes
wireless data transfer, something available on every computer laptop today. But wireless is not really wireless if the power is delivered through
wires. Both power and data need to be delivered wirelessly for the true potential impact of wireless to be made on spacecraft design, build,
integration, and test. Integrating today's commonplace wireless data systems into spacecraft would seem to be a logical step in spacecraft
development, but to date has not been implemented widely if at all. AeroAstro proposes an innovative solution to design and build microspacecraft (and spacecraft components) harnessing the true promise of wireless systems. The overall objective of the proposal is to develop
and demonstrate a truly wireless spacecraft bus - exhibiting not only wireless data, but also wireless power distribution. By definition, this
wireless approach is inherently modular, and alleviates the need for wire harnesses of any type while simultaneously making staged built-in-test
possible concurrently during spacecraft assembly.
Low-Cost, Rapid Spacecraft Design and Multi-Subsystem Functionality
Structures – Launch and Flight Vehicles
Company Name
CSA Engineering, Inc
Title
Structural Attachments for Rapid Assembly of Satellites
INFOSCITEX CORPORATION
Multifunctional Programmable Structure (IPS) for a Modular
Satellite Architecture
DOD SBIR Phase I
Quad Chart
The objective of this study is to develop low cost, modular, multi-functional, primary spacecraft structures using a “plywood and two-byfour” approach for rapid assembly. These structures are comprised of aluminum facesheet/honeycomb-core sandwich panels
(plywood) bonded to aluminum extrusions (two-by-four). The structures require simple hand tools to assemble and leverage
nomograph based design guides for the required structural/thermal sizing for rapid response missions. The flat sandwich panels are
integrated into hexagonal (box) structures by using basic extruded shapes which are structural connections between panels. This
approach allows open architecture structural designs with scalable sizing. The bonded assemblies include fastening hard points for
attachment of flight sensor payloads (200kg class) and supporting subsystems. Since these structures are aluminum the basic
assembly provides inherent damage robustness; EMI and radiation hardened capability; and allows for multiple levels of thermal
management. These thermal management levels are: - The basic thermal continuity of the bonded aluminum panels and extrusions Easily added custom heat straps to augment thermal paths within the bonded the assembly or at attached components - Direct
interfacing to heat pipes at the mechanical hard points of the attached components. This study targets Air Force warfighter missions
with satellite build processing within 8 hours.
The DOD’s Operationally Responsive Space paradigm calls for ambitiously reducing the time required to design, fabricate and deploy a
satellite to days or even hours, from the current requirement of months or years. This timeline will require modular satellite structures
available on site that support plug-n-play integration with payload components. Infoscitex proposes a multi-functional structural panel
with programmable electronic and thermal management capabilities. A light-weight, carbon fiber composite core structure offers
structural integrity and acts to efficiently conduct heat away from thermally-controlled payloads. An innovative heat spreading
technology is applied over the core, allowing it to act as either an insulator or a heat sink via real-time electronic control. This feature
allows the proposed panel to provide appropriate thermal control to many payloads with disparate cooling requirements. The integrated
panel supports intra-panel wiring and component interface ports
compatible with emerging designs for a programmable wiring harness. During Phase I the project team will evaluate preliminary
concepts through modeling and analysis, optimize the design of an individual standardized panel, define inter-panel interfaces, and
demonstrate feasibility of a multi-panel structure. The Phase I objective will be to develop an optimum panel design that can be
prototyped and demonstrated during the Phase II.
Low-Cost, Rapid Spacecraft Design and Multi-Subsystem Functionality
Structures – Launch and Flight Vehicles
Company Name
MicroSat Systems, Inc.
Title
RAPID MANUFACTURING OF
MULTI-FUNCTIONAL SATELLITE
PANELS WITH EMBEDDED
COMPONENTS USING
ULTRASONIC CONSOLIDATION
Physical Sciences, Inc.
Physical Sciences, Inc.
DOD SBIR Phase II
Quad Chart
This Phase II effort focuses on the development of an ultrasonic consolidation (UC) process for creating AI 6061 alloy components with embedded functionality
that will meet the spacecraft mechanical, thermal and electrical requirements of the Microsatellite Target System and the Plug and Play Satellite. The result is a
validated process for rapidly manufacturing multi-functional satellite panels, built at lower production cost, with less labor.
A modular, reconfigurable small spacecraft that can be assembled within 7 days and integrated with payloads within 24 hours. Our approach would allow
spacecraft configuration on-line from modules selected from a PC-inspired inventory scheme and rapid integration and test in the field via web-based real time
technical support from the manufacturer.
Low-Cost, Rapid Spacecraft Design and Multi-Subsystem Functionality
Structures – Launch and Flight Vehicles
USPTO
Patent Number
Title
Assignee
Abstract
US6543724
Miniature Satellite Design
Lockheed Martin Corporation
The invention is a micro-satellite assembly. In detail, the invention includes first and second flat structural members containing the satellite
payload. First and second tubular elements connect first and second structural members such that they are in a spaced relationship. A
plurality of solar panels are movably to the tubular elements between the first and second structural elements, movable from a stored
position between the structural elements to an deployed position external of these structural members. A mechanism is provided for biasing
the plurality of the solar panels to the deployed position. A second mechanism is used to releasably secure the plurality of solar panels in
the stored position.
Low-Cost, Rapid Spacecraft Design and Multi-Subsystem Functionality
Structures – Modular Interconnects
U.S. Patent Applications
Patent Number
Title
Assignee
Abstract
US2003150958A1
Miniature Spacecraft
None
A miniature spacecraft is constructed using modular features so as to provide a wide range of possible sizes with a choice of physical attributes.
Radially directed bays are arranged around a central cylindrical element which may be used for storing a propulsion tank or canister. Octagonal or
other configurations provide rigidity and strength through the use of triangular planetary cylinders which form outwardly facing bays for storage.
US2007029446A1
Modular platform architecture
for satellites
None
A method for implementing a modular platform for the construction of satellites and other spacecraft based on modular platform architecture, the
method comprising: (a) identifying a plurality of functional elements and their associated functional routines that may be operable within at least
one satellite; (b) associating the functional routines with one another in a strategic manner; (c) dividing the functional routines to define a plurality
of subsystems; and (d) deriving a plurality of modules from the plurality of subsystems, each of the modules being configured to operably interface
with at least one other module to construct a working satellite capable of carrying out a pre-determined number of the functional routines.
Low-Cost, Rapid Spacecraft Design and Multi-Subsystem Functionality
Structures – Structural Modeling and Tools
U.S. Patent Applications
Patent Number
Title
Assignee
Abstract
US2006016935A1
Modular Spacecraft Design
Architecture
AeroAstro, Inc.
A spacecraft architecture and accompanying standard allows for the creation of a spacecraft using an assortment of
modules that comply with the standard. The standard preferably includes both mechanical and electrical compatibility
criteria. To assure physical/mechanical compatibility, the structure of each module is constrained to be compatible with
any other compatible module. To minimize the interference among modules, the extent of each module in select
dimensions is also constrained. To assure functional compatibility, a common communication format is used to interface
with each module, and each public-function module is configured to respond to requests for function capabilities that it
can provide to other functions. Each module is preferably designed to provide structural support to the assemblage of
modules, and an anchor module is provided or defined for supporting the entire assemblage and coupling the assemblage
to other structures, such as a launch vehicle.
Low-Cost, Rapid Spacecraft Design and Multi-Subsystem Functionality
Structures – Structural Modeling and Tools
USPTO
Patent Number
Title
US7110842
Modular design
method
Assignee
PACCAR Inc.
Abstract
A method of modular design suitable for use, for example, in a CAD program is provided. In one aspect, the method is embodied in a CAD program
specifically configured to support modular design, or more specifically, concurrent and consistent design of multiple installation assemblies of a product.
Each installation assembly consists of a reference assembly and a Bill-of-Material (BOM) assembly. The BOM assembly contains component models
that are the modular elements to be designed. The reference assembly contains referenced assemblies that are installed to make up the design
environment for the component models in the BOM assembly. The BOM assembly in one installation assembly can be reused as a referenced assembly
in the reference assembly of another installation assembly. Any changes made to the BOM assembly are automatically reflected in the reference
assembly that contains the same BOM assembly as a referenced assembly.
Low-Cost, Rapid Spacecraft Design and Multi-Subsystem Functionality
Verification and Vaildation – Operations Concepts and Requirements
NTRS
Abstract
Modular, Reconfigurable, and Rapid-response (MR(sup 2)) space systems represent a paradigm shift in the way space assets of all sizes are designed, manufactured, integrated, tested,
and flown. This paper will describe the MR(sup 2) paradigm in detail, and will include guidelines for its implementation. The Remote Sensing Advanced Technology microsatellite (RSAT) is
a proposed flight system test-bed used for developing and implementing principles and best practices for MR(sup 2) spacecraft, and their supporting infrastructure. The initial goal of this
test-bed application is to produce a lightweight (approx. 100 kg), production-minded, cost-effective, and scalable remote sensing micro-satellite capable of high performance and broad
applicability. Such applications range from future distributed space systems, to sensor-webs, and rapid-response satellite systems. Architectures will be explored that strike a balance
between modularity and integration while preserving the MR(sup 2) paradigm. Modularity versus integration has always been a point of contention when approaching a design whereas oneof-a-kind missions may require close integration resulting in performance optimization, multiple and flexible application spacecraft benefit and ampom modularity, resulting in maximum
flexibility. The process of building spacecraft rapidly ( and lt 7 days), requires a concerted and methodical look at system integration and test processes and pitfalls. Although the concept of
modularity is not new and was first developed in the 1970s by NASA's Goddard Space Flight Center (Multi-Mission Modular Spacecraft), it was never modernized and was eventually
abandoned. Such concepts as the Rapid Spacecraft Development Office (RSDO) became the preferred method for acquiring satellites. Notwithstanding, over the past 30 years technology
has advanced considerably, and the time is ripe to reconsider modularity in its own right, as enabler of R(sup 2), and as a key element of transformational systems. The MR2 architecture
provides a competitive advantage over the old modular approach in its rapid response to market needs that are difficult to predict both from the perspectives of evolving technology, as well
as mission and application requirements.
The days of watching a massive manned cylinder thrust spectacularly off a platform into space might rapidly become ancient history when the National Aeronautics and Space
Administration (NASA) introduces its new millenium mission class. Motivated by the need to gather more data than is possible with a single spacecraft, scientists have developed a new
class of missions based on the efficiency and cooperative nature of a hive culture. The missions, aptly dubbed nanoswarm will be little more than mechanized colonies cooperating in their
exploration of the solar system. Each swarm mission can have hundreds or even thousands of cooperating intelligent spacecraft that work in teams. The spacecraft must operate
independently for long periods both in teams and individually, as well as have autonomic properties - self-healing, -configuring, -optimizing, and -protecting- to survive the harsh space
environment. One swarm mission under concept development for 2020 to 2030 is the Autonomous Nano Technology Swarm (ANTS), in which a thousand picospacecraft, each weighing
less than three pounds, will work cooperatively to explore the asteroid belt. Some spacecraft will form teams to catalog asteroid properties, such as mass, density, morphology, and chemical
composition, using their respective miniature scientific instruments. Others will communicate with the data gatherers and send updates to mission elements on Earth. For software and
systems development, this is uncharted territory that calls for revolutionary techniques.
The Rapid Spacecraft Development Office (RSDO) at NASA's Goddard Space Flight Center is responsible for the management and direction of a dynamic and versatile program for the
definition, competition, and acquisition of multiple indefinite delivery and indefinite quantity contracts - resulting in a catalog of spacecraft buses. Five spacecraft delivery orders have been
placed by the RSDO and one spacecraft has been launched. Numerous concept and design studies have been performed, most with the intent of leading to a future spacecraft acquisition.
A collection of results and lessons learned is recorded to highlight management techniques, methods and processes employed in the conduct of spacecraft acquisition. Topics include
working relationships under fixed price delivery orders, price and value, risk management, contingency reserves, and information restrictions.
FC
GSFC
Title
Modular, Reconfigurable, and
Rapid Response Space Systems
The Remote Sensing Advanced
Technology Microsatellite
GSFC
NASA's Swarm Missions The
Challenge of Building
Autonomous Software
GSFC
Rapid Spacecraft Development
Results and Lessons Learned
SSC
Small Satellite Constellations:
The Future for Operational Earth
Observation
Nanosat, microsat and minisat are low-cost, rapid-response small-satellites built from advanced terrestrial technology. SSTL delivers the benefits of affordable access to space through lowcost, rapid response, small satellites designed and built with state-of-the-art COTS technologies by: a) reducing the cost of entry into space; b) Achieving more missions within fixed
budgets; c) making constellations and formation flying financially viable; d) responding rapidly from initial concept to orbital operation; and e) bringing the latest industrial COTS component
advances to space. Growth has been stimulated in constellations for high temporal revisit and amp;persistent monitoring and military responsive space assets.
Jet
Propulsion
Laboratory
An Integrated Approach to Risk
Assessment for Concurrent
Design
This paper describes an approach to risk assessment and analysis suited to the early phase, concurrent design of a space mission. The approach integrates an agile, multi-user risk
collection tool, a more in-depth risk analysis tool, and repositories of risk information. A JPL developed tool, named RAP, is used for collecting expert opinions about risk from designers
involved in the concurrent design of a space mission. Another in-house developed risk assessment tool, named DDP, is used for the analysis.
Goddard
Space Flight
Center
Analysis of Formation Flying in
Eccentric Orbits Using Linearized
Equations of Relative Motion
Geometrical methods for formation flying design based on the analytical solution to Hill's equations have been previously developed and used to specify desired relative motions in near
circular orbits. By generating relationships between the vehicles that are intuitive, these approaches offer valuable insight into the relative motion and allow for the rapid design of satellite
configurations to achieve mission specific requirements, such as vehicle separation at perigee or apogee, minimum separation, or a specific geometrical shape. Furthermore, the results
obtained using geometrical approaches can be used to better constrain numerical optimization methods allowing those methods to converge to optimal satellite configurations faster. This
paper presents a set of geometrical relationships for formations in eccentric orbits, where Hill's equations are not valid, and shows how these relationships can be used to investigate
formation designs and how they evolve with time.
Low-Cost, Rapid Spacecraft Design and Multi-Subsystem Functionality
Verification and Vaildation – Operations Concepts and Requirements
FC
Jet
Propulsion
Laboratory
Marshall
Space Flight
Center
Title
JIMO Follow-On Mission Studies
NTRS
Abstract
Team Prometheus is a geographically-distributed, collaborative engineering team composed of representatives from several NASA centers and the Department of Energy (DOE). During the
months of April through September 2004, Team Prometheus performed studies of representative Saturn Titan and Neptune Triton science missions based on proposed Jupiter Icy Moons
Orbiter (JIMO) technology. The principal objectives of these studies were 1) to assess the feasibility of using direct copies of the baseline JIMO flight system designs to perform these followon missions, and 2) to identify and assess technologies or potential enhancements or alterations to the baseline JIMO designs which could reduce the cruise durations to meet NASA
Headquarters-specified programmatic goals. A tertiary objective was to provide feedback to the JIMO Project on the suitability of the JIMO reference designs for potential follow-on
applications. copyright 2005 American Institute of Physics.
Low-Cost, Rapid Spacecraft Design and Multi-Subsystem Functionality
Verification and Vaildation – Simulation Modeling Enviornment
Company Name
NextGen
Aeronautics, Inc.
Title
Advanced Modeling Concepts
for Conceptual Design
Field Center
ARC
NASA SBIR Phase I
Quad Chart
Preliminary design of aircraft structures is multidisciplinary, involving knowledge of structural mechanics, aerodynamics, aeroelasticity, structural dynamics and design
concepts. Even though analysis tools and computing resources have improved significantly, very little attention is being paid to the required data exchange and
seamless interoperability for true multi-disciplinary analysis. Indeed there is a lot to be gained in performing multi-disciplinary analysis up front in the design cycle
during the conceptual design phase as changes during mid or tail end of the design process are often expensive and difficult to implement. The net result is increased
cost and performance limiting weight penalties on the air vehicle. It is essential to consider the impact of several design disciplines simultaneously to arrive at a
satisfactory design where generated data for each discipline is seamlessly generated or available to each design discipline.
Low-Cost, Rapid Spacecraft Design and Multi-Subsystem Functionality
Verification and Vaildation – Simulation Modeling Enviornment
Company Name
Lynguent Inc.
Title
MODEL-BASED DESIGN TOOLS FOR
EXTENDING COTS COMPONENTS TO
EXTREME ENVIRONMENTS
Field Center
JPL
NASA SBIR Phase II
Quad Chart
These MBD tools will enable lower-cost system development and cost versus lifetime assessment, shorten development
time, and extend flight-proven technology to broader applications. Plans to develop MBD tools based on its Phase I
feasibility study and to utilize a high temperature test bed as a case study to demonstrate a calibration methodology for the
tools to insure accuracy with respect to accelerated testing results.
Low-Cost, Rapid Spacecraft Design and Multi-Subsystem Functionality
Verification and Vaildation – Simulation Modeling Enviornment
Company Name
Advatech Pacific, Inc.
Title
Space Propulsion Modeling and Simulation
DOD SBIR Phase I
Quad Chart
There is a requirement for improving specific impulse and mass fraction in propulsion systems, while providing precision and predictability in
thrust to enhance accuracy. These systems must be dependable and cost-effective at a high level of operational readiness. This contributes
significantly to performance of systems such as Land Based Strategic Deterrent, Prompt Global Strike and Responsive Launch to Space and
Responsive Satellites. Advatech Pacific (API) proposes to provide modeling and simulation of critical system elements (utilizing legitimate
mathematical, aeronautical, astronautical, operational, and chemical principles, ultimately to encompass life cycle cost) in electric, chemical,
and high-energy density propulsion, which will be used with laboratory research to advance the state of the art. API, working with AFRL, will
create an architecture of software systems and integrate these to provide comprehensive simulation and modeling tools. Electric propulsion
models (Coliseum, HPHall, Colloid) have state of the art physics, but have a basic architecture, which provides limited functionality. API is
suited to the task of re-architecting and parallelizing these models. API will also support chemical propulsion technology including CFD
modeling of several key experimental systems. Finally, API will develop plasma physics and high-energy propulsion tool concepts in support
of a joint PRSA/PRSS effort at AFRL.
Low-Cost, Rapid Spacecraft Design and Multi-Subsystem Functionality
Verification and Vaildation – Simulation Modeling Enviornment
Company Name
Star Technologies Corporation
Title
3D SATELLITE BUILDER
DOD SBIR Phase II
Quad Chart
Star Technologies Corporation proposes to develop "3D Satellite Builder" that provides (1) 3D Visualization User Interface (UI) to
support spacecraft design, development, and testing; (2) a flexible Plug-n-Play software architecture for prototyping of subsystems.
Low-Cost, Rapid Spacecraft Design and Multi-Subsystem Functionality
Verification and Vaildation – Simulation Modeling Enviornment
U.S. Patent Applications
Patent Number
Title
Assignee
Abstract
US2005028133A1
System and method for rapid
design, prototyping, and
implementation of distributed
scalable architecture for task
control and automation
None
US2007244678A1
Design Optimization System and
Method
Board of Trustees of
Michigan State University
The present invention provides a system and method for simplifying and accelerating the process of prototyping, real-world
simulation, and implementation of virtually any task performance system or device, thereby dramatically reducing the design-toimplementation cycle time and expense. The inventive system includes a development system that provides a user, with visual tools
to interactively and dynamically partition a previously designed visual system model of the task performance system or device, and
then interactively or automatically assign the partitions to corresponding selectable target components, to produce a prototyped
system ready for conversion to executable form suitable for implementation. The inventive system and method can also be readily
used to automatically generate any instruction sets that are necessary for implementing the prototyped task performance system in
actual target components of one or more emulation and/or production target systems. A novel automatic executable program code
generation process that can be advantageously utilized is also provided in accordance with the present invention. Finally, the present
invention may optionally include a data handling device that enables real-time monitoring and management of a remote target
system from one or more user systems, as well as a set of tools for designing interactive visual instrument panels for that purpose.
A design optimization system and method includes recursive evaluation of a system model and/or a set of subsystem models using
subsystem design(s) to extract interactions and perform subsystem design optimization. Selective update of the boundary conditions
occurs according to: <maths id="MATH-US-00001" num="1"> <MATH OVERFLOW="SCROLL"> <MROW> <MSUB>
<MOVER> <MI>u</MI> <MO>^</MO> </MOVER> <MROW> <MI>subsystem</MI> <MO>b</MO> <MSTYLE> <MTEXT>
</MTEXT> </MSTYLE> <MO>b</MO> <MROW> <MO>(</MO> <MI>i</MI> <MO>)</MO> </MROW> </MROW>
</MSUB> <MO>=</MO> <MROW> <MUNDEROVER> <MO>∑</MO> <MROW> <MI>s</MI> <MO>=</MO>
<MN>1</MN> </MROW> <MI>k</MI> </MUNDEROVER> <MO>b</MO> <MROW> <MSUB> <MOVER> <MI>w</MI>
<MO>~</MO> </MOVER> <MI>s</MI> </MSUB> <MO>b</MO> <MSUB> <MOVER> <MI>u</MI> <MO>^</MO>
</MOVER> <MI>s</MI> </MSUB> <MO>b</MO> <MSTYLE> <MTEXT> </MTEXT> </MSTYLE> <MO>b</MO>
<MI>on</MI> <MO>b</MO> <MSTYLE> <MTEXT> </MTEXT> </MSTYLE> <MO>b</MO> <MSUB> <MI>Gamma</MI>
<MROW> <MI>subsystem</MI> <MO>b</MO> <MSTYLE> <MTEXT> </MTEXT> </MSTYLE> <MO>b</MO> <MROW>
<MO>(</MO> <MI>i</MI> <MO>)</MO> </MROW> </MROW> </MSUB> </MROW> </MROW> </MROW> </MATH>
</MATHS> where û<SUB>s</SUB>(x,y,z,t) are the interactions between subsystems and/or between the system and its
subsystem(s) on Gamma<SUB>subsystem(i) </SUB>at iteration s, and w<SUB>s</SUB>(x,y,z,t) are weight functions whose
spatial and temporal distributions are predetermined and whose magnitudes are varied stochastically within a selected range.
Low-Cost, Rapid Spacecraft Design and Multi-Subsystem Functionality
Verification and Vaildation – Simulation Modeling Enviornment
NTRS
Abstract
A multidisciplinary, collaborative simulation has been performed on a Grid of geographically distributed PC clusters. The multiscale simulation approach seamlessly combines
i) atomistic simulation backed on the molecular dynamics (MD) method and ii) quantum mechanical (QM) calculation based on the density functional theory (DFT), so that
accurate but less scalable computations are performed only where they are needed. The multiscale MD QM simulation code has been Grid-enabled using i) a modular,
additive hybridization scheme, ii) multiple QM clustering, and iii) computation communication overlapping. The Gridified MD QM simulation code has been used to study
environmental effects of water molecules on fracture in silicon. A preliminary run of the code has achieved a parallel efficiency of 94 on 25 PCs distributed over 3 PC clusters
in the US and Japan, and a larger test involving 154 processors on 5 distributed PC clusters is in progress.
FC
Ames Research
Center
Title
Collaborative Simulation Grid Multiscale
Quantum-Mechanical Classical
Atomistic Simulations on Distributed PC
Clusters in the US and Japan
Langley Research
Center
Development of a Dynamically
Configurable, Object-Oriented
Framework for Distributed, Multi-modal
Computational Aerospace Systems
Simulation
Goddard Space Flight
Center
Optics Program Simplifies Analysis and
Design
Stennis Space Center
Promising More Information
When NASA needed a real-time, online database system capable of tracking documentation changes in its propulsion test facilities, engineers at Stennis Space Center joined
with ECT International, of Brookfield, Wisconsin, to create a solution. Through NASA's Dual-Use Program, ECT developed Exdata, a software program that works within the
company's existing Promise software. Exdata not only satisfied NASA s requirements, but also expanded ECT s commercial product line. Promise, ECT s primary product, is
an intelligent software program with specialized functions for designing and documenting electrical control systems. An addon to AutoCAD software, Promise generates
control system schematics, panel layouts, bills of material, wire lists, and terminal plans. The drawing functions include symbol libraries, macros, and automatic line breaking.
Primary Promise customers include manufacturing companies, utilities, and other organizations with complex processes to control.
Jet Propulsion
Laboratory
SCENAREO - Science and engineering
activity reasoning and optimization
SCENAREO is a PC-based multi-mission scenario design tool for space missions based on inputs of subsystem performance models, information on various targets, and the
specifications for the science objectives. This tool provides model-based analysis to compose and schedule subsystems' operations automatically and optimally to meet the
needs of the science objectives. Moreover, SCENAREO has analytical capabilities to validate the feasibility of the science objectives based on subsystem design parameters
and to recommend the valid setting in terms of subsystem performance and resources usages to fulfill the proposed science objectives. This tool's automation, optimization,
and accuracy capabilities reduce the mission design lifecycle.
JPL
Model-based engineering design for
space missions
The basic elements of model-based design for space missions have existed for almost a decade, awaiting an opportunity to implement them in the same place at the same
time. In early design phases, combinations of models, concurrent engineering methods, and scenario-driven design have been used for several years with results that have
exceeded even optimistic expectations but the goal of extending these methods to later phases of design has been more elusive. JPL's Model-Based Engineering Design
(MBED) initiative will provide opportunity to reach that goal. It enables advanced systems engineering practice through a series of integrated, increasingly detailed models that
provide continuity from architectural concept through detailed design. It extends current capability for rapid conceptual design, allowing thorough exploration of design
tradespaces and selection of an optimal design point with associated cost and rationale and it provides seamless connection to subsystem models and detailed design tool
suites. In this paper we will review the goals and status of MBED and show the expected interconnectivity between conceptual and detailed design.
This research is aimed at developing a neiv and advanced simulation framework that will significantly improve the overall efficiency of aerospace systems design and
development. This objective will be accomplished through an innovative integration of object-oriented and Web-based technologies with both new and proven simulation
methodologies. The basic approach involves Ihree major areas of research Aerospace system and component representation using a hierarchical object-oriented component
model which enables the use of multimodels and enforces component interoperability. Collaborative software environment that streamlines the process of developing, sharing
and integrating aerospace design and analysis models. . Development of a distributed infrastructure which enables Web-based exchange of models to simplify the
collaborative design process, and to support computationally intensive aerospace design and analysis processes. Research for the first year dealt with the design of the basic
architecture and supporting infrastructure, an initial implementation of that design, and a demonstration of its application to an example aircraft engine system simulation.
Engineers at Goddard Space Flight Center partnered with software experts at Mide Technology Corporation, of Medford, Massachusetts, through a Small Business Innovation
Research (SBIR) contract to design the Disturbance-Optics-Controls-Structures (DOCS) Toolbox, a software suite for performing integrated modeling for multidisciplinary
analysis and design. The DOCS Toolbox integrates various discipline models into a coupled process math model that can then predict system performance as a function of
subsystem design parameters. The system can be optimized for performance; design parameters can be traded; parameter uncertainties can be propagated through the math
model to develop error bounds on system predictions; and the model can be updated, based on component, subsystem, or system level data. The Toolbox also allows the
definition of process parameters as explicit functions of the coupled model and includes a number of functions that analyze the coupled system model and provide for
redesign. The product is being sold commercially by Nightsky Systems Inc., of Raleigh, North Carolina, a spinoff company that was formed by Mide specifically to market the
DOCS Toolbox. Commercial applications include use by any contractors developing large space-based optical systems, including Lockheed Martin Corporation, The Boeing
Company, and Northrup Grumman Corporation, as well as companies providing technical audit services, like General Dynamics Corporation.
Low-Cost, Rapid Spacecraft Design and Multi-Subsystem Functionality
Verification and Vaildation – Testing Requirements and Architectures
Company Name
GMA Industries, Inc.
Title
Automated Test Program Set Development using Integrated
Circuit Electromagnetic Emissions
DOD SBIR Phase I
Quad Chart
This proposal describes the automated development of a test program set utilizing electromagnetic emissions from integrated circuits to
determine UUT operational status. Electromagnetic emissions from rapidly changing voltages and currents within high-speed logic and other
circuits have traditionally been seen as a problematic source of electromagnetic interference that must be eliminated as much as possible.
However, changes within these emissions can be a highly significant indicator that a failure has occurred within the IC as well as within
adjacent circuit card paths. Further, the hardware and software requirements to capture and analyze these emissions are relatively simple in
comparison to today's highly complex automatic test equipment, however this approach was previously not possible until the recent
development of comprehensive analysis capability within common oscilloscopes and spectrum analyzers. Our approach requires basic power
supplies, logic stimulus, programmable oscilloscope/spectrum analyzer, and simple probes with a test fixture having no active circuitry, as
opposed to the large number of stimulus and response equipment commonly found in ATE. We present an approach towards capturing
electromagnetic emission signatures through automated means, and ascribing failures to the circuit board as a whole, and to the individual
circuit board components.
Project Management, Systems, Engineering and Mission Assurance Tools
Communication – Architectures and Networks
Company Name
ICs
Title
Flight Lossless Data Compression
Electronics
Field Center
GSFC
NASA SBIR Phase I
Quad Chart
The proposed work seeks to drastically increase the capability of the lossless data compression technology embedded in the currently used flight
part known as USES (Universal Source Encoder for Space). USES met the CCSDS 121-0-B 1 recommendation. New advances to the lossless data
compression electronic technology which advances the current flight electronics device: ? Increase quantization levels to 32 bits; the current device
supports only 15 bits. ? Support multi-frequency simultaneous inputs, at least three to represent color inputs. ? Increase speed from 20
MSamples/sec to 200 M Samples/sec ? Realize in a radiation tolerant 0.25 micron CMOS process
Project Management, Systems, Engineering and Mission Assurance Tools
Communication – Architectures and Networks
U.S. Patent Apps
Patent Number
Title
Assignee
Abstract
US20030212749A1
Method and system
for secure electronic
distribution,
archiving and
retrieval
GE Financial Assurance
Holdings, Inc
The present invention provides a secure electronic information distribution, archiving and retrieval method and system for communication
and correspondence purposes using an integrated system of email, email links and secured online web sites. An information bulletin may be
maintained in a secure, easy to use searchable archive. A bulletin notification may be used to send information bulletin summaries and links
on a regular or predetermined basis, or on an ad hoc basis, to inform and/or alert one or more recipients of receipt of the information
bulletin. The information bulletin may be accessed through a direct link, which may be randomly generated to preserve security and
authenticity of the information bulletin. The information bulletin may further include links to attachments and other interactive information.
US20060085238A1
Method and system
for monitoring an
issue
Oden Insurance Services, Inc.,
The present invention is a web-based software application designed to monitor implementation of an issue. The invention includes
receiving information from a user on the issue to be monitored; receiving information from a user on an assignment on the issue; and
sending the assignment to a recipient. An acknowledgement is received from the recipient on the assignment. Reminders are forwarded to
the recipient when the acknowledgement is not received within a predetermined time period, and the reminders can be automatically
forwarded at predetermined time periods. Information is received on the issues from a user or an external source, and listed in a form which
can be filtered and sorted according to predetermined criteria. A completion status of the regulatory issue can be displayed.
US20070011202A1
System and method
for electronic
message notification
None
US20060048050A1
Method for providing
both automated and
on demand project
performance
measurements
International Business Machines
Corporation
A method, article of manufacture and system for electronic mail notification including determining that it is a time for at least one of a
reminder notification action and a report notification action. For a reminder notification action, all action items within a given range are
obtained and the action items are processed into at least one action item report which is e-mailed to at least one designated recipient. For a
report notification action, the a list of completed action items in a given range are obtained and the list of completed action items is
processed into at least one completed action report that is e-mailed to at least one designated recipient.
A method and system for displaying and reporting project completion information of a large-scale project having a number of individual
subprojects. A rollup agent is employed for acquiring project completion information from a number of subproject databases to provide
overall project-wide display and reporting capabilities. The rollup agent collects and organizes the information, and stores it on a rollup
database. The rollup agent also interacts with a user to select aspects of any subprojects to report or display. The rollup agent then creates
and stores reports of the selected information, or displays the report interactively as a pop-up summary report or a spreadsheet format report
to a user. A number of scheduled agents provide scheduled reporting at the subproject level. A number of on demand agents provides the
interactive database view, pop-up reporting and spreadsheet reporting for each of the individual subprojects.
Project Management, Systems, Engineering and Mission Assurance Tools
Communication – Architectures and Networks
U.S. Patent Apps
Patent Number
Title
Assignee
Abstract
US2006293939A1
Design Managing
Means, Design Tool
and Method for Work
Breakdown Structure
None
US20070078792A1
One view integrated
project management
system
4 U Services DBA Stellar
Services
A WBS design managing means in project management, comprising: a WBS tree-like architecture designing means for implementing
graphical tree-like design of the work items included in a project and their relations, and mapping them into an enhanced WBS data
structure; a WBS attribute editing means for defining and editing attributes of work items at each node of the tree-like architecture; and a
WBS data managing means for storing and managing data constructed in accordance with said enhanced WBS data structure. The WBS
design managing means may connect to a converter for converting said constructed data into the format required by a project management
tool so as to input said structured data to the project management tool. Besides, there is provided an enhanced WBS design tool. The
present invention makes the system design tool and project management tool be integrated, so that the system architect and the project
management personal have more fluent channel for information exchange under the support of the present invention's tool, improving
efficiency and accuracy of project management.
Projects of any type require sophisticated management software programs. In reality, these management software programs are provided
by various vendors and for different professional fields. For example, scheduling programs for construction professionals, blue print
programs for drafting professionals, accounting programs for cost-control professionals. In addition, there are numerous other unofficial
and official documents generated by managers of different levels for reporting or track-record-keeping purposes using popular word
processing programs and spread-sheet programs. Documents generated by different software programs can only be viewed under the
programs they are generated under or compatible programs. This poses a problem in that there is a lack of organized control resulting in
difficulty in locating and viewing documents in a timely manner. The present invention provides an one view software program that would
be able to provide an organized control and viewing of all documents regardless which software programs they are generated under.
Project Management, Systems, Engineering and Mission Assurance Tools
Information – Computer Systems Architectures
Company Name
Advanced Solutions, Inc
Title
Modular Autonomous C&DH
Software with Built-In
Simulation/Test
Field Center
ARC
NASA SBIR Phase I
Quad Chart
NASA, the Department of Defense (DoD), and Commercial spacecraft programs have the desire to minimize program costs while continuing to ensure
mission robustness and flexibility. The spacecraft system that is a cost driver throughout the program life cycle is the Command and Data Handling
System (C&DH), along with the associated algorithms and software. Advanced Solutions Inc (ASI) plans to develop C&DH software which can be
targeted and adapted to a wide variety of C&DH hardware architectures and mission requirements. We also recognize the need to streamline the
entire spacecraft development lifecycle and provide a product that not only provides highly autonomous core flight software that is adaptable to any
mission, but also has the ability to replace traditional development, integration and test elements. This will be accomplished by expanding upon our
revolutionary On-board Dynamic Simulation System (ODySSy) to allow the C&DH system to support all phases of the spacecraft lifecycle. Additionally,
the traditional test control room is now unnecessary and is replaced by the mission control architecture to provide a true test-like-you-fly environment.
The test team, mission control team, and data analyst's can be in remote locations through use of the Web-Based
Data Distribution Network (WebDDN).
Project Management, Systems, Engineering and Mission Assurance Tools
Information – Computer Systems Architectures
Company Name
Datatek Applications, Inc.
Title
Mobile IPv6 in a Low Bandwidth Tactical Environment
DOD SBIR Phase I
Quad Chart
The purpose of this proposal is to demonstrate the feasiblity of innovative extentions to the use of Mobile IPv6 to allow US Army legacy tactical
devices to interoperate in the new IPv6 mobile networks being deployed. The Army is planning the migration of its core network to Internet
Protocol version 6 (IPv6) over the next several years as part of its Global Information Grid (GIG) expansion. The Army has many legacy tactical
IPv4-based mobile devices, such as radios, data and navigation units, etc. that may never be replaced or updated for many years or may be costprohibitive to do so. These IPv4-based tactical devices are not interoperable with the IPv6-based GIG. Our company, Datatek Applications, is
developing an experimental IPv4-IPv6 Translator that shows promise in solving this interoperability problem facing the Army and DoD. Mobility is
of paramount importance to the DoD. We plan on continuing to develop our Translator for use in Army Net Centric Warfare Command and Control
(C2) Systems to rapidly convert these IPv4 legacy devices to Mobile IPv6 nodes with all the benefits that IPv6 brings to the mobile warfighter.
Project Management, Systems, Engineering and Mission Assurance Tools
Information – Computer Systems Architectures
Company Name
Trident Systems, Inc.
Title
INTEGRATED SIMULATION-BASED DESIGN
ENVIRONMENT
Modus Operandi, Inc.
CODA: A COLLABORATIVE DATA ACCESS SYSTEM
DOD SBIR Phase II
Quad Chart
This project is attempting to realize a prototype of an environment for the design and optimization of undersea weapons and vehicles.
InterchangeSE is a distributed, multi-user infrastructure that integrates tools such as systems engineering, design optimization, network
simulation, requirements, CAD, cost, human engineering, and project management tools.
Systems integration problems are magnified on large DoD projects built by distributed multi-contractor teams using heterogeneous
information systems. The challenge is to rapidly access and exploit these information systems while managing costs and risks.
Project Management, Systems, Engineering and Mission Assurance Tools
Information – Data Acquisition and End-to-End Management
Company Name
Odyssey Space Research
Title
Wiki-based Data and Information Integration
(WikiDI2) System
Field Center
ARC
NASA SBIR Phase I
Quad Chart
The proposed innovation is: A data and information integration (DI2) system built from the methods and tools used to create Wiki
websites. Wiki (Hawaiian for "quick") is software that allows users to create and edit Web page content using any Web browser.
The significance of the innovation is that a Wiki-based DI2 system will: (1) Produce a collaborative, interactive website designed
for the unique needs of government and commercial spacecraft projects for capturing, disseminating, managing and linking data
resources across multiple projects and among distributed teams, (2) Provide "out of the box" advanced DI2 capabilities needed
by project management and technical personnel from day one, (3) Provide a low cost DI2 solution for small companies and
teams needing alternatives to expensive, complicated and inflexible project management tools, (4) Be easy to install, use and
maintain without requiring programming or webmaster skills. The proposed innovation merges the Wiki philosophy of fast, easy
and rewarding online content creation with standard project management functions and capabilities for data and information
integration within secure, user-authenticated environments. This facilitates critical elements of communication, interaction and
data capture/synthesis that are often missing or underdeveloped in many traditional project management efforts. Phase 1 TRL
will be 6.
Project Management, Systems, Engineering and Mission Assurance Tools
Information – Data Acquisition and End-to-End Management
USPTO
Patent Number
Title
Assignee
Abstract
US7058588
Dependency-based
work flow integration
and reduction
International Business
Machines Corporation
A process and system is disclosed to assist work planners by assembling a work breakdown structure (WBS) and work flow for a project based on the
explicit selection or deselection of outcome(s) by a work planner from a defined set of possible outcomes. The process and system ensure that the
resulting project WBS and work flow is composed of the minimum set of activities required to produce the set of outcomes desired for the project. The
process and system further ensure that the project's activities are organized into an activity hierarchy defined by a WBS template designated by the
work planner, and that each of the project's activities is linked into an appropriate work flow, supported by appropriate instructional or descriptive
content.
US7058660
System and method for
network-based project
management
Bank One Corporation
A system for managing a project and its associated information over a network is presented. The system includes a project information database
associated with the network and a project information management tool. The project information database stores information associated with a project
being carried out by a corporate entity having at least one sub-entity. The project information management tool includes a project information module
to manage the information associated with the project, and at least one cross-sub-entity project-component module. The cross-sub-entity project
module manages and tracks overall information and sub-entity specific information, of the information associated with the project, related to the
project-component.
US7222130 B1
Report then query
capability for a
multidimensional
database model
Business Objects S. A.
A system and a method for creating an analytical report on top of a multidimensional data model built on top of a relational or multidimensional
database. The database operates in a computer system and provides returned values responsive to queries. Such a query is generated automatically and
is deduced from a report definition. According to one embodiment, a report specification is used by the system and method of the present invention is
able to defer the initial query of the data source, as is the case with conventional reporting tools and methods, until after the report has been defined.
That is, the manner in which a analytic report is defined provides for an automatically generated query. Once the report has been defined, the data to
populate such a report is retrieved to build the document.
US7243110
Searchable archive
Sand Technology Inc.
A searchable archiving system. A searchable archiving system includes one or more compacted files of archive data loosely coupled to a search
process. To create a compacted file, an archiving process tokenizes the archive data, optimizes the tokenized archive data, and extracts archive
metadata from the tokenized data. The tokenized data may then be compressed in a variety of ways into compressed segments that may be individually
accessed and decompressed by the search agents. Before compression, segment metadata is extracted from the segments. The compressed segments
and segment metadata are then combined to create a compacted file. The search process accesses the compacted files by consulting locally stored
archive metadata extracted from the files during the compaction process. The search process then invokes one or more search agents that actively
search the compacted files. The search agents do so by using the segment metadata to identify segments to decompress and search.
Project Management, Systems, Engineering and Mission Assurance Tools
Information – Data Acquisition and End-to-End Management
U.S. Patent Applications
Patent Number
Title
Assignee
Abstract
US20030004922A1
System and method
for data management
ONTRAK Data
International, Inc
US20050022198A1
Computerimplemented process
management system
Task Server
An automated data management system and method for logging, processing, and reporting a large volume of data having different file types,
using different versions, stored on different media, and/or run by different operating systems, includes a first processor for restoring a plurality
of received data files, the data files being capable of being different file types; a file organizing/categorizing processor for organizing the
received data files into data slices, each data slice including an identification number and a descriptor that describes characteristics of the
received data file; a file logging processor for logging the received data files into a first database based on the data slices; a data uploading
processor for uploading the first database to a second database; a de-duplicate processor for calculating a SHA value of the received data files
to determine whether the received data files have duplicates and flagging duplicated data files in the second database; an image conversion
processor for converting at least a portion of the received data files into image files; and a second processor for exporting the image files.
A task management system including a task server linking a plurality of system users, including at least one task definer, at least one task
requestor and at least one task fulfiller all linked over a communications link. The task server includes a task processor for processing tasks, a
task memory for storing task definitions and one or more graphical user interfaces (GUIs) for interfacing the system users to the task server to
facilitate operation of said task processing system. The GUIs include task view interfaces, task fulfiller interfaces, which are used by task
requesters and task fulfillers to request and fulfill tasks, respectively. The GUIs also include a plurality of administrative editor interfaces,
which are used by task definers to define, group and sequence tasks.
US20050034064A1
Method and system
for creating and
following drill links
ActiveViews, Inc.
US20050195660A1
Clustered hierarchical
file services
None
US20050283391A1
Tool for organizing
project-executionmanagement-related
knowledge
None
A method and system for creating and following drill links in a report are disclosed. A relational abstraction of a data store is defined, the
definition including a plurality of views, scalar or aggregate fields associated with the views, and relations between the views. A report is
generated that includes at least one drill link associated with a sequence of one or more relations originating at a base view of the relational
abstraction. Upon selecting a drill link contained in a first report, information about the drill link is extracted from the report, and the extracted
information and the destination view associated with the drill link are used to create a second report.
A system for object-based archival data storage includes an object-based storage subsystem having respective data storage devices, at least one
file presentation interface that interfaces to client platforms, an administration interface having graphical user interface (GUI) and a command
line interface (CLI), a meta data subsystem for storing meta data about files, and includes a virtual file subsystem having a virtual file server
(VFS), a policy subsystem, and a scalable interconnect to couple the object-based storage subsystem, the at least one file presentation interface,
the administration interface, the meta data subsystem, and the policy subsystem, wherein the policy subsystem provides system rules
predetermined by a user for at least one of hash based integrity checking, read-only/write-ability/erase-ability control, and duplicate data
treatment corresponding to files and file objects.
A tool for organizing the project manager's project-execution-management-related (PEMR) knowledge. The tool includes a framework which
is a hierarchically organized categorical structure utilized to categorize the project manager's PEMR knowledge. The tool includes a framework
content editor which performs text creating, modifying, storing and retrieving for editing clues for the knowledge categorized into the
framework. The framework and the clues together form a schema. Accessed and used through the well organized schema, the PEMR
knowledge becomes structured.
Project Management, Systems, Engineering and Mission Assurance Tools
Information – Data Acquisition and End-to-End Management
WIPO
Patent Number
Title
Assignee
Abstract
WO0167291B1
Method, Process and System
for Optimized Outcome
Driven Workflow Synthesis
and Reduction
Pricewaterhousecoopers L L P
WO0211008A1
System and Method for
Project Management
Kokuyo Co., Ltd.
WO03056448A1
Workflow Systems and
Methods for Project
Management and Information
Management
Solectron Corporation
WO03091919A1
Project Management System
Oracle International
Corporation
A process and system is disclosed to assist work planners by assembling a work breakdown structure (WBS) and work flow for a
project based on the explicit selection or de-selection of outcome(s) by a work planner from a defined set of possible outcomes. The
process and system ensure that the resulting project WBS and work flow is composed of the minimum set of activities required to
produce the set of outcomes desired for the project. The process and system further ensure that the project's activities are organized into
an activity hierarchy defined by a WBS template designated by the work planner, and that each of the project's activities is linked into
an appropriate work flow, supported by appropriate instructional content.
A project management system is disclosed which is easy to use by members belonging to a project and other persons. The system
comprises a server (2) and a DB (3) which stores project by project, contents belonging to those projects participated in by some or all
of users. The server (2) comprises a communication controller (4) for transmitting prescribed pages to user terminals (1) and receiving
operation messages from the pages, a project desktop sheet generator (10) for reading out contents data from the database (3) in
response to those operation messages and generating pages for displaying or accessing all the contents belonging to those projects,
project by project, as project desktop sheets, and an access controller (12) for controlling communications with user terminals, when
there has been an access made via the communication controller (4) for the content of the contents, in unit of project desktop.
The present invention relates to workflow systems and methods. In one embodiment, the invention relates to integration of a calendar
system with a workflow system where a calendar event can initiate a workflow by sending a message to a form route manager. The
completion of a workflow or step in the workflow can result in sending a message to a calendar system to generate an event. In another
embodiment, the invention relates to the integration of workflow with a project management system that includes project segments,
which are processes that can be defined and controlled by workflow routes. The project management system sends a message to the
workflow system to initiate workflow and the workflow system sends a message to the project management system, for example, at the
completion of the workflow route. The relationship between the workflow routes can maintain the relationship between the project
segments. If the relationships between project segments change, the relationship between the workflow routes changes without added
effort.
A project management system comprises a database (2) for storing objects relating to a project together with data defining
interdependencies of the objects. A control system (1) responsive to entities (5) connected to it provides access to selected objects
stored on the database (2) and stores on the database (2) updated versions of selected objects received from the entities (5). The control
system (1) is adapted to produce copies of an object that has been accessed by a first entity and all objects that are interdependent on
that object if that object or an object that is interdependent on that object is modified by a second entity between the object being
accessed by the first entity and an updated version being received by the database.
Project Management, Systems, Engineering and Mission Assurance Tools
Information – Data Acquisition and End-to-End Management
WIPO
Patent Number
Title
Assignee
Abstract
WO06055803A2
Multiple-Party Project
Management System and
Method
None
A computerized system of project management is provided. A plurality of selectable pre-determined project parameters are used and, at
least one of the pluralities of selectable pre-determined project parameters has an associated follow-up project parameter. A graphical
user interface is configured to allow selection from said plurality of selectable pre-determined project parameters, display a chronology
object and, in association with the chronology object, at least selected ones of said plurality of selectable pre-determined project
parameters and any associated follow-up project parameters. A processor is configured to receive selections of said plurality of
selectable pre-determined project parameters and to update the chronology object in substantially real time.
WO06076398A2
Predictive Analytic Method
and Apparatus
Metier Ltd
WO07081919A2
Project Management System
and Method
Marware, Inc.
A computerized project management analytical system and method that develops and manages an ontology that links objects and is
capable of being mined. The ontology is comprised of a project ontology framework, a matching engine and a project status matrix that
illustrates a multi-relational view of the project status, of confidence levels, or interdiction points and/or positions on project timelines.
A project management system and method are provided wherein projects can be managed easily and with minimal manual data entry.
Project management software, embodying the project management system of the present invention, can run on a computer network or
user workstation, without requiring a dedicated host server. The project management system of the present invention provides at least
one graphical interface, to permit a user to easily create and/or edit a project, tasks, subprojects and milestones, using the mouse or other
pointer device. A resource window can be provided, conveniently and consistently located in all main views of the system, to assist the
user in efficiently operating the project management system. Further, projects can be automatically updated without repetitive data
entry, using data entered once by the user performing a task.
Project Management, Systems, Engineering and Mission Assurance Tools
Information – Data Acquisition and End-to-End Management
NTRS
Abstract
An online systems engineering tool for flight research projects has been developed through the use of a workgroup database. Capabilities are implemented for typical flight research systems
engineering needs in document library, configuration control, hazard analysis, hardware database, requirements management, action item tracking, project team information, and technical
performance metrics. Repetitive tasks are automated to reduce workload and errors. Current data and documents are instantly available online and can be worked on collaboratively. Existing
forms and conventional processes are used, rather than inventing or changing processes to fit the tool. An integrated tool set offers advantages by automatically cross-referencing data,
minimizing redundant data entry, and reducing the number of programs that must be learned. With a simplified approach, significant improvements are attained over existing capabilities for
minimal cost. By using a workgroup-level database platform, personnel most directly involved in the project can develop, modify, and maintain the system, thereby saving time and money. As a
pilot project, the system has been used to support an in-house flight experiment. Options are proposed for developing and deploying this type of tool on a more extensive basis.
FC
DFRC
Title
Computer-Aided Systems
Engineering for Flight
Research Projects Using a
Workgroup Database
JPL
System, cost, and risk
analysis for access to space
This paper proposes the use of a new tool more quickly develop initial cost and risk estimates of alternative flight options for both single missions and the partnering of missions into a single
space flight. this work is particularly useful for small missions that require low-cost opportunities for accessing space.
MSFC
The Effect of Infrastructure
Sharing in Estimating
Operations Cost of Future
Space Transportation
Systems
NASA and the aerospace industry are extremely serious about reducing the cost and improving the performance of launch vehicles both manned or unmanned. In the aerospace industry,
sharing infrastructure for manufacturing more than one type spacecraft is becoming a trend to achieve economy of scale. An example is the Boeing Decatur facility where both Delta II and Delta
IV launch vehicles are made. The author is not sure how Boeing estimates the costs of each spacecraft made in the same facility. Regardless of how a contractor estimates the cost, NASA in its
popular cost estimating tool, NASA Air force Cost Modeling (NAFCOM) has to have a method built in to account for the effect of infrastructure sharing. Since there is no provision in the most
recent version of NAFCOM2002 to take care of this, it has been found by the Engineering Cost Community at MSFC that the tool overestimates the manufacturing cost by as much as 30.
Therefore, the objective of this study is to develop a methodology to assess the impact of infrastructure sharing so that better operations cost estimates may be made.
JPL
POLARIS: Helping Managers
Get Answers Fast
This viewgraph presentation reviews the Project Online Library and Resource Information System (POLARIS) system. It is NASA-wide, web-based system, providing access to information
related to Program and Project Management. It will provide a one-stop shop for access to: a searchable, sortable database of all requirements for all product lines, project life cycle diagrams
with reviews, project life cycle diagrams with reviews, project review definitions with products review information from NPR 7123.1, NASA Systems Engineering Processes and Requirements,
templates and examples of products, project standard WBSs with dictionaries, and requirements for implementation and approval, information from NASA s Metadata Manager (MdM): Attributes
of Missions, Themes, Programs and amp; Projects, NPR7120.5 waiver form and instructions and much more. The presentation reviews the plans and timelines for future revisions and
modifications.
Project Management, Systems, Engineering and Mission Assurance Tools
Information – Database Development and Interfacing
USPTO
Patent Number
Title
US6613099
Process and System for Providing a
Table View of a Form Layout for a
Database
US7000182
Assistant for creation of layouts or
reports for databases
Assignee
Apple Computer, Inc
Abstract
The present invention provides methods and apparatus for displaying data associated with a plurality of records in a
database, including obtaining a first layout including a body defining a plurality of fields having a specified order and
associated attributes, each of the associated attributes having a corresponding attribute value. A second layout is created
from the first layout such that the second layout includes selected ones of the plurality of fields, the second layout being
adapted for displaying a plurality of records such that values associated with the same field of the plurality of records are
displayed adjacent to one another. Attribute values corresponding to at least some of the attributes associated with the
plurality of fields are then copied from the first layout to the second layout. The second layout is then displayed.
Sun Microsystems, Inc
An assistant for the creation of layouts/reports for databases is disclosed. A layout for a database is the arrangement of
information for the database such as for data entry or screen viewing, and a report (or report format) for a database is the
arrangement of information from the database for presentation of the data in a printed document or with on-line viewing.
The assistant serves to automate in the creation of the layout/reports after an interview sequence with a user.
Project Management, Systems, Engineering and Mission Assurance Tools
Information – Database Development and Interfacing
U.S. Patent Applications
Patent Number
Title
US20030217264A1
System and method for
providing a secure
environment during theUSe
of electronic documents and
data
US20050086221A1
Method of providing data
dictionary-driven web-based
database applications
Assignee
Signitas Corporation
Abstract
The illustrative embodiment of the present discloses a method of providing a secure environment during the use of electronic documents
and data. Authenticated users are able to access, act upon and sign, via a secure connection, a workflow object that is stored on a remote
server. The workflow object includes a sequence of action items, the steps in a workflow, and includes documents or references to
documents required by the workflow. Also included in the workflow object is an Access Control List ( ACL ) which specifies which
users can access which documents at which times. Each document has its own ACL which allows the access of each document to be
specified independently from other documents at a given time. The documents may be encrypted and decrypted using a variety of
methods designed to enhance security, including the use of digital signatures. Once a document is decrypted ( if encrypted), the user
performs a task specified in the workflow using the decrypted document. The workflow is updated to reflect completed tasks, the
document may be electronically signed, and the altered document is then re-encrypted.
None
A method for creating a web-based database application that is data dictionary driven is disclosed. A web site, containing various
computer programs, data for a database application and a data dictionary describing both the structure of an application database and the
requirements for the database application, creates web pages for facilitating the execution of a database application over the internet.
The method comprises the steps of creating, updating and maintaining an on-line data dictionary, and creating and initially populating a
database for the application. The various programs at the web site create web pages for the application, update the data dictionary and
facilitate modification of the data structure for the application. The data dictionary may be initialized: from electronic data uploaded to
the web site; from the data-dictionary records describing an existing web-based application; or directly by the user via a web interface.
The resulting application allows the user to enter information, view information, select the records to be viewed, and make changes to
the application such as form captions, fields displayed, colors used, database structure, and other contents of the data dictionary.
Project Management, Systems, Engineering and Mission Assurance Tools
Information – Human-Computer Interfaces
WIPO
Patent Number
Title
Assignee
Abstract
WO02101534A1
Graphical User Interface with Zoom for
Detail-In-Context Presentations
Idelix Software Inc.
A graphical user interface (GUI) is provided for manipulating a presentation of a region of interest within visual information
displayed on a display screen of a computer display system. The GUI includes: a first bounding shape surrounding the focal
region; a second bounding shape surrounding the shoulder region; a base outline; a pickup point; a slide bar; a move area within
the region of interest; at least one zoom area; and, a zoom button.
Project Management, Systems, Engineering and Mission Assurance Tools
Information – Human-Computer Interfaces
FC
GSFC
Title
A Recipe for Streamlining Mission
Management
ARC
A Reliable Service-Oriented
Architecture for NASA's Mars
Exploration Rover Mission
GAFC
A Virtual Mission Operations Center
Collaborative Environment
NTRS
Abstract
This paper describes a project's design and implementation for streamlining mission management with knowledge capture processes across multiple organizations of a NASA
directorate. Thc project's focus is on standardizing processes and reports enabling secure information access and case of maintenance automating and tracking appropriate workflow
rules through process mapping and infusing new technologies. This paper will describe a small team's experiences using XML technologies through an enhanced vendor suite of
applications integrated on Windows-based platforms called the Wallops Integrated Scheduling and Document Management System (WISDMS). This paper describes our results using
this system in a variety of endeavors, including providing range project scheduling and resource management for a Range and Mission Management Office implementing an automated
Customer Feedback system for a directorate streamlining mission status reporting across a directorate and initiating a document management, configuration management and portal
access system for a Range Safety Office's programs. The end result is a reduction of the knowledge gap through better integration and distribution of information, improved process
performance, automated metric gathering, and quicker identification of problem areas and issues. However, the real proof of the pudding comes through overcoming the user's
reluctance to replace familiar, seasoned processes with new technology ingredients blended with automated procedures in an untested recipe. This paper shares some of the team's
observations that led to better implementation techniques, as well as an IS0 9001 Best Practices citation. This project has provided a unique opportunity to advance NASA's competency
in new technologies, as well as to strategically implement them within an organizational structure, while wetting the appetite for continued improvements in mission management.
The Collaborative Information Portal (CIP) was enterprise software developed jointly by the NASA Ames Research Center and the Jet Propulsion Laboratory (JPL) for NASA's highly
successful Mars Exploration Rover (MER) mission. Both MER and CIP have performed far beyond their original expectations. Mission managers and engineers ran CIP inside the
mission control room at JPL, and the scientists ran CIP in their laboratories, homes, and offices. All the users connected securely over the Internet. Since the mission ran on Mars time,
CIP displayed the current time in various Mars and Earth time zones, and it presented staffing and event schedules with Martian time scales. Users could send and receive broadcast
messages, and they could view and download data and image files generated by the rovers' instruments. CIP had a three-tiered, service-oriented architecture (SOA) based on industry
standards, including J2EE and web services, and it integrated commercial off-the-shelf software. A user's interactions with the graphical interface of the CIP client application generated
web services requests to the CIP middleware. The middleware accessed the back-end data repositories if necessary and returned results for these requests. The client application could
make multiple service requests for a single user action and then present a composition of the results. This happened transparently, and many users did not even realize that they were
connecting to a server. CIP performed well and was extremely reliable; it attained better than 99 uptime during the course of the mission. In this paper, we present overviews of the MER
mission and of CIP. We show how CIP helped to fulfill some of the mission needs and how people used it. We discuss the criteria for choosing its architecture, and we describe how the
developers made the software so reliable. CIP's reliability did not come about by chance, but was the result of several key design decisions. We conclude with some of the important
lessons we learned form developing, deploying, and supporting the software.
The Virtual Mission Operations Center - Collaborative Environment (VMOC-CE) intent is to have a central access point for all the resources used in a collaborative mission operations
environment to assist mission operators in communicating on-site and off-site in the investigation and resolution of anomalies. It is a framework that as a minimum incorporates online
chat, realtime file sharing and remote application sharing components in one central location. The use of a collaborative environment in mission operations opens up the possibilities for
a central framework for other project members to access and interact with mission operations staff remotely. The goal of the Virtual Mission Operations Center (VMOC) Project is to
identify, develop, and infuse technology to enable mission control by on-call personnel in geographically dispersed locations. In order to achieve this goal, the following capabilities are
needed Autonomous mission control systems Automated systems to contact on-call personnel Synthesis and presentation of mission control status and history information Desktop
tools for data and situation analysis Secure mechanism for remote collaboration commanding Collaborative environment for remote cooperative work The VMOC-CE is a collaborative
environment that facilitates remote cooperative work. It is an application instance of the Virtual System Design Environment (VSDE), developed by NASA Goddard Space Flight Center's
(GSFC) Systems Engineering Services and Advanced Concepts (SESAC) Branch. The VSDE is a web-based portal that includes a knowledge repository and collaborative environment
to serve science and engineering teams in product development. It is a 'one stop shop' for product design, providing users real-time access to product development data, engineering
and management tools, and relevant design specifications and resources through the Internet. The initial focus of the VSDE has been to serve teams working in the early portion of the
system product lifecycle - concept development, proposal preparation, and formulation. The VMOC-CE expands the application of the VSDE into the operations portion of the system
lifecycle. It will enable meaningful and real-time collaboration regardless of the geographical distribution of project team members. Team members will be able to interact in satellite
operations, specifically for resolving anomalies, through access to a desktop computer and the Internet. Mission Operations Management will be able to participate and monitor up to the
minute status of anomalies or other mission operations issues. In this paper we present the VMOC-CE project, system capabilities, and technologies.
Project Management, Systems, Engineering and Mission Assurance Tools
Information – Human-Computer Interfaces
NTRS
Abstract
What do you get when you cross rocket scientists with computer geeks It is an interactive, distributed computing web of tools and services providing a more productive environment for
propulsion research and development. The Rocket Engine Advancement Program 2 (REAP2) project involves researchers at several institutions collaborating on propulsion experiments
and modeling. In an effort to facilitate these collaborations among researchers at different locations and with different specializations, researchers at the Information Technology and
Systems Center,' University of Alabama in Huntsville, are creating a prototype web-based interactive information system in support of propulsion research. This system, to be based on
experience gained in creating similar systems for NASA Earth science field experiment campaigns such as the Convection and Moisture Experiments (CAMEX), will assist in the
planning and analysis of model and experiment results across REAP2 participants. The initial version of the Propulsion Experiment Project Management Environment (PExPM) consists
of a controlled-access web portal facilitating the drafting and sharing of working documents and publications. Interactive tools for building and searching an annotated bibliography of
publications related to REAP2 research topics have been created to help organize and maintain the results of literature searches. Also work is underway, with some initial prototypes in
place, for interactive project management tools allowing project managers to schedule experiment activities, track status and report on results. This paper describes current successes,
plans, and expected challenges for this project.
FC
MSFC
Title
Building a Propulsion Experiment
Project Management Environment
ARC
CIP Training Manual Collaborative
Information Portal Advance Training
Information for Field Test Participants
The Collaborative Information Portal (CIP) is a web-based information management and retrieval system. Its purpose is to provide users at MER (Mars Exploration Rover) mission
operations with easy access to a broad range of mission data and products and contextual information such as the current operations schedule. The CIP web-server provides this
content in a user customizable web-portal environment. Since CIP is still under development, only a subset of the full feature set will be available for the EDO field test. The CIP webportal will be accessed through a standard web browser. CIP is intended to be intuitive and simple to use, however, at the training session, users will receive a one to two page
reference guide, which should aid them in using CIP. Users must provide their own computers for accessing CIP during the field test. These computers should be configured with Java
1.3 and a Java 2 enabled browser. Macintosh computers should be running OS 10.1.3 or later. Classic Mac OS (OS 9) is not supported. For more information please read section 7.3 in
the FIASCO Rover Science Operations Test Mission Plan. Several screen shots of the Beta Release of CIP are shown on the following pages.
Goddard
Space
Flight
Center
REACH Real-Time Data Awareness in
Multi-Spacecraft Missions
Missions have been proposed that will use multiple spacecraft to perform scientific or commercial tasks. Indeed, in the commercial world, some spacecraft constellations already exist.
Aside from the technical challenges of constructing and flying these missions, there is also the financial challenge presented by the tradition model of the flight operations team (FOT)
when it is applied to a constellation mission. Proposed constellation missions range in size from three spacecraft to more than 50. If the current ratio of three-to-five FOT personnel per
spacecraft is maintained, the size of the FOT becomes cost prohibitive. The Advanced Architectures and Automation Branch at the Goddard Space Flight Center (GSFC Code 588) saw
the potential to reduce the cost of these missions by creating new user interfaces to the ground system health-and-safety data. The goal is to enable a smaller FOT to remain aware and
responsive to the increased amount of ground system information in a multi-spacecraft environment. Rather than abandon the tried and true, these interfaces were developed to run
alongside existing ground system software to provide additional support to the FOT. These new user interfaces have been combined in a tool called REACH. REACH-the Real-time
Evaluation and Analysis of Consolidated Health-is a software product that uses advanced visualization techniques to make spacecraft anomalies easy to spot, no matter how many
spacecraft are in the constellation. REACH reads a real-time stream of data from the ground system and displays it to the FOT such that anomalies are easy to pick out and investigate.
Data visualization has been used in ground system operations for many years. To provide a unique visualization tool, we developed a unique source of data to visualize the REACH
Health Model Engine. The Health Model Engine is rule-based software that receives real-time telemetry information and outputs 'health' information related to the subsystems and
spacecraft that the telemetry belong to. The Health Engine can run out-of-the-box or can be tailored with a scripting language. Out of the box, it uses limit violations to determine the
health of subsystems and spacecraft when tailored, it determines health using equations combining the values and limits of any telemetry in the spacecraft. The REACH visualizations
then 'roll up' the information from the Health Engine into high level, summary displays. These summary visualizations can be 'zoomed' into for increasing levels of detail. Currently
REACH is installed in the Small Explorer (SMEX) lab at GSFC, and is monitoring three of their five spacecraft. We are scheduled to install REACH in the Mid-sized Explorer (MIDEX)
lab, which will allow us to monitor up to six more spacecraft. The process of installing and using our 'research' software in an operational environment has provided many insights into
which parts of REACH are a step forward and which of our ideas are missteps. Our paper explores both the new concepts in spacecraft health-and-safety visualization, the difficulties of
such systems in the operational environment, and the cost and safety issues of multi-spacecraft missions.
Project Management, Systems, Engineering and Mission Assurance Tools
Information – Human-Computer Interfaces
FC
ARC/MSFC
Title
Spaceflight Operations Services Grid
(SOSG)
ARC
THE COLLABORATIVE
INFORMATION PORTAL AND NASA'S
MARS EXPLORATION ROVER
MISSION
ARC
The MER CIP Portal for Ground
Operations
NTRS
Abstract
I In an effort to adapt existing space flight operations services to new emerging Grid technologies we are developing a Grid-based prototype space flight operations Grid. This prototype
is based on the operational services being provided to the International Space Station's Payload operations located at the Marshall Space Flight Center, Alabama. The prototype
services will be Grid or Web enabled and provided to four user communities through portal technology. Users will have the opportunity to assess the value and feasibility of Grid
technologies to their specific areas or disciplines. In this presentation descriptions of the prototype development, User-based services, Grid-based services and status of the project will
be presented. Expected benefits, findings and observations (if any) to date will also be discussed. The focus of the presentation will be on the project in general, status to date and future
plans. The End-use services to be included in the prototype are voice, video, telemetry, commanding, collaboration tools and visualization among others. Security is addressed
throughout the project and is being designed into the Grid technologies and standards development. The project is divided into three phases. Phase One establishes the baseline Userbased services required for space flight operations listed above. Phase Two involves applying Gridlweb technologies to the User-based services and development of portals for access
by users. Phase Three will allow NASA and end users to evaluate the services and determine the future of the technology as applied to space flight operational services. Although,
Phase One, which includes the development of the quasi-operational User-based services of the prototype, development will be completed by March 2004, the application of Grid
technologies to these services will have just begun. We will provide status of the Grid technologies to the individual User-based services. This effort will result in an extensible
environment that incorporates existing and new spaceflight services into a standards-based framework providing current and future NASA programs with cost savings and new and
evolvable methods to conduct science. This project will demonstrate how the use of new programming paradigms such as web and grid services can provide three significant benefits to
the cost-effective delivery of spaceflight services. They will enable applications to operate more efficiently by being able to utilize pooled resources. They will also permit the reuse of
common services to rapidly construct new and more powerful applications. Finally they will permit easy and secure access to services via a combination of grid and portal technology by
a distributed user community consisting of NASA operations centers, scientists, the educational community and even the general population as outreach. The approach will be to deploy
existing mission support applications such as the Telescience Resource Kit (TReK) and new applications under development, such as the Grid Video Distribution System (GViDS),
together with existing grid applications and services such as high-performance computing and visualization services provided by NASA s Information Power Grid (IPG) in the MSFC s
Payload Operations Integration Center (POIC) HOSC Annex. Once the initial applications have been moved to the grid, a process will begin to apply the new programming paradigms to
integrate them where possible. For example, with GViDS, instead of viewing the Distribution service as an application that must run on a single node, the new approach is to build it such
that it can be dispatched across a pool of resources in response to dynamic loads. To make this a reality, reusable services will be critical, such as a brokering service to locate
appropriate resource within the pool. This brokering service can then be used by other applications such as the TReK. To expand further, if the GViDS application is constructed using a
services-based mel, then other applications such as the Video Auditorium can then use GViDS as a service to easily incorporate these video streams into a collaborative conference.
Finally, as these applications are re-factored into this new services-based paradigm, the construction of portals to integrate them will be a simple process. As a result, portals can be
tailored to meet the requirements of specific user communities.
THE COLLABORATIVE INFORMATION PORTAL WAS ENTERPRISE SOFTWARE DEVELOPED JOINTLY BY THE NASA AMES RESEARCH CENTER AND THE JET
PROPULSION LABORATORY FOR NASA'S MARS EXPLORATION ROVER MISSION. MISSION MANAGERS, ENGINEERS, SCIENTISTS, AND RESEARCHERS USED THIS
INTERNET APPLICATION TO VIEW CURRENT STAFFING AND EVENT SCHEDULES, DOWNLOAD DATA AND IMAGE FILES GENERATED BY THE ROVERS, RECEIVE
BROADCAST MESSAGES, AND GET ACCURATE TIMES IN VARIOUS MARS AND EARTH TIME ZONES. THIS ARTICLE DESCRIBES THE FEATURES, ARCHITECTURE, AND
IMPLEMENTATION OF THIS SOFTWARE, AND CONCLUDES WITH LESSONS WE LEARNED FROM ITS DEPLOYMENT AND A LOOK TOWARDS FUTURE MISSIONS.
We developed the Mars Exploration Rover Collaborative Information Portal (MER CIP) to facilitate MER operations. MER CIP provides a centralized, one-stop delivery platform
integrating science and engineering data from several distributed heterogeneous data sources. Key issues for MER CIP include 1) Scheduling and schedule reminders 2) Tracking the
status of daily predicted outputs 3) Finding and analyzing data products 4) Collaboration 5) Announcements 6) Personalization.
Project Management, Systems, Engineering and Mission Assurance Tools
Information – Software Tools for Distributed Analysis and Simulation
Company Name
Trident Systems, Inc.
Title
MODELING THE IMPACT OF TECHNOLOGY TRANSITION ON
SHIP OPERATIONAL CAPABILITIES
DOD SBIR Phase II
Quad Chart
The result of this combined effort will be a Collaborative Tool Framework (CTF) that will provide open Application Programming
Interfaces (APIs) allowing companies to plug-in tools that can aid the decision makers in the task to determine which
technologies to add to a system, when to add them, and once added, accurately predict the impact on the rest of the existing
system.
Project Management, Systems, Engineering and Mission Assurance Tools
Structures – Launch and Flight Vehicle
Company Name
Infoscitex Corp.
Title
Multifunctional Programmable Structure
(IPS) for a Modular Satellite Architecture
DOD SBIR Phase I
Quad Chart
The DOD’s Operationally Responsive Space paradigm calls for ambitiously reducing the time required to design, fabricate and deploy a satellite to
days or even hours, from the current requirement of months or years. This timeline will require modular satellite structures available on site that support
plug-n-play integration with payload components. Infoscitex proposes a multi-functional structural panel with programmable electronic and thermal
management capabilities. A light-weight, carbon fiber composite core structure offers structural integrity and acts to efficiently conduct heat away from
thermally-controlled payloads. An innovative heat spreading technology is applied over the core, allowing it to act as either an insulator or a heat sink
via real-time electronic control. This feature allows the proposed panel to provide appropriate thermal control to many payloads with disparate cooling
requirements. The integrated panel supports intra-panel wiring and component interface ports compatible with emerging designs for a programmable
wiring harness. During Phase I the project team will evaluate preliminary concepts through modeling and analysis, optimize the design of an individual
standardized panel, define inter-panel interfaces, and demonstrate feasibility of a multi-panel structure. The Phase I objective will be to develop an
optimum panel design that can be prototyped and demonstrated during the Phase II.
Project Management, Systems, Engineering and Mission Assurance Tools
Structures – Modular Interconnects
Company Name
Spaceworks, Inc.
Title
Reconfigurable PnP Spacecraft Structure
DOD SBIR Phase I
Quad Chart
Plug-and-play avionics for spacecraft are being developed by the Air Force to create systems that promise revolutionary improvements in
responsiveness for military users. To keep pace, similar developments are needed in other supporting systems. SpaceWorks proposes to develop plugand-play structures for spacecraft that have features that support these advanced avionics systems and that also contribute to the overall goals of
reconfigurability, flexibility, modularity, and rapid assembly, integration, and checkout. The approach involves a small number of spacecraft panel types
that can be assembled rapidly to form many useful spacecraft configurations. The panels have identical mechanical and electrical interfaces at discrete
sites and between panels to enable rapid plug-and-play functionality for payloads and components. The plug-and-play electrical network consisting of
electronics, harnessing, and connectors is recessed within each panel to maximize useable volume and footprint. Panel design and materials maximize
the flexibility to incorporate multifunctional features such as thermal management devices, integrated structural sensors, and layered radiation shielding.
SpaceWorks will fabricate a set of these panels leading to a demonstration showing how the approach supports plug-and-play avionics, provides a high
degree of reconfigurability, and can meet a very demanding integration and checkout timeline.
Project Management, Systems, Engineering and Mission Assurance Tools
Thermal – Control Instrumentation
FC
GSFC
Title
Controlling Micro ElectroMechanical Systems (MEMS)
in Space
NTRS
Abstract
Small spacecraft, including micro and nanosats, as they are envisioned for future missions, will require an alternative means to achieve thermal control due to
their small power and mass budgets. One of the proposed alternatives is Variable Emittance (Vari-E) Coatings for spacecraft radiators. Space Technology-5
(ST-5) is a technology demonstration mission through NASA Goddard Space Flight Center (GSFC) that will utilize Vari-E Coatings. This mission involves a
constellation of three (3) satellites in a highly elliptical orbit with a perigee altitude of approx200 km and an apogee of approx38,000 km. Such an environment
will expose the spacecraft to a wide swing in the thermal and radiation environment of the earth's atmosphere. There are three (3) different technologies
associated with this mission. The three technologies are electrophoretic, electrochromic, and Micro ElectroMechanical Systems (MEMS). The ultimate goal is
to make use of Vari-E coatings, in order to achieve various levels of thermal control. The focus of this paper is to highlight the Vari-E Coating MEMS
instrument, with an emphasis on the Electronic Control Unit responsible for operating the MEMS device. The Test and amp Evaluation approach, along with
the results, is specific for application on ST-5, yet the information provides a guideline for future experiments and or thermal applications on the exterior
structure of a spacecraft. copyright 2003 American Institute of Physics
Project Management, Systems, Engineering and Mission Assurance Tools
Verification and Validation – Operations Concepts and Requirements
FC
GSFC
Title
Space Technology 5 Changing the Mission Design without
Changing the Hardware
NTRS
Abstract
The Space Technology 5 (ST-5) Project is part of NASA's New Millennium Program. The validation objectives are to demonstrate the research-quality
science capability of the ST-5 spacecraft to operate the three spacecraft as a constellation and to design, develop, test and flight-validate three capable
micro-satellites with new technologies. A three-month flight demonstration phase is planned, beginning in March 2006. This year, the mission was replanned for a Pegasus XL dedicated launch into an elliptical polar orbit (instead of the Originally-planned Geosynchronous Transfer Orbit.) The re-plan
allows the mission to achieve the same high-level technology validation objectives with a different launch vehicle. The new mission design involves a
revised science validation strategy, a new orbit and different communication strategy, while minimizing changes to the ST-5 spacecraft itself. The
constellation operations concepts have also been refined. While the system engineers, orbit analysts, and operations teams were re-planning the
mission, the implementation team continued to make progress on the flight hardware. Most components have been delivered, and the first spacecraft is
well into integration and test.
Project Management, Systems, Engineering and Mission Assurance Tools
Verification and Validation – Testing Requirements and Architectures
FC
GSFC
Title
Integration and Testing Challenges of Small Satellite Missions
Experiences from the Space Technology 5 Project
NTRS
Abstract
The Space Technology 5(ST5) payload was successfully carried into orbit on an OSC Pegasus XL launch vehicle, which was carried aloft
and dropped from the OSC Lockheed L-1011 from Vandenberg Air Force Base March 22,2006, at 903 am Eastern time, 603 am Pacific
time. In order to reach the completion of the development and successful launch of ST 5, the systems integration and test(I and ampT)
team determined that a different approach was required to meet the project requirements rather than the standard I and ampT approach
used for single, room-sized satellites. The ST5 payload, part of NASA's New Millennium Program headquartered at JPL, consisted of three
micro satellites (approximately 30 kg each) and the Pegasus Support Structure (PSS), the system that connected the spacecrafts to the
launch vehicle and deployed the spacecrafts into orbit from the Pegasus XL launch vehicle. ST5 was a technology demonstration payload,
intended to test six (6) new technologies for potential use for future space flights along with demonstrating the ability of small satellites to
perform quality science. The main technology was a science grade magnetometer designed to take measurements of the earth's magnetic
field. The three spacecraft were designed, integrated, and tested at NASA Goddard Space Flight Center with integration and
environmental testing occurring in the Bldg. 7-1 0-15-29. The three spacecraft were integrated and tested by the same I and ampT team.
The I and ampT Manager determined that there was insufficient time in the schedule to perform the three I and ampT spacecraft activities
in series used standard approaches. The solution was for spacecraft 1 to undergo integration and test first, followed by spacecraft 2 and 3
simultaneously. This simultaneous integration was successful for several reasons. Each spacecraft had a Lead Test Conductor who
planned and coordinated their spacecraft through its integration and test activities. One team of engineers and technicians executed the
integration of all three spacecraft, learning and gaining knowledge and efficiency as spacecraft 1 integration and testing progressed. They
became acutely familiar with the hardware, operation and processes for I and ampT, thus each team member had the experience and
knowledge to safely execute I and ampT for spacecraft 2 and 3 together. The integration team was very versatile and each member could
perform many different activities or work any spacecraft, when needed. Daily meetings between the three Lead TCs and technician team
allowed the team to plan and implement activities efficiently. The three (3) spacecraft and PSS were successfully integrated and tested,
shipped to the launch site, and ready for launch per the I and ampT schedule that was planned three years previously.
Smart Autonomous Command and Data Handling System, Algorithms and Data Management
Avionics and Astrionics – Guidance Navigation and Control
FC
GSFC
Title
Disturbance Reduction Control Design for the
ST7 Flight Validation Experiment
NTRS
Abstract
The Space Technology 7 experiment will perform an on-orbit system-level validation of two specific Disturbance Reduction System technologies a gravitational
reference sensor employing a free-floating test mass, and a set of micro-Newton colloidal thrusters. The ST7 Disturbance Reduction System is designed to maintain the
spacecraft's position with respect to a free-floating test mass to less than 10 nm Hz, over the frequency range of 1 to 30 mHz. This paper presents the design and
analysis of the coupled, drag-free and attitude control systems that close the loop between the gravitational reference sensor and the micro-Newton thrusters, while
incorporating star tracker data at low frequencies. A full 18 degree-of-freedom model, which incorporates rigid-body models of the spacecraft and two test masses, is
used to evaluate the effects of actuation and measurement noise and disturbances on the performance of the drag-free system.
Smart Autonomous Command and Data Handling System, Algorithms and Data Management
Avionics and Astrionics – Onboard Computing and Data Management
FC
JPL
Title
Autonomous Detection of Dust Devils and Clouds on
Mars
GSFC
Evaluation of Algorithms for Compressing
Hyperspectral Data
NTRS
Abstract
Acquisition of science in space applications is shifting from teleoperated gathering to an automated on-board analysis with improvements in the use of
on-board memory, CPU, bandwidth and data quality. In this paper, we describe algorithms to autonomously detect dust devils and clouds from a rover
and summarize the results. These algorithms meet high hit-to-miss ratios and satisfy strict requirements of CPU, memory usage and bandwidth. The
detectors have been scheduled for upload to the Mars Exploration Rovers (MER) in 2006. These are the first autonomous science processes in the
rovers.
With EO-1 Hyperion in orbit NASA is showing their continued commitment to hyperspectral imaging (HSI). As HSI sensor technology continues to
mature, the ever-increasing amounts of sensor data generated will result in a need for more cost effective communication and data handling systems.
Lockheed Martin, with considerable experience in spacecraft design and developing special purpose onboard processors, has teamed with Applied
Signal and Image Technology (ASIT), who has an extensive heritage in HSI spectral compression and Mapping Science (MSI) for JPEG 2000 spatial
compression expertise, to develop a real-time and intelligent onboard processing (OBP) system to reduce HSI sensor downlink requirements. Our
goal is to reduce the downlink requirement by a factor greater than 100, while retaining the necessary spectral and spatial fidelity of the sensor data
needed to satisfy the many science, military, and intelligence goals of these systems. Our compression algorithms leverage commercial-off-the-shelf
(COTS) spectral and spatial exploitation algorithms. We are currently in the process of evaluating these compression algorithms using statistical
analysis and NASA scientists. We are also developing special purpose processors for executing these algorithms onboard a spacecraft.
Smart Autonomous Command and Data Handling System, Algorithms and Data Management
Avionics and Astrionics – Telemetry, Tracking and Control
Company Name
Asier Technology Corp.
Title
INTEGRATED DATA COMPRESSION AND
SECURITY ALGORITHMS
DOD SBIR Phase II
Quad Chart
Asier determined that two different types of MDA telemetry data could be compressed and encrypted in real-time (no added latency)
and without damaging either the integrity or synchronicity of the data stream.
Phase II will continue the testing schedule and characterize the compressibility of data from all telemetry-sending devices involved in a
THAAD test flight.
Smart Autonomous Command and Data Handling System, Algorithms and Data Management
Avionics and Astrionics – Telemetry, Tracking and Control
FC
GSFC
Title
Autonomous Telemetry Collection for
Single-Processor Small Satellites
NTRS
Abstract
For the Space Technology 5 mission, which is being developed under NASA's New Millennium Program, a single spacecraft processor will be required to do on-board realtime computations and operations associated with attitude control, up-link and down-link communications, science data processing, solid-state recorder management,
power switching and battery charge management, experiment data collection, health and status data collection, etc. Much of the health and status information is in analog
form, and each of the analog signals must be routed to the input of an analog-to-digital converter, converted to digital form, and then stored in memory. If the microoperations of the analog data collection process are implemented in software, the processor may use up a lot of time either waiting for the analog signal to settle, waiting
for the analog-to-digital conversion to complete, or servicing a large number of high frequency interrupts. In order to off-load a very busy processor, the collection and
digitization of all analog spacecraft health and status data will be done autonomously by a field-programmable gate array that can configure the analog signal chain,
control the analog-to-digital converter, and store the converted data in memory.
Smart Autonomous Command and Data Handling System, Algorithms and Data Management
Communications – Architectures and Networks
U.S. Patent Applications
Patent Number
Title
Assignee
Abstract
US2005117610A1
Compressor, decompressor, data block and
resource management method
ALCATEL
The invention relates to a compressor of data from data frames. According to the invention, the compressor detects "active"
elements that vary from one frame to another and transmits only the "active" elements in a data block. This block also
contains a state code locating this "active" data in the original frame, in order to be able to reconstitute it. The invention also
relates to a data decompressor, a data block containing a data group compressed in accordance with the invention, and a
bandwidth resource management method. Particular application in a satellite telecommunication system.
Smart Autonomous Command and Data Handling System, Algorithms and Data Management
Communications – Architectures and Networks
WIPO
Patent Number
Title
Assignee
Abstract
WO05057342A2
A Method and System of Bandwidth
Management for Streaming Data
Wilife, Inc.
A data networking system and method which allows efficient use of bandwidth for data streams such as video and audio. This
invention allows network nodes to dynamically identify changing network conditions which are typical on wireless and power line
networks. The system and method dynamically adapt to the changes which affect network bandwidth by changing compression rates,
compression types, audio/video quality, motion masks, throughput for specific connections, or mass storage of data streams until the
network is capable of sending the data. The result is an improved system that requires little or no user intervention as network
conditions change.
Smart Autonomous Command and Data Handling System, Algorithms and Data Management
Communications – Architectures and Networks
FC
Jet Propulsion Laboratory
Title
A Multi-mission Event-Driven Component-Based System
for Support of Flight Software Development, ATLO, and
Operations first used by the Mars Science Laboratory
(MSL) Project
GRC
Advanced Communications Architecture Demonstration
Made Significant Progress
NTRS
Abstract
This paper details an architectural description of the Mission Data Processing and Control System (MPCS), an event-driven, multi-mission ground
data processing components providing uplink, downlink, and data management capabilities which will support the Mars Science Laboratory (MSL)
project as its first target mission. MPCS is developed based on a set of small reusable components, implemented in Java, each designed with a
specific function and well-defined interfaces. An industry standard messaging bus is used to transfer information among system components.
Components generate standard messages which are used to capture system information, as well as triggers to support the event-driven
architecture of the system. Event-driven systems are highly desirable for processing high-rate telemetry (science and engineering) data, and for
supporting automation for many mission operations processes.
Simulation for a ground station located at 44.5 deg latitude. The Advanced Communications Architecture Demonstration (ACAD) is a concept
architecture to provide high-rate Ka-band (27-GHz) direct-to-ground delivery of payload data from the International Space Station. This new
concept in delivering data from the space station targets scientific experiments that buffer data onboard. The concept design provides a method to
augment the current downlink capability through the Tracking Data Relay Satellite System (TDRSS) Ku-band (15-GHz) communications system.
The ACAD concept pushes the limits of technology in high-rate data communications for space-qualified systems. Research activities are ongoing
in examining the various aspects of high-rate communications systems including (1) link budget parametric analyses, (2) antenna configuration
trade studies, (3) orbital simulations (see the preceding figure), (4) optimization of ground station contact time (see the following graph), (5)
processor and storage architecture definition, and (6) protocol evaluations and dependencies.
Smart Autonomous Command and Data Handling System, Algorithms and Data Management
Communications – Autonomous Control and Monitoring
FC
Goddard Space Flight Center
Title
Agent Based Software for the Autonomous Control of
Formation Flying Spacecraft
ARC
Automated Recognition of Geologically Significant Shapes
in MER PANCAM and MI Images
NTRS
Abstract
Distributed satellite systems is an enabling technology for many future NASA DoD earth and space science missions, such as MMS,
MAXIM, Leonardo, and LISA 1, 2, 3. While formation flying offers significant science benefits, to reduce the operating costs for these
missions it will be essential that these multiple vehicles effectively act as a single spacecraft by performing coordinated observations.
Autonomous guidance, navigation, and control as part of a coordinated fleet-autonomy is a key technology that will help accomplish this
complex goal. This is no small task, as most current space missions require significant input from the ground for even relatively simple
decisions such as thruster burns. Work for the NMP DS1 mission focused on the development of the New Millennium Remote Agent
(NMRA) architecture for autonomous spacecraft control systems. NMRA integrates traditional real-time monitoring and control with
components for constraint-based planning, robust multi-threaded execution, and model-based diagnosis and reconfiguration. The
complexity of using an autonomous approach for space flight software was evident when most of its capabilities were stripped off prior to
launch (although more capability was uplinked subsequently, and the resulting demonstration was very successful).
Autonomous recognition of scientifically important information provides the capability of 1) Prioritizing data return 2) Intelligent data
compression 3) Reactive behavior onboard robotic vehicles. Such capabilities are desirable as mission scenarios include longer durations
with decreasing interaction from mission control. To address such issues, we have implemented several computer algorithms, intended to
autonomously recognize morphological shapes of scientific interest within a software architecture envisioned for future rover missions.
Mars Exploration Rovers (MER) instrument payloads include a Panoramic Camera (PANCAM) and Microscopic Imager (MI). These
provide a unique opportunity to evaluate our algorithms when applied to data obtained from the surface of Mars. Early in the mission we
applied our algorithms to images available at the mission web site (http marsrovers.jpl.nasa.gov gallery images.html), even though these
are not at full resolution. Some algorithms would normally use ancillary information, e.g. camera pointing and position of the sun, but
these data were not readily available. The initial results of applying our algorithms to the PANCAM and MI images are encouraging. The
horizon is recognized in all images containing it such information could be used to eliminate unwanted areas from the image prior to data
transmission to Earth. Additionally, several rocks were identified that represent targets for the mini-thermal emission spectrometer. Our
algorithms also recognize the layers, identified by mission scientists. Such information could be used to prioritize data return or in a
decision-making process regarding future rover activities. The spherules seen in MI images were also autonomously recognized. Our
results indicate that reliable recognition of scientifically relevant morphologies in images is feasible.
Smart Autonomous Command and Data Handling System, Algorithms and Data Management
Communications – RF
DOD SBIR Phase II
Quad Chart
Video data compression is a powerful enabling technology for surveillance applications. It permits the transmission of high data-volume video
surveillance data in real-time or near real-time, by encoding the data in an efficient compressed format, to be reconstructed at the receiver.
Company Name
FASTVDO LLC
Title
VIDEO DATA COMPRESSION
Physical Optics Corporation
MOBILE ADVANCED DATA COMPRESSION
ADAPTIVE TECHNOLOGY
Mobile Advanced Data Compression Adaptive Technology (MADCAT) integrates video transmission capability into tactical radios without
requiring new data link equipment, significantly reducing platform weight and integration cost. In addition, MADCAT offers secure transmission by
means of FHSS and crypto, and strong error resilience for maximum video stream continuity through noisy channels.
RNET Technologies, Inc.
VIRTUAL OBJECTS BASED COMPRESSION
(VOBC) OF VIDEO
A completely new video compression approach, known as VOBC, which is based on “Virtual Objects (VO)” was developed. From source video,
VOs are extracted over a group of frames, i.e., GOF (typically 16) and the extracted VOs and background are compressed separately. As a
result, very high compression ratios are obtained. The VOs, which include “real objects”, can be compressed in a “loss-less” manner, while the
background is compressed in a “lossy” manner.
Smart Autonomous Command and Data Handling System, Algorithms and Data Management
Electronics – Highly Reconfigurable
FC
Johnson Space Center
Title
Use of Field Programmable Gate Array
Technology in Future Space Avionics
NTRS
Abstract
Fulfilling NASA's new vision for space exploration requires the development of sustainable, flexible and fault tolerant spacecraft control systems. The
traditional development paradigm consists of the purchase or fabrication of hardware boards with fixed processor and or Digital Signal Processing
(DSP) components interconnected via a standardized bus system. This is followed by the purchase and or development of software. This paradigm
has several disadvantages for the development of systems to support NASA's new vision. Building a system to be fault tolerant increases the
complexity and decreases the performance of included software. Standard bus design and conventional implementation produces natural
bottlenecks. Configuring hardware components in systems containing common processors and DSPs is difficult initially and expensive or impossible
to change later. The existence of Hardware Description Languages (HDLs), the recent increase in performance, density and radiation tolerance of
Field Programmable Gate Arrays (FPGAs), and Intellectual Property (IP) Cores provides the technology for reprogrammable Systems on a Chip
(SOC). This technology supports a paradigm better suited for NASA's vision. Hardware and software production are melded for more effective
development they can both evolve together over time. Designers incorporating this technology into future avionics can benefit from its flexibility.
Systems can be designed with improved fault isolation and tolerance using hardware instead of software. Also, these designs can be protected from
obsolescence problems where maintenance is compromised via component and vendor availability. To investigate the flexibility of this technology,
the core of the Central Processing Unit and Input Output Processor of the Space Shuttle AP101S Computer were prototyped in Verilog HDL and
synthesized into an Altera Stratix FPGA.
Smart Autonomous Command and Data Handling System, Algorithms and Data Management
Information – Data Acquisition and End-to-End Management
U.S. Patent Applications
Patent Number
Title
Assignee
Abstract
US2005187777A1
Layer 2
compression/decompression for
mixed
synchronous/asynchronous
transmission of data frames
within a communication
network
ALCATEL
A device is disclosed for compressing data contained in input frames to be compressed constituted of stream frames defining portions of TRAU and
signaling frames that have to be transmitted within a communication network and each of which is constituted of at least a header containing control data
representative at least of the type of stream frame and where applicable payload data, certain types containing critical and/or non-critical data. The device
analyzes each TRAU or signaling frame header contained in successively received input frames in order to determine its type and generates periodically
compressed frames to be transmitted that are divided into first and second sections of variable size. The first section contains critical data compressed
synchronously and the second section contains non-critical data compressed asynchronously.
Smart Autonomous Command and Data Handling System, Algorithms and Data Management
Information – Data Acquisition and End-to-End Management
NTRS
Abstract
A highly performing image data compression technique is currently being developed for space science applications under the requirement of high-speed and pushbroom
scanning. The technique is also applicable to frame based imaging data. The algorithm combines a two-dimensional transform with a bitplane encoding this results in an
embedded bit string with exact desirable compression rate specified by the user. The compression scheme performs well on a suite of test images acquired from
spacecraft instruments. It can also be applied to three-dimensional data cube resulting from hyper-spectral imaging instrument. Flight qualifiable hardware
implementations are in development. The implementation is being designed to compress data in excess of 20 Msampledsec and support quantization from 2 to 16 bits.
This paper presents the algorithm, its applications and status of development.
FC
GSFC
Title
A High Performance Image Data Compression
Technique for Space Applications
GSFC
Data compression using Chebyshev transform
The present invention is a method, system, and computer program product for implementation of a capable, general purpose compression algorithm that can be engaged
on the fly. This invention has particular practical application with time-series data, and more particularly, time-series data obtained form a spacecraft, or similar situations
where cost, size and/or power limitations are prevalent, although it is not limited to such applications. It is also particularly applicable to the compression of serial data
streams and works in one, two, or three dimensions. The original input data is approximated by Chebyshev polynomials, achieving very high compression ratios on serial
data streams with minimal loss of scientific information.
JPL
Data Management for Mars Exploration Rovers
Data Management for the Mars Exploration Rovers (MER) project is a comprehensive system addressing the needs of development, test, and operations phases of the
mission. During development of flight software, including the science software, the data management system can be simulated using any POSIX file system. During
testing, the on-board file system can be bit compared with files on the ground to verify proper behavior and end-to-end data flows. During mission operations, end-to-end
accountability of data products is supported, from science observation concept to data products within the permanent ground repository. Automated and human-in-theloop ground tools allow decisions regarding retransmitting, re-prioritizing, and deleting data products to be made using higher level information than is available to a
protocol-stack approach such as the CCSDS File Delivery Protocol (CFDP).
GSFC
Onboard Processor for Compressing HSI Data
JPL
Science-based Region-of-Interest Image
Compression
With EO-1 Hyperion and MightySat in orbit NASA and the DoD are showing their continued commitment to hyperspectral imaging (HSI). As HSI sensor technology
continues to mature, the ever-increasing amounts of sensor data generated will result in a need for more cost effective communication and data handling systems.
Lockheed Martin, with considerable experience in spacecraft design and developing special purpose onboard processors, has teamed with Applied Signal and Image
Technology (ASIT), who has an extensive heritage in HSI, to develop a real-time and intelligent onboard processing (OBP) system to reduce HSI sensor downlink
requirements. Our goal is to reduce the downlink requirement by a factor greater than 100, while retaining the necessary spectral fidelity of the sensor data needed to
satisfy the many science, military, and intelligence goals of these systems. Our initial spectral compression experiments leverage commercial-off-the-shelf (COTS) spectral
exploitation algorithms for segmentation, material identification and spectral compression that ASIT has developed. ASIT will also support the modification and integration
of this COTS software into the OBP. Other commercially available COTS software for spatial compression will also be employed as part of the overall compression
processing sequence. Over the next year elements of a high-performance reconfigurable OBP will be developed to implement proven preprocessing steps that distill the
HSI data stream in both spectral and spatial dimensions. The system will intelligently reduce the volume of data that must be stored, transmitted to the ground, and
processed while minimizing the loss of information.
As the number of currently active space missions increases, so does competition for Deep Space Network (DSN) resources. Even given unbounded DSN time, power and
weight constraints onboard the spacecraft limit the maximum possible data transmission rate. These factors highlight a critical need for very effective data compression
schemes. Images tend to be the most bandwidth-intensive data, so image compression methods are particularly valuable. In this paper, we describe a method for
prioritizing regions in an image based on their scientific value. Using a wavelet compression method that can incorporate priority information, we ensure that the highest
priority regions are transmitted with the highest fidelity.
GSFC
SDS data compression recommendations
Development and status
The Consultative Committee for Space Data Systems (CCSDS) has been engaging in recommending data compression standards for space applications. The first effort
focused on a lossless scheme that was adopted in 1997. Since then, space missions benefiting from this recommendation range from deep space probes to near Earth
observatories. The cost savings result not only from reduced onboard storage and reduced bandwidth, but also in ground archive of mission data. In many instances, this
recommendation also enables more science data to be collected for added scientific value. Since 1998, the compression sub-panel of CCSDS has been investigating
lossy image compression schemes and is currently working towards a common solution for a single recommendation. The recommendation will fulfill the requirements for
remote sensing conducted on space platforms.
Smart Autonomous Command and Data Handling System, Algorithms and Data Management
Information – Data Acquisition and End-to-End Management
WIPO
Patent Number
WO02060106A2
Title
WO02067429A2
System and Method for Enhanced
Error Correction in Trellis Decoding,
Also Unequal Error Protection of
Variable-Length Data Packets Based
on Recursive Systematic
Convolutional Coding, also FeedbackBased Unequal Error Protection
Coding
System Initialization of MicrocodeBased Memory Built-In Self-Test
Cute Ltd.
System and Method for Encoding and
Decoding Data Files
Cyber Operations,
LLC
WO0208904A2
WO02093358A1
System and Method for Data ReCompression for Communication Over
Ip
Assignee
Flash Networks
Ltd.
International
Business Machines
Corporation, IBM
United Kingdom
Limited
Abstract
The transmission of image data through a network or other transmission medium is bandwidth intensive. To help reduce the amount of
bandwidth utilized for the transmission of image data, compression algorithms are used. However, even once compressed, if a large
amount of image data is being transferred, such as in a multimode network that shares graphic intensive information (i.e., the world wide
web) bandwidth can be at a premium. The present invention provides a technique for receiving multiple image files from multiple sources
and applying further compression, or recompression to help alleviate the bandwidth constraints. The present invention operates to further
compress image files on-the-fly so that processing delays for the retransmission of the image files are limited. To effectuate the on-the-fly
re-compression, image structures to indicate the status of re-compressing the various image files are maintained. Thus, segments of the
image files can be re-compressed as they are received regardless of whether the reception of segments from various image files are
received in an interleaved manner.
An encoding device for encoding a real time data stream for transfer over a noisy channel, the data stream comprising data bits in a
succession of data packets, the system comprising: a data transmitter for sending said data bits in said packets in a utilization order, a data
interleaver for interleaving said data bits into an interleaved order, and an encoder for encoding said data bits in said interleaved order to
form parity bits for insertion into said data stream as a parity set, such that said parity bits are differentially distributed over said packets
from said data bits.
The functionality of a programmable memory built-in self-test (BIST) arrangement for testing an embedded memory structure of an
integrated circuit is extended to system level testing to ascertain operability of the system after the integrated circuits and boards including
them have been placed in service in larger systems, by generating default test signals which are loaded in an instruction store module when
test instructions are not provided from an external tester. This additional utility of the BIST arrangement, increases efficiency of chip
space utilization and improves the system level test. Loading of test instructions from an external tester during chip manufacture and/or
board assembly is unaffected.
Distributed compression of a data file can comprise a master server module for breaking the data file into data blocks and transmitting the
data blocks to worker server modules. A first worker server module can compress a first data block using a first compression algorithm,
resulting in a first compressed data block. A second worker server module can compress the second data block using a second
compression algorithm, resulting in a second compressed data block. The first and second compression algorithmscan comprise the same
algorithm or different algorithms. An archive module can save the first and second compressed data blocks in an archive file for storage or
for transmission over a communication network. The worker server modules also can compress the respective data blocks using multiple
compression algorithms and can choose the highest compressed result (Figure 1, 100, 102, 110, 112, 114, 116, 118, 120, 122, 124, 126,
127, 128, 129).
Smart Autonomous Command and Data Handling System, Algorithms and Data Management
Information – Data Acquisition and End-to-End Management
WIPO
Patent Number
WO0241498A2
Title
WO03001748A1
Method and Apparatus for
Compression and Decompression of
Data
Ziplabs Pte Ltd.
WO03044963A1
Data Compression Method and System
California Institute
of Technology
WO03056704A1
Method and Apparatus for Adapting
Data Compression According to
Previous Exchange of Information
Nokia Corporation
WO03069873A2
Audio Enhancement Communication
Techniques
Tellabs Operations,
Inc.
Communication System and Method
Utilizing Request-Reply
Communication Patterns for Data
Compression
Assignee
Telefonaktiebolaget
Lm Ericsson (Publ)
Abstract
A method, system, and apparatus for increasing the efficiency and robustness of the compression of a communication protocol for
communication between entities (210, 230, 410, 440) over bandwidth limited communication links. The present invention uses the
request-reply nature of communication protocols to update compression and decompression dictionaries (220, 240, 420, 430, 450, 460).
Each communication entity (210, 230, 410, 440) will update its dictionary (220, 240, 420, 430, 450, 460) with a new message as soon as it
is known that the other communication entity has access to teh message. In one aspect of the present invention, an entity (210, 230, 410,
440) updates a compression/decompression dictionary (220, 240, 420, 430, 450, 460) by updating the dictionary with sent messages as
soon as a reply arrives from the other entity (210, 230, 410, 440), and by updating the dictionary (220, 240, 420, 430, 450, 460) with
received messages immediately. In another aspect of the present invention, received messages are used to update an entity's decompression
dictionary (430, 450) and sent messages are used to update an entity's compression dictionary (420, 460).
A real time compression and decompression client-server architecture over bandwidth limited data communications network uses a client
application (120) located between a client (130) and a bandwidth limited network (110) and a server (100) located between the bandwidth
limited network (110) and an Internet server to send and receive compressed data in an efficient manner across the bandwidth limited
network. The native web object formats at the Internet Server are replaced and transformed at the application to a more efficient lossless or
lossy compression format. The client application (130) performs the task of decompressing format. The client application (130) performs
the task of decompressing and decrypting the new compressed object format from the bandwidth limited network back to the native web
object format.
A signal encoding/decoding system and method are provided, where a signal quantizer is designed by minimizing a target function in
accordance with a rate/distortion tradeoff. The system and method can be applied to give lossy compression of data for graphic or music
files and also to smart-server applications, in accordance with the available band.
A method and apparatuses are disclosed for compressing digital data before transmitting it from a communications device of a first party
to a communications device of a second party. There is maintained (302, 303) a collection of digital data (105, 115) that represents an
archive of information (104,114, 201) that has been exchanged between the first party and the second party. When new information (202)
to be transmitted from the communications device of the first party to the communications device of the second party is composed (308), it
is represented as digital data. Said digital data is compressed (311) by utilizing correspondences between the new information (202), so
that at least some of the resulting compressed digital data (203) refers to information in said archive (201).
A communication system (10) receives a communication signal comprising first and second data with different compression levels, such
as highly compressed and weakly compressed levels. A mode detector (15) detects the level of compression. One or more signal decoders
(20, 22) decode the highly compressed data. An analyzer (30) determines the type of enhancement required. One or more processors (48,
50, 80) enhance the data as required. An encoder (60) reencodes the enhanced decoded data. Metrics (90) may aid the operation of the
analyzer (30). The communication system may include telephones (120, 122, 124, 126). Processors (103, 104) enhance signals in opposite
first and second directions between pairs of the telephones. A path (106) connects the processors in tandem. One or more switches (101,
102) disable signal enhancement for one of the processors depending on the compression level of the signals to avoid degrading call
quality.
Smart Autonomous Command and Data Handling System, Algorithms and Data Management
Information – Data Acquisition and End-to-End Management
WIPO
Patent Number
WO03084205A2
Title
WO04008672A2
Wireless Communications System
Having Built-In Packet Data
Corporation and Support for Enabling
Non-Standard Features Between
Network Elements
Methods and Apparatus for Network
Signal Aggregation and Bandwidth
Reduction
Nokia Corporation,
Nokia Inc.
NMS
Communications
Wireless network (10) demands continually increase as wireless service providers pursue additional service capabilities. In a cellular
communication system (10), leased lines between remote cell sites (22) and the corresponding Mobile Switching Offices (MSOs) (29)
remain a major operating cost. Bandwidth reduction by identification and elimination of payload data and control information which need
not be fully replicated because it can be deduced from information accessible or previously transmitted allows fewer lines to support the
same bandwidth. A wireless access gateway (30) is operable to aggregate such redundant and regenerable data on a backhaul link between
a wireless cell site and the corresponding mobile switching office (MSO) (29) to provide low-latency, type specific lossless bandwidth
reduction. The wireless access gateway (30) identifies regenerable information and eliminates portions of the data which the device need
not transmit because the data is redundant, or accessible or recreatable, at the receiving side.
A Method and System to Compress
and to Decompress Data
Van Gucht, Jurgen
Compressing data by sorting these data according to a specific order. The sorted data is compressed into a dictionary. The comparison
results obtained during the sorting of the data are stored. Decompressing the compressed data using the compressed dictionary and the
stored comparison results.
WO04014001A1
WO04109931A1
Repetition Coded Compression for
Highly Correlated Image Data
Assignee
Matrixview Pte.
Ltd.
Abstract
This invention is related to both a process and a system for compressing highly correlated image data. The system for compressing image
and other highly correlated data comprises means for capturing the image, means for converting to digital form, means for reshaping the
data, means for encoding the repetitions, means for storing the compressed data and means for retrieving the data. The method for
compressing image and other highly correlated data comprises of steps like capturing the image, converting into digital form, reshaping
the data into matrix form, encoding the repetitions into a bit-plane index and stored data values, storing the compressed data in storage
memory and retrieving the data for decompression. The system and method for compressing image and other highly correlated data is
described in the description and illustrated by the way of drawings
A method for performing data compression in a wireless communications system, comprising receiving an original data packet at a
network node (30) located between a packet data network (35) and an Internet Protocol network (70) of the wireless communications
system, compressing the original data packet at the network node, and transmitting the compressed original data packet through the IP
network (70) and through a wireless link to a mobile station (100).
Smart Autonomous Command and Data Handling System, Algorithms and Data Management
Information – Data Acquisition and End-to-End Management
WIPO
Patent Number
WO05004061A1
Title
WO05022399A2
Data Compression Engines and RealTime Wideband Compressor for
Multi-Dimensional Data
Canadian Space
Agency
WO05043766A1
Method and System for Loss-Less
Data Compression
Sunsail
Development Avv
Method for lossless compression of an original dataset (D l) in a binary format, including a sequence of a first number (N1) of set bits and
a second number (NO) of clear bits, wherein the method: - determines a ranking number (RN) of the sequence of set bits and clear bits in
the original dataset (D1) relative to a number (TC) of possible combinations including the first number (N1) of set bits and the second
number (NO) of clear bits, and - generates an encoded dataset (D2) which includes the ranking number (RN), the ranking number (RN)
being a lossless compressed representation of the original dataset (D 1).
WO05057937A1
Compressing Image Data
Matrixview Limited
A method for compressing image data of an image, comprising; transforming the image data into a bit plane of first and second values;
comparing each image element with a previous image element and if they are equal; recording a first value into a bit plane; and if they are
not equal, recording a second value into the bit plane; and encoding repeating first and second values in the bit plane into a bit plane index;
wherein the compressed image is able to be decompressed using the bit plane index and the bit plane.
Method of Coding a Continuous Data
FlowUSing Vector Quantization
Assignee
Canadian Space
Agency
Abstract
The present invention relates to a method and system for compressing a continuous data flow in real-time based on lossy compression. In
real-time data compression, a series of multi-dimensional data subsets acquired in a given period of time are treated as a regional data cube
for the purpose of dividing a continuous series of data subsets into a plurality of data cubes. In a first embodiment implementation of
parallel processing using a plurality of compression engines is facilitated by separating a data cube into a plurality of clusters comprising
similar spectral vectors. By separating the data cube into clusters of similar spectral vectors no artificial spatial boundaries are introduced
substantially improving image quality. Furthermore, the spectral vectors within a cluster are more easily compressed due to their
similarity. In a second embodiment a predetermined number of 2D focal plane frames in a boundary area of a previous regional data cube
close to a current regional data cube are included in a training set used for codevector training for the current region. Therefore, no
artificial boundary occurs between the two adjacent regions when codevectors trained in this way are used for codebook generation and
encoding of the spectral vectors of the current regional data cube substantially reducing image artifacts between adjacent regions. A
remedy for the single bit error problem is provided in a third embodiment. Full redundancy of compressed data for a regional data cube is
obtained by combining the previous regional data cube and the current regional data cube for codebook training. In order to obtain
redundancy for the index map, the codebook is used to encode the current regional data cube as well as the previous regional data cube
producing a baseline index map for the current regional data cube and a redundant index map for the previous regional data cube.
Therefore, full redundancy for a regional data cube is provided allowing restoration of a regional data cube if its codebook and/or index
map are corrupted or lost due to single bit errors.
The present invention relates to a real-time wideband compressor for multi-dimensional data. The compressor comprises a plurality of
compression engines for simultaneously compressing a plurality of data subsets of a set of input data vectors and providing compressed
data thereof using one of SAMVQ or HSOCVQ data compression. Each compression engine comprises an along spectral vectors
codevector trainer as well as an across spectral bands codevector trainer. The compression engines are programmable to perform either
along spectral vectors codevector training or across spectral bands codevector training in combination with one of the SAMVQ or
HSOCVQ techniques without changing hardware. The compressor further comprises a network switch for partitioning the set of input data
vectors into the plurality of data subsets, for providing each of the plurality of data subsets to one of the plurality of compression engines,
and for transmitting the compressed data. The real-time wideband compressor is highly advantageous in, for example, space applications
by programmable enabling performance of different techniques of codevector training as well as different techniques of VQ. Furthermore,
after the compression process is started the compression process is performed autonomously without external communication.
Smart Autonomous Command and Data Handling System, Algorithms and Data Management
Information – Data Acquisition and End-to-End Management
WIPO
Patent Number
WO06017382A1
Title
WO06060224A1
Golomb-Rice Lossless Compression of Honeywell
Satellite Images
International Inc.
A Rice coding data compression module includes a memory interface operable to receive sensor data from memory, a data normalization
module operable to normalize received sensor data, an encoder operable to apply a Rice compression algorithm to the normalized data to
produce compressed sensor data, a data management module operable to apply packet formatting to the compressed sensor data to produce
formatted compressed sensor data packets; and a memory interface operable to store the formatted compressed sensor data packets to
memory.
WO07019388A2
Data Compression and Abnormal
Situation Detection in a Wireless
Sensor Network
Honeywell
International Inc.
Wireless communication systems adapted for compressing data prior to certain communications. Data compression may be limited or
skipped when it is determined that the data compression may cause an unacceptable amount of data to be lost. Abnormal situation
detection as part of data compression is included. Methods associated with such systems are also encompassed.
WO07071541A1
Providing an Independent
Compression Server Within a
Network, As Well As a Method,
Network Station and Dhcp Server
Thomson Licensing
WO07078600A2
Digital Image ReconstructionUSing
Inverse Spatial Filtering
Eastman Kodak
Company
WO07117970A2
System and Method for Enhancing
Data Speed Over Communication
Lines
Sbc Knowledge
Ventures, L.P.
The invention is related with the problem of utilizing data compression in a network of distributed stations. Often header compression is
used to improve the bandwidth usage in networks, in particular wireless networks. Header compression could be implemented in access
points or routers, but both implementations have serious problems, e.g. due to limited CPU power, lack of scalability, or handover latency.
To resolve the problems the invention proposes to use a dedicated data compression server in the network and a new protocol to
transparently deploy data compression in the network.
A method for processing a source digital image wherein the source digital image is comprised of a plurality of pixels. A spatial filter is
applied to the source digital image to produce an enhanced digital image. An inverse spatial filter is applied to the enhanced digital image
to produce an estimated digital image. A difference digital image is then produced from the estimated digital image and the source digital
image, wherein the difference digital image is representative of a difference between the source digital image and the estimated digital
image. The difference digital image and the enhanced digital image can be transmitted from a first device to a second device remote. At
the second device, a reconstructed digital image can be generated from the difference digital image and the enhanced digital image,
wherein the reconstructed digital image is substantially equivalent to the source digital image.
The present disclosure provides a system and method for compressing data in a communications network. The downstream data is
compressed at a network element and sent to a customer premises over a data communication link, and is decompressed by a device at the
customer end. The upstream data is compressed by the device at the customer end and decompressed by the network element.
Methods and Apparatus for
Communicating and Displaying
Compressed Image Data
Assignee
Electronics for
Imaging, Inc.
Abstract
Methods and apparatus are provided for creating one or more compressed image tiles based on a compressed file that describes a digital
image. In particular, the compressed image tiles are created without fully decompressing the compressed file. Each compressed image tile
includes data corresponding to a portion of the digital image, and is independent of other compressed image tiles (i.e., may be
decompressed without decompressing any other tile). In response to requests to display a desired portion of the digital image at a specific
resolution, the compressed image tiles corresponding to the desired portion and the specified resolution are communicated via a band
limited communication channel. In this regard, the portions of the digital image may be quickly communicated and displayed, without
having to wait for the entire compressed file to be communicated over the band limited channel.
Smart Autonomous Command and Data Handling System, Algorithms and Data Management
Information – Data Acquisition and End-to-End Management
WIPO
Patent Number
WO07149358A1
Title
WO28004837A1
Apparatus and Method for Estimating
Compression Modes for H.264
Codings
Data Compression
Assignee
Essex Pa, L.L.C.
Abstract
The present invention concerns a method of compressing a grouping of associated data. The method performs the steps of (a) selecting a
symbol string that occurs within the data to be compressed, and generating a symbol string code indicative of one or more positions of the
symbol string within the grouping of data to be compressed; (b) successively repeating (a) for further symbol strings that occur within the
grouping; and (c) combining respective symbol string codes into a compressed data code.
Libertron Co., Ltd.,
Konkuk University
Industrial
Cooperation Corp
The present invention relates to technology of compressing large image data in order to effectively use a storage medium and efficiently
use a communication medium in an image data compression technique field, and more particularly, to fast compression mode calculation
in H.264. The present invention provides an H.264 compression mode estimation apparatus comprising: a macroblock image characteristic
calculator calculating a macroblock image characteristic from macroblock image information; a macroblock mode estimator which
comprises mode history tables and a table management unit managing these tables, in order to estimate a mode of a current macroblock
from macroblock coordinates; and a mode estimation optimization judgment unit calculating RD-cost for the mode estimation and
selecting an optimal mode in order to prevent the spread of errors of the estimated mode.
Smart Autonomous Command and Data Handling System, Algorithms and Data Management
Information – Human-Computer Interfaces
Company Name
Distributed Simulation Technology Inc.
Title
OPENGL GRAPHICS FOR ROTORCRAFT
DISPLAYS
DOD SBIR Phase II
Quad Chart
This effort will develop a rapid application development and virtual prototyping tool that automatically generates reusable object oriented
source code for use in safety critical avionics display systems. Object oriented development will reduce development costs, program
schedules and risk for new rotorcraft displays and will increase safety through a reduction in software defects.
Smart Autonomous Command and Data Handling System, Algorithms and Data Management
Information – Human-Computer Interfaces
FC
Goddard Space Flight Center
Title
REACH Real-Time Data Awareness in MultiSpacecraft Missions
NTRS
Abstract
Missions have been proposed that will use multiple spacecraft to perform scientific or commercial tasks. Indeed, in the commercial world,
some spacecraft constellations already exist. Aside from the technical challenges of constructing and flying these missions, there is also the
financial challenge presented by the tradition model of the flight operations team (FOT) when it is applied to a constellation mission. Proposed
constellation missions range in size from three spacecraft to more than 50. If the current ratio of three-to-five FOT personnel per spacecraft is
maintained, the size of the FOT becomes cost prohibitive. The Advanced Architectures and Automation Branch at the Goddard Space Flight
Center (GSFC Code 588) saw the potential to reduce the cost of these missions by creating new user interfaces to the ground system healthand-safety data. The goal is to enable a smaller FOT to remain aware and responsive to the increased amount of ground system information in
a multi-spacecraft environment. Rather than abandon the tried and true, these interfaces were developed to run alongside existing ground
system software to provide additional support to the FOT. These new user interfaces have been combined in a tool called REACH. REACHthe Real-time Evaluation and Analysis of Consolidated Health-is a software product that uses advanced visualization techniques to make
spacecraft anomalies easy to spot, no matter how many spacecraft are in the constellation. REACH reads a real-time stream of data from the
ground system and displays it to the FOT such that anomalies are easy to pick out and investigate. Data visualization has been used in ground
system operations for many years. To provide a unique visualization tool, we developed a unique source of data to visualize the REACH
Health Model Engine. The Health Model Engine is rule-based software that receives real-time telemetry information and outputs 'health'
information related to the subsystems and spacecraft that the telemetry belong to. The Health Engine can run out-of-the-box or can be tailored
with a scripting language. Out of the box, it uses limit violations to determine the health of subsystems and spacecraft when tailored, it
determines health using equations combining the values and limits of any telemetry in the spacecraft. The REACH visualizations then 'roll up'
the information from the Health Engine into high level, summary displays. These summary visualizations can be 'zoomed' into for increasing
levels of detail. Currently REACH is installed in the Small Explorer (SMEX) lab at GSFC, and is monitoring three of their five spacecraft. We
are scheduled to install REACH in the Mid-sized Explorer (MIDEX) lab, which will allow us to monitor up to six more spacecraft. The process of
installing and using our 'research' software in an operational environment has provided many insights into which parts of REACH are a step
forward and which of our ideas are missteps. Our paper explores both the new concepts in spacecraft health-and-safety visualization, the
difficulties of such systems in the operational environment, and the cost and safety issues of multi-spacecraft missions.
Smart Autonomous Command and Data Handling System, Algorithms and Data Management
Information –Software Development Environments
Company Name
GrammaTech, Inc.
Title
STATIC DETECTION OF BUGS
IN EMBEDDED SOFTWARE
USING LIGHTWEIGHT
VERIFICATION
Field Center
ARC
NASA SBIR Phase II
Quad Chart
By combining lightweight verification techniques with other scalable analysis techniques that target syntactic properties we will create a tool that
flags violations for almost all rules typically applied to high assurance code.
Smart Autonomous Command and Data Handling System, Algorithms and Data Management
Information –Software Development Environments
Company Name
Interface & Control Systems, Inc.
Title
Sleuth & Guide-Auto Checkout and Tasking for PnP
DOD SBIR Phase I
Quad Chart
The problem of long and costly development cycles for satellite command and control flight software is not unique to the AFRL's TacSat
Programs. In order to shorten the time-to-flight for spacecraft software systems, Interface and Control Systems has devised a software
ontology which will greatly reduce the development and validation time required to field new flight software systems on Spacecraft. Our Phase
1 SLEUTH and GUIDE products will automate the discovery of device and service capabilities and adapt the onboard system based on the
information provided by SPA compatible, plug and play avionics hardware and software. The proposed architecture will replace labor
intensive, hand-generated logic with auto-generated embedded systems code in an extremely cost-effective manner. The round trip time to
sense PnP device capabilities, generate code, load the code to the target processor, and validate functionality will be reduced to minutes.
Auto-coding tools will allow rapid reconfiguration of algorithms and flight profiles. ICS will demonstrate a highly automated test and checkout
capability with the Sleuth architecture and provide an advanced planning and tasking capability (Guide) which shares a common infrastructure
for definition of relationships between onboard sensors and subsystems.
Smart Autonomous Command and Data Handling System, Algorithms and Data Management
Information –Software Development Environments
Company Name
Cognitive Concepts
Title
MODEL-BASED SPECIFICATION AND TESTING FOR
RAPID SOFTWARE TRANSITION
DOD SBIR Phase II
Quad Chart
Model-Based Testing with the SpecTest toolset provides user with methods and tools that enable them to achieve reliability and safety
requirements while mitigating cost and schedule risk. This will help alleviate mission critical failures as well as schedule delays and
cost overruns.
Smart Autonomous Command and Data Handling System, Algorithms and Data Management
Information –Software Tools for Distributed Analysis and Simulation
FC
JPL
Title
Scalable collaborative risk management
technology for complex critical systems
NTRS
Abstract
We describe here our project and plans to develop methods, software tools, and infrastructure tools to address challenges relating to geographically distributed software
development. Specifically, this work is creating an infrastructure that supports applications working over distributed geographical and organizational domains and is using
this infrastructure to develop a tool that supports project development using risk management and analysis techniques where the participants are not collocated.
Smart Autonomous Command and Data Handling System, Algorithms and Data Management
Propulsion – Micro Thrusters
FC
JPL
Title
Colloid micro-Newton thruster development for the ST7DRS and LISA missions
NTRS
Abstract
We present recent progress and development of the Busek Colloid Micro-Newton Thruster (CMNT) for the Space Technology 7 Disturbance
Reduction System (ST7-DRS) and Laser Interferometer Space Antenna (LISA) Missions.
Smart Autonomous Command and Data Handling System, Algorithms and Data Management
Robotics – Integrated Robotic Concepts and Systems
FC
JSC
Title
AERCam Autonomy Intelligent Software Architecture for
Robotic Free Flying Nanosatellite Inspection Vehicles
NTRS
Abstract
The NASA Johnson Space Center has developed a nanosatellite-class Free Flyer intended for future external inspection and remote
viewing of human spacecraft. The Miniature Autonomous Extravehicular Robotic Camera (Mini AERCam) technology demonstration unit
has been integrated into the approximate form and function of a flight system. The spherical Mini AERCam Free Flyer is 7.5 inches in
diameter and weighs approximately 10 pounds, yet it incorporates significant additional capabilities compared to the 35-pound, 14-inch
diameter AERCam Sprint that flew as a Shuttle flight experiment in 1997. Mini AERCam hosts a full suite of miniaturized avionics,
instrumentation, communications, navigation, power, propulsion, and imaging subsystems, including digital video cameras and a high
resolution still image camera. The vehicle is designed for either remotely piloted operations or supervised autonomous operations,
including automatic stationkeeping, point-to-point maneuvering, and waypoint tracking. The Mini AERCam Free Flyer is accompanied by a
sophisticated control station for command and control, as well as a docking system for automated deployment, docking, and recharge at a
parent spacecraft. Free Flyer functional testing has been conducted successfully on both an airbearing table and in a six-degree-offreedom closed-loop orbital simulation with avionics hardware in the loop. Mini AERCam aims to provide beneficial on-orbit views that
cannot be obtained from fixed cameras, cameras on robotic manipulators, or cameras carried by crewmembers during extravehicular
activities (EVA s). On Shuttle or International Space Station (ISS), for example, Mini AERCam could support external robotic operations
by supplying orthogonal views to the intravehicular activity (IVA) robotic operator, supply views of EVA operations to IVA and or ground
crews monitoring the EVA, and carry out independent visual inspections of areas of interest around the spacecraft. To enable these future
benefits with minimal impact on IVA operators and ground controllers, the Mini AERCam system architecture incorporates intelligent
systems attributes that support various autonomous capabilities. 1) A robust command sequencer enables task-level command scripting.
Command scripting is employed for operations such as automatic inspection scans over a region of interest, and operator-hands-off
automated docking. 2) A system manager built on the same expert-system software as the command sequencer provides detection and
smart-response capability for potential system-level anomalies, like loss of communications between the Free Flyer and control station. 3)
An AERCam dynamics manager provides nominal and off-nominal management of guidance, navigation, and control (GN and ampC)
functions. It is employed for safe trajectory monitoring, contingency maneuvering, and related roles. This paper will describe these
architectural components of Mini AERCam autonomy, as well as the interaction of these elements with a human operator during
supervised autonomous control.
Smart Autonomous Command and Data Handling System, Algorithms and Data Management
Verification and Validation – Operations Concepts and Requirements
FC
JSC
Title
An Introduction to Flight Software
Development FSW Today, FSW
2010
JPL
Design and architecture of the Mars
relay network planning and analysis
framework
NTRS
Abstract
Experience and knowledge gained from ongoing maintenance of Space Shuttle Flight Software and new development projects including Cockpit Avionics Upgrade are applied to projected
needs of the National Space Exploration Vision through Spiral 2. Lessons learned from these current activities are applied to create a sustainable, reliable model for development of critical
software to support Project Constellation. This presentation introduces the technologies, methodologies, and infrastructure needed to produce and sustain high quality software. It will
propose what is needed to support a Vision for Space Exploration that places demands on the innovation and productivity needed to support future space exploration. The technologies in
use today within FSW development include tools that provide requirements tracking, integrated change management, modeling and simulation software. Specific challenges that have been
met include the introduction and integration of Commercial Off the Shelf (COTS) Real Time Operating System for critical functions. Though technology prediction has proved to be
imprecise, Project Constellation requirements will need continued integration of new technology with evolving methodologies and changing project infrastructure. Targets for continued
technology investment are integrated health monitoring and management, self healing software, standard payload interfaces, autonomous operation, and improvements in training.
Emulation of the target hardware will also allow significant streamlining of development and testing. The methodologies in use today for FSW development are object oriented UML design,
iterative development using independent components, as well as rapid prototyping . In addition, Lean Six Sigma and CMMI play a critical role in the quality and efficiency of the workforce
processes. Over the next six years, we expect these methodologies to merge with other improvements into a consolidated office culture with all processes being guided by automated office
assistants. The infrastructure in use today includes strict software development and configuration management procedures, including strong control of resource management and critical
skills coverage. This will evolve to a fully integrated staff organization with efficient and effective communication throughout all levels guided by a Mission-Systems Architecture framework
with focus on risk management and attention toward inevitable product obsolescence. This infrastructure of computing equipment, software and processes will itself be subject to
technological change and need for management of change and improvement,
In this paper we describe the design and architecture of the Mars Network planning and analysis framework that supports generation and validation of efficient planning and scheduling
strategy. The goals are to minimize the transmitting time, minimize the delaying time, and or maximize the network throughputs. The proposed framework would require (1) a client-server
architecture to support interactive, batch, WEB, and distributed analysis and planning applications for the relay network analysis scheme, (2) a high-fidelity modeling and simulation
environment that expresses link capabilities between spacecraft to spacecraft and spacecraft to Earth stations as time-varying resources, and spacecraft activities, link priority, Solar System
dynamic events, the laws of orbital mechanics, and other limiting factors as spacecraft power and thermal constraints, (3) an optimization methodology that casts the resource and constraint
models into a standard linear and nonlinear constrained optimization problem that lends itself to commercial off-the-shelf (COTS)planning and scheduling algorithms.
Smart Autonomous Command and Data Handling System, Algorithms and Data Management
Verification and Validation – Simulation Modeling Environment
Company Name
Advatech Pacific Inc
Title
Virtual Satellite Integration Environment
Field Center
ARC
NASA SBIR Phase I
Quad Chart
Advatech Pacific proposes to develop a Virtual Satellite Integration Environment (VSIE) for the NASA Ames Mission Design Center. The
VSIE introduces into NASA Mission Design Commercial Off-The-Shelf (COTS) Product Lifecycle Management (PLM) tools and
processes, which haven proven themselves in the industrial manufacturing world. In addition to COTS PLM tools, the VSIE hinges on
two key concepts: An enhanced Digital Mock-Up (termed DMU++) and the so-called Common Geometry Strategy. DMU typically
addresses mechanical form, fit and function of a component or sub-assembly in the assembly context. We propose to go one step further
and address electrical power, data, and other similar interfaces in the DMU to automatically detect compatibility issues beyond mere
mechanical fit. This will be an enabling functionality for rapid mission design and integration of components from a database of existing
off-the-shelf hardware such as the database currently under development at the NASA Ames Mission Design Center. The Common
Geometry Strategy was introduced in the late 1990s in both the commercial aircraft engine and automotive industries, however, it has so
far not found its way into satellite design or satellite mission design. The fundamental idea is that the same Master Model geometric
information is readily available to all disciplines and individuals that need geometric information to perform their job. Since a particular
disciplinary specialist may require only an abstraction of the detailed 3D-geometry, the idea of a "Context Model" is introduced. The
Context Model is a simplified representation of the detailed 3D CAD model, which is simplified precisely to the level of detail required by
the specialist, while maintaining full associativity to the Master Model geometry, so that it either automatically updates when the Master
Model geometry changes, or at least notifies the specialist that it is out of date.
Smart Autonomous Command and Data Handling System, Algorithms and Data Management
Verification and Validation – Simulation Modeling Environment
Company Name
Edaptive Computing, Inc.
Title
Software Formalisms
DOD SBIR Phase I
Quad Chart
Modern software development processes increasingly depend on complete specifications, the ability to execute the specifications expressed in a
lightweight formalism, and the ability to verify properties expressed in that formalism. EDAptive Computing, Inc. (ECI) team has designed the Model
Checking and Execution of Specifications using Lightweight Formalisms (ModSpec) program to deliver this combination. Our existing innovative
technology building blocks will be augmented with model checking, computerized assistants, and multi-domain component libraries developed in Rosetta,
a non-proprietary lightweight formal language. Specifically meeting the topic requirements this combination forms an end-to-end capability to enter or
capture a technology-independent specification, evaluate it for completeness, check it for correctness, and report the results in an intuitive form. We will
leverage our extensive experience with graphical systems entry, formal languages, formal methods in general, and model checking in particular. We will
demonstrate the feasibility of ModSpec by extending an experimental foundation first produced under a related MDA project, or a suitable mutually agreed
alternative. The Phase I result will clearly show that this powerful mix of innovative tools and mature methods will result in faster, more accurate, and more
error-free software development, integration, reuse, and deployment.
Smart Autonomous Command and Data Handling System, Algorithms and Data Management
Verification and Validation – Simulation Modeling Environment
FC
JPL
Title
On the validity of the double integrator approximation in
deep space formation flying
NTRS
Abstract
Free-flying models are commonly used f o r path planning and open loop control design (i. e., guidance design) and translational feedback control design (i. e.,
control design) for deep space precision formation flying. The free flying model, essentially a double integrator, results from discarding small terms in the
relative spacecraft equations of motion. While the magnitude of these discarded terms may be small, one must show that their dynamic effects are small as
compared to the precision performance requirements. We do so by deriving a theoretical method for bounding the difference between the solution of a
nonlinear truth model of the relative translational spacecraft dynamics and a Simplified linear time-invariant model. Presently, the method incorporates
feedforward and static output feedback control. The method is applied to a Terrestrial Planet Finder- based example. Using only feedforward control
(guidance) the free-flying model and a Hill- Clohessy- Wiltshire Equations-based model are shown to be accurate to 1 c m for up to 4 and 30 hours,
respectively. Also shown is that the simplest free-flying model may not be sufficient for low-gain feedback control design-closed-loop tracking errors can be as
large as 8 meters.
Smart Autonomous Command and Data Handling System, Algorithms and Data Management
Verification and Validation – Testing Requirements and Architectures
USPTO
Patent Number
Title
US6625758
Method and apparatus for automated system Advanced Micro
level testing
Devices, Inc.
Assignee
Abstract
A system-level (SLT) of a CPU device is performed in an automated test environment. Each device under test is automatically
placed an SLT station and a test is performed at an initial operating speed. A CPU device which passes the test is then
automatically removed and placed in a storage container based on that operating speed, also known as a rating (or rated) speed. If
the device fails the test, however, then it remains in the test station and the operating speed of the station is adjusted until the
device is able to pass the test. Once successful, the device is automatically removed and placed in a storage container based on the
operating speed at which it finally was successful. A device which is unable to pass a system-level test at any speed is
automatically removed and placed in a reject bin. This testing procedure is repeated for a number of devices without requiring
manual intervention to place the device in the SLT station, adjust the test operating speed, or binning the CPU device according to
its rated speed.
Smart Autonomous Command and Data Handling System, Algorithms and Data Management
Verification and Validation – Testing Requirements and Architectures
WIPO
Patent Number
WO0232071A2
Title
Method and Apparatus for Optimizing
Data Compression in a Wireless Digital
Access System
Assignee
Motorola,
Inc.
Abstract
A method and apparatus for optimizing data compression in a wireless digital access system (606) is described. The capability to establish a
data compression session (216) spanning two communication links (604 and 605) is provided. Common compression parameters
compatible with both communication links (604 and 605) are coordinated. Unnecessary processing is avoided, thereby reducing the
processing load of a control processor (204) used in conjunction with the invention. [
Smart Autonomous Command and Data Handling System, Algorithms and Data Management
Verification and Validation – Testing Requirements and Architectures
FC
DFRC
Title
Development and Evaluation of Fault-Tolerant
Flight Control Systems
JPL
Fault Injection Campaign for a Fault Tolerant
Duplex Framework
NTRS
Abstract
The research is concerned with developing a new approach to enhancing fault tolerance of flight control systems. The original motivation for fault-tolerant control
comes from the need for safe operation of control elements (e.g. actuators) in the event of hardware failures in high reliability systems. One such example is modem
space vehicle subjected to actuator sensor impairments. A major task in flight control is to revise the control policy to balance impairment detectability and to achieve
sufficient robustness. This involves careful selection of types and parameters of the controllers and the impairment detecting filters used. It also involves a decision,
upon the identification of some failures, on whether and how a control reconfiguration should take place in order to maintain a certain system performance level. In this
project new flight dynamic model under uncertain flight conditions is considered, in which the effects of both ramp and jump faults are reflected. Stabilization algorithms
based on neural network and adaptive method are derived. The control algorithms are shown to be effective in dealing with uncertain dynamics due to external
disturbances and unpredictable faults. The overall strategy is easy to set up and the computation involved is much less as compared with other strategies. Computer
simulation software is developed. A serious of simulation studies have been conducted with varying flight conditions.
Fault tolerance is an efficient approach adopted to avoid or reduce the damage of a system failure. In this work we present the results of a fault injection campaign we
conducted on the Duplex Framework (DF). The DF is a software developed by the UCLA group [1, 2] that uses a fault tolerant approach and allows to run two replicas
of the same process on two different nodes of a commercial off-the-shelf (COTS) computer cluster. A third process running on a different node, constantly monitors the
results computed by the two replicas, and eventually restarts the two replica processes if an inconsistency in their computation is detected. This approach is very cost
efficient and can be adopted to control processes on spacecrafts where the fault rate produced by cosmic rays is not very high.
Advanced Avionics
Avionics and Astrionics – Attitude Determination and Control
Company Name
Systems & Processes Engineering
Corp. (SP)
Title
Low Cost Satellite Inertial Measurement System
(LCSIM)
DOD SBIR Phase I
Quad Chart
Satellite Inertial Measurement (LCSIM) sensor suite which incorporates a grouping of sensors that are capable of both short-term lockdown
of the inherently quick moving objects, as well as long term positional and angular update capability to maintain zero drift orientation control.
In order to allow cost effective quick turn satellites for a variety of missions, there is a need for a miniature, low cost, low power, and
extremely capable inertial measurement unit. Recent advancements in MEMS sensors, GPS sensors and imaging sensors allow such a unit
to be constructed using Commercial Off The Shelf (COTS) components. By optimizing the synergy between the sensors, the strengths of
each sensor can be used to enhance the overall suite, compensating for sensor weaknesses. The result is a high speed, quick lock down,
ultra high accuracy suite made up of a large number of miniature sensors rather than a few ultra high accuracy (and high cost) ones. With
the miniature high accuracy system small satellites could be used, in place of some sensor suites now capable of flying only on larger
attitude controlled busses.
Advanced Avionics
Avionics and Astrionics – Attitude Determination and Control
DOD SBIR Phase II
Quad Chart
AeroAstro proposes to leverage its existing innovative miniature star-tracker design to develop an improved star tracker that provides both spacecraft
attitude and angular rate information. The resulting system will have minimal cost and impact to spacecraft resources and can easily handle typical
spacecraft tumble conditions and provide for recovery from a lost-in-space (LIS) condition.
Company Name
AeroAstro, Inc.
Title
FAST ANGULAR RATE MINIATURE STAR
TRACKER
Continental Controls and Design, Inc.
FOUR-DIMENSIONAL (4-D) OCEANOGRAPHIC
INSTRUMENTATION
In Phase 2, CCD proposes to build and test directional wave buoy prototypes that are smaller than a soda can and could cost less than $1,000 in
production. The heart of this new directional micro-buoy is a Tiny Guidance Engine (TGE) developed originally for smart munitions. In less than a
cubic inch, TGE measures all 6 degrees of freedom inertially, 3 axes magnetically, pressure, and temperature.
Microcosm, Inc.
PLUG-AND-PLAY INERTIAL MEASUREMENT
UNIT
In Phase II, the Microcosm team proposes to develop a MEMS IMU that will provide nanosatellites with a low-cost, low-power, small-footprint
navigation sensor that outputs information on par with LN-200 class IMUs. The MEMS IMU will be designed with plug-and-play (PnP) capability
consistent with proposed spacecraft PnP avionics (SPA) interface standards being developed at AFRL.
Advanced Avionics
Avionics and Astrionics – Attitude Determination and Control
U.S. Patent Applications
Patent Number
Title
Assignee
Abstract
US2005133670A1
Unified sensor-based attitude
determination and control for spacecraft
operations
None
Systems and method of attitude determination and control for spacecraft include the use of a unified set of sensors for all phases of
space flight. For example, the same set of sensors may be used for on-station operations and transfer orbit operations. The use of a
unified set of sensors reduces complexity, spacecraft cost, and spacecraft weight
US2007023579A1
Unified attitude control for spacecraft
transfer orbit operations
The Boeing Company
A method of controlling attitude of a spacecraft during a transfer orbit operation is provided. The method includes providing a
slow spin rate, determining the attitude of the spacecraft using a unified sensor set, and controlling the attitude of the spacecraft
using a unified control law. The use of a unified set of sensors and a unified control law reduces spacecraft complexity, cost, and
weight.
Advanced Avionics
Avionics and Astrionics – Attitude Determination and Control
USPTO
Patent
Number
US7136752
Title
Assignee
Abstract
Method and apparatus for onboard autonomous pair catalog
generation
The Boeing
Company
A system ( 18 ) includes: a) A vehicle ( 12 ) includes an attitude or angular velocity control system ( 38 ), a plurality of star trackers or star
sensors ( 22 ) each having a field of view ( 28 ); b) a memory ( 30 ) having a star catalog ( 32 ), an allocated area for a star pair catalog ( 58 ) and
a reference table ( 56 ) stored therein; and c) a processor ( 24 ) coupled to the attitude or angular velocity control system ( 38 ), the star trackers or
star sensors ( 22 ), and the memory ( 30 ). The processor ( 24 ) populates the star pair catalog ( 58 ), using the method described herein. The
processor ( 24 ) then periodically determines the vehicle inertial attitude or angular velocity or sensor alignment, based, in part, on the star pair
catalog ( 58 ) and reference table ( 56 ). The novel ability of the software to autonomously populate the star pair catalog ( 58 ) allows users to
avoid uploading a large amount of data, and the problems associated with such an upload.
US7228231
Multiple stay-out zones for
ground-based bright object
exclusion
The Boeing
Company
A vehicle ( 12 ) including a control system ( 18 ) is used for controlling vehicle attitude or angular velocity ( 38 ). The processor ( 24 ) is coupled
to a star sensor or tracker ( 22 ) and a memory ( 30 ) that may include a star catalog ( 32 ), and an exclusion list ( 36 ). The exclusion list ( 36 ), a
list of stars to be temporarily excluded from consideration when determining attitude or angular velocity or relative alignment of star sensors or
trackers, is calculated on board. Such a calculation prevents the necessity for a costly, periodic, ground calculation and upload of such data. By
manipulating the star catalog, or sub-catalogs derived from said catalog, based upon the exclusion list ( 36 ), measurements of such excluded stars
are prevented from corrupting the attitude or angular velocity or alignment estimates formulated on board. The system uses multiple stayout zones
for excluding stars from the exclusion list. A central exclusion zone excludes all stars while a second or more exclusion zones allow some stars to
be used in the attitude determination
US7310578
Fast access, low memory, pair
catalog
The Boeing
Company
A system ( 18 ) includes: a) A vehicle ( 12 ) includes an attitude or angular velocity control system ( 38 ), a plurality of star trackers or star
sensors ( 22 ) each having a field of view ( 28 ); b) a memory ( 30 ) having a star catalog ( 32 ), a star pair catalog ( 58 ) and a reference table ( 56
) stored therein; and c) a processor ( 24 ) coupled to the attitude or angular velocity control system ( 38 ), the star trackers or star sensors ( 22 ),
and the memory ( 30 ). The processor ( 24 ) determines the vehicle inertial attitude or angular velocity or sensor alignment, based, in part, on the
star pair catalog ( 58 ) and reference table ( 56 ). The design of the star pair catalog ( 58 ) and reference table ( 56 ) is suitable for rapid
determination of attitude or angular velocity or sensor alignment, and an efficient use of memory.
Advanced Avionics
Avionics and Astrionics – Attitude Determination and Control
WIPO
Patent Number
WO0206115A1
Title
WO04080133A2
Integrated Sensor And Electronics
Package
Flight Control Modules Merged
Into The Integrated Modular
Avionics
Assignee
Honeywell International Inc.
Abstract
IIn an aircraft using fly-by-wire technology, the flight control functions have been integrated into the integrated modular avionics
("IMA") (410). The new flight control module ("FCM") (402) resides on the same data bus as the other modules in the IMA and
receives power from the same power supply. In addition, the FCM is also connected to a separate power supply to add redundancy
to the system. Several benefits arise from this configuration of an FCM. There is no longer a separate chassis needed for the flight
control functions, thus resulting in a reduction in weight. In addition, the FCM now has access to all of the data on the IMA bus,
instead of a limited amount of data over an ARINC 629 bus. The FCM provides augmentation signals to the actuator control
electronics ("ACE") to aid in the flying of the aircraft. In the event of a failure of the FCM, the ACE still provides enough control
to fly the airplane.
The Charles Stark Draper
Laboratory, Inc.
An integrated sensor and electronics package wherein a micro-electromechanical sensor die is bonded to one side of the package
substrate, one or more electronic chips are bonded to an opposite side of the package substrate, internal electrical connections run
from the sensor die, through the package substrate, and to the one or more electronic chips, and input/output connections on the
package substrate are electrically connected to one or more of the electronic chips.
Advanced Avionics
Avionics and Astrionics – Attitude Determination and Control
FC
GSFC/JPL
Title
The Inertial Stellar Compass (ISC) A Multifunction, Low
Power, Attitude Determination Technology
Breakthrough
JPL
Isolated post resonator mesogyroscope
NTRS
Abstract
The Inertial Stellar Compass (ISC) is a miniature, low power, stellar inertial attitude determination system with an accuracy of better than 0.1 degree (1
sigma) in three axes. The ISC consumes only 3.5 Watts of power and is contained in a 2.5 kg package. With its embedded on-board processor, the ISC
provides attitude quaternion information and has Lost-in-Space (LIS) initialization capability. The attitude accuracy and LIS capability are provided by
combining a wide field of view Active Pixel Sensor (APS) star camera and Micro- ElectroMechanical System (MEMS) inertial sensor information in an
integrated sensor system. The performance and small form factor make the ISC a useful sensor for a wide range of missions. In particular, the ISC
represents an enabling, fully integrated, micro-satellite attitude determination system. Other applications include using the ISC as a single sensor
solution for attitude determination on medium performance spacecraft and as a bolt on independent safe-hold sensor or coarse acquisition sensor for
many other spacecraft. NASA's New Millennium Program (NMP) has selected the ISC technology for a Space Technology 6 (ST6) flight validation
experiment scheduled for 2004. NMP missions, such a s ST6, are intended to validate advanced technologies that have not flown in space in order to
reduce the risk associated with their infusion into future NASA missions. This paper describes the design, operation, and performance of the ISC and
outlines the technology validation plan. A number of mission applications for the ISC technology are highlighted, both for the baseline ST6 ISC
configuration and more ambitious applications where ISC hardware and software modifications would be required. These applications demonstrate the
wide range of Space and Earth Science missions that would benefit from infusion of the ISC technology.
A new symmetric vibratory gyroscope principle has been devised in which a central post proof mass is counter-rocked against an outer sensing plate
such that the motion is isolated from the gyroscope case. Prototype gyroscopes have been designed and fabricated with micromachined silicon at
mesoscale (20-cm resonator width), vs. microscale (e.g., 2-mm resonator width) to achieve higher sensitivity and machined precision. This novel
mesogyro design arose out of an ongoing technical cooperation between JPL and Boeing begun in 1997 to advance the design of micro-inertial sensors
for low-cost space applications. This paper describes the theory of operation of the mesogyro and relationships with other vibratory gyroscopes, the
mechanical design, closed loop electronics design, bulk silicon fabrication and packaged gyroscope assembly and test methods. The initial packaged
prototype test results are reported for what is believed to be the first silicon mesogyroscope.
Advanced Avionics
Avionics and Astrionics – Guidance Navigation and Control
Company Name
Milli Sensor Systems and
Actuators, Inc
Title
Stable Tactical-Grade MEMS IMU for SpinStabilized Rockets
Field Center
GSFC
Stellar Exploration, Inc.
Low-Cost Suite of COTS GNC Sensors for
Precision Lunar Lander
ARC
NASA SBIR Phase I
Quad Chart
An Integrated MEMS IMU is proposed that will operate effectively in a spinning rocket up to 7 revs/sec. The IMU contains three
gyroscopes and nine accelerometers on the same chip. The instrument designs have the low cross-axis sensitivity that permit the
orthogonal gyros to sense the relatively smaller pitch and yaw rates in the presence of the much larger rate about the spin axis. An
algorithm is proposed to combine the signals from the instruments to co-operatively obtain spatial reference. A lab experiment is
planned during Phase I that will use available equipment and MSSA IMUs to prove the concept.
We are proposing to exploit (in an innovative way) existing, readily available, GNC sensors for the purpose of precision lunar landing.
Majority of previous lunar lander concepts with the precision/pinpoint landing capability required expensive and risky development of
new GNC and landing sensors (scanning lidars, multi-beam mm-ww radar, etc.). Our proposed alternative consists solely of existing and
low-cost sensors that synergistically leverage each capability and compensate for individual sensor weaknesses. For example, we can
use a simple single-beam low-frequency radar altimeter (available at low-cost off-the-shelf, and proven on several Mars lander
missions). The low-frequency radar can meet the maximum slant range requirements much easier than the mm-wave sensor but it does
not have the adequate multiple narrow beam capability of the Apollo LM or Viking lander radar. However, the optical descent imaging
measurement (using DSMAC-type sensor) can supplement the single beam radar measurement and obtain the same information about
the complete state vector. There are several similar concepts implemented in this sensor suite of complementing strengths and
weakness of individual sensors.
Advanced Avionics
Avionics and Astrionics – Guidance Navigation and Control
Company Name
NAVSYS
Corporation
Title
Modular, Scalable Propulsion Module for ESPABased Satellite Dispensing Systems
Microcosm, Inc.
Innovative, Low Cost, Plug-and-Play Intertial
Measurement System
DOD SBIR Phase I
Quad Chart
Current GPS technologies for satellite navigation are costly, heavy, and utilize high amount of power. This makes such systems difficult for microsatellites to
support. Under this Phase I SBIR, we propose to develop a low-cost, low-weight, low-power GPS navigation system to support micro/nanosatellites. A key
component of our solution involves our patented TIDGET based receiver design, which takes a brief snapshot of GPS data and powers off until the next position
fix is desired. Also, we plan to implement our software defined radio (SDR) GPS waveform technology to process the recorded snapshots onboard using the
existing avionics processor, or as an option, downlink the data, when convenient, for ground-based processing. An important aspect of the TIDGET is it’s modular
size, which allow multiple TIDGETs to be placed on the satellite shell for full GPS visibility and robustness to satellite spin. Under this Phase I SBIR, we propose
to construct a prototype TIDGET receiver to collect representative simulated GPS RF data. The data will be processed via MATLAB for analysis. Furthermore, we
propose to develop a Phase II design for possible follow-on development of a ruggidized, space-qualified TIDGET with embedded avionics processing for the
ESPA module.
For this Phase I SBIR, Microcosm proposes to coordinate with the Air Force Research Laboratory (AFRL) Space Plug-and-play Avionics (SPA) working groups to
develop a breadboard Inertial Measurement Unit (IMU) interface that meets the latest SPA interface standards. In addition, Microcosm will study implementations
for other plug-and-play GN&C components. Today's spacecraft components come with a variety of electrical interfaces and non-standard data interfaces.
Changing the IMU on a spacecraft usually requires custom changes to the flight software. Plug-and-play spacecraft components offer the flexibility and simplified
integration that will be required to realize a 6-day spacecraft. With a standard interface, spacecraft components will be interchangeable and can be selected
specifically to meet mission needs. The AFRL is coordinating SPA working groups to define a plug-and-play standard. The final result of this Phase I SBIR effort
will be a breadboard plug-and-play IMU unit based on a Microelectromechanical (MEMS) device that meets those standards. In the Phase II SBIR effort,
Microcosm will apply the same underlying technology to build other plug-and-play GN&C components, such as magnetic torquer drive electronics and star
sensors. In a related Phase II SBIR, Microcosm built a breadboard GN&C system to internally defined plug-and-play standards.
Advanced Avionics
Avionics and Astrionics – Guidance Navigation and Control
Company Name
Waddan Systems
Title
TIME SPACE POSITION INFORMATION
GPS INS INSTRUMENTATION (TSPIGII)
Systems & Processes Engineering
Corporation
NANOSATELLITE SENSOR SUITE
DOD SBIR Phase II
Quad Chart
In Phase II, the TSPIGII prototype will be built to size using commercial hardware and Waddan’s inertial module. The baseline effort planned for this phase
include integration of COTS, development of TSPIGII specific system software, development of the flat pack inertial module, Kalman filter design for the
tightly coupled GSP/INS and its related software, benchmark testing, cost modeling, etc.
SPEC’s approach is to highly integrate an optimal mix of low-cost, low-power, miniature sized sensors into a package capable of high accuracy and high
bandwidth attitude/position determination. External interfaces have been configured to accommodate open standards and “plug-and-play” system
architectures. The results of this program will be a flight ready prototype NSS system capable of immediate follow on flight test.
Advanced Avionics
Avionics and Astrionics – Guidance Navigation and Control
WIPO
Patent Number
WO04105269A1
Title
System and Method For
Multiplexing And Trasmitting Dc
Power, Imu Data And Rf Data On A
Single Cable
Assignee
Honeywell
International Inc.
Abstract
A system for providing connectivity to co-located instruments includes a processor for receiving and processing GPS data and IMU data, a GPS
receiver antenna operable to supply the GPS data to the processor, and an IMU, co-located with the GPS receiver antenna, operable to provide
the IMU data to the processor. A single cable is provided between the processor and co-located equipment and, by using filtering mechanisms,
the single cable is operable to simultaneously supply DC power to the IMU and to transmit the GPS data and the IMU data to the processor.
Advanced Avionics
Avionics and Astrionics – Guidance Navigation and Control
FC
Marshall Space
Flight Center
Title
A Framework for Integration of IVHM
Technologies for Intelligent
Integration for Vehicle Management
Jet Propulsion
Laboratory
Equipping an FPGA-Based Mars
Rover With an LN-200 IMU
Goddard Space
Flight Center
Flying the ST-5 Constellation with
"Plug and Play" Autonomy
Components and the GMSEC Bus
NTRS
Abstract
As a part of the overall goal of developing Integrated Vehicle Health Management (IVHM) systems for aerospace vehicles, the NASA Faculty Fellowship Program (NFFP) at
Marshall Space Flight Center has performed a pilot study on IVHM principals which integrates researched IVHM technologies in support of Integrated Intelligent Vehicle
Management (IIVM). IVHM is the process of assessing, preserving, and restoring system functionality across flight and ground systems (NASA NGLT 2004). The framework
presented in this paper integrates advanced computational techniques with sensor and communication technologies for spacecraft that can generate responses through detection,
diagnosis, reasoning, and adapt to system faults in support of IIVM. These real-time responses allow the IIVM to modify the effected vehicle subsystem(s) prior to a catastrophic
event. Furthermore, the objective of this pilot program is to develop and integrate technologies which can provide a continuous, intelligent, and adaptive health state of a vehicle and
use this information to improve safety and reduce costs of operations. Recent investments in avionics, health management, and controls have been directed towards IIVM. As this
concept has matured, it has become clear the IIVM requires the same sensors and processing capabilities as the real-time avionics functions to support diagnosis of subsystem
problems. New sensors have been proposed, in addition, to augment the avionics sensors to support better system monitoring and diagnostics. As the designs have been
considered, a synergy has been realized where the real-time avionics can utilize sensors proposed for diagnostics and prognostics to make better real-time decisions in response to
detected failures. IIVM provides for a single system allowing modularity of functions and hardware across the vehicle. The framework that supports IIVM consists of 11 major onboard functions necessary to fully manage a space vehicle maintaining crew safety and mission objectives Guidance and Navigation Communications and Tracking Vehicle
Monitoring Information Transport and Integration Vehicle Diagnostics Vehicle Prognostics Vehicle mission Planning Automated Repair and Replacement Vehicle Control Human
Computer Interface and Onboard Verification and Validation. Furthermore, the presented framework provides complete vehicle management which not only allows for increased
crew safety and mission success through new intelligence capabilities, but also yields a mechanism for more efficient vehicle operations. The representative IVHM technologies for
IIVH includes 1) robust controllers for use in re-usable launch vehicles, 2) scaleable flexible computer platform using heterogeneous communication, 3) coupled electromagnetic
oscillators for enhanced communications, 4) Linux-based real-time systems, 5) genetic algorithms, 6) Bayesian Networks, 7) evolutionary algorithms, 8) dynamic systems control
modeling, and 9) advanced sensing capabilities. This paper presents IVHM technologies developed under NASA's NFFP pilot project. The integration of these IVHM technologies
forms the framework for IIVM.
The Mars Exploration Rovers (MER) currently navigating the surface of Mars are outfitted with an advanced stereovision correlation algorithm which allows them to "see" threedimensionally and autonomously avoid obstacles in their path. A bottleneck of this system is that it is computationally intense and requires 3 minutes of processing for every
correlated image and path choice. Taking advantage of the optimization and reprogrammability of FPGAs, the Mobility Avionics lab has reduced this process to under a second. The
lab is demonstrating the advancement with a prototype rover, complete with an LN-200 inertial measurement unit (IMU), which is a flight spare from MER. The LN-200 is a spacegrade, six degrees-of-freedom IMU using three fiber-optic gyroscopes and three silicon accelerometers and no moving parts. It has particular power-sequencing needs and
communicates with a specialized serial protocol (SDLC over RS-422), requiring specific hardware and software for proper functionality and interfacing with an FPGA. The process of
incorporating the LN-200 into the system is described herein.
The Space Technology 5 (ST5) Project, part of NASA's New Millennium Program, will consist of a constellation of three micro-satellites. This viewgraph document presents the
components that will allow it to operate in an autonomous mode. The ST-5 constellation will use the GSFC Mission Services Evolution Center (GMSEC) architecture to enable cost
effective model based operations. The ST-5 mission will demonstrate several principles of self managing software components.
Advanced Avionics
Avionics and Astrionics – Guidance Navigation and Control
FC
Marshall Space
Flight Center
Title
Intelligent Vehicle Health
Management
Marshall Space
Flight Center
IVHM Framework for Intelligent
Integration for Vehicle Health
Management
Goddard Space
Flight Center
Low Power Transceiver
NTRS
Abstract
As a part of the overall goal of developing Integrated Vehicle Health Management systems for aerospace vehicles, the NASA Faculty Fellowship Program (NFFP) at Marshall Space
Flight Center has performed a pilot study on IVHM principals which integrates researched IVHM technologies in support of Integrated Intelligent Vehicle Management (IIVM). IVHM
is the process of assessing, preserving, and restoring system functionality across flight and ground systems (NASA NGLT 2004). The framework presented in this paper integrates
advanced computational techniques with sensor and communication technologies for spacecraft that can generate responses through detection, diagnosis, reasoning, and adapt to
system faults in support of INM. These real-time responses allow the IIVM to modify the affected vehicle subsystem(s) prior to a catastrophic event. Furthermore, the objective of
this pilot program is to develop and integrate technologies which can provide a continuous, intelligent, and adaptive health state of a vehicle and use this information to improve
safety and reduce costs of operations. Recent investments in avionics, health management, and controls have been directed towards IIVM. As this concept has matured, it has
become clear the INM requires the same sensors and processing capabilities as the real-time avionics functions to support diagnosis of subsystem problems. New sensors have
been proposed, in addition, to augment the avionics sensors to support better system monitoring and diagnostics. As the designs have been considered, a synergy has been
realized where the real-time avionics can utilize sensors proposed for diagnostics and prognostics to make better real-time decisions in response to detected failures. IIVM provides
for a single system allowing modularity of functions and hardware across the vehicle. The framework that supports IIVM consists of 11 major on-board functions necessary to fully
manage a space vehicle maintaining crew safety and mission objectives Guidance and Navigation Communications and Tracking Vehicle Monitoring Information Transport and
Integration Vehicle Diagnostics Vehicle Prognostics Vehicle mission Planning Automated Repair and Replacement Vehicle Control Human Computer Interface and Onboard
Verification and Validation. Furthermore, the presented framework provides complete vehicle management which not only allows for increased crew safety and mission success
through new intelligence capabilities, but also yields a mechanism for more efficient vehicle operations. The representative IVHM technologies for computer platform using
heterogeneous communication, 3) coupled electromagnetic oscillators for enhanced communications, 4) Linux-based real-time systems, 5) genetic algorithms, 6) Bayesian
Networks, 7) evolutionary algorithms, 8) dynamic systems control modeling, and 9) advanced sensing capabilities. This paper presents IVHM technologies developed under NASA's
NFFP pilot project and the integration of these technologies forms the framework for IIVM.
Integrated Vehicle Health Management (IVHM) systems for aerospace vehicles, is the process of assessing, preserving, and restoring system functionality across flight and
techniques with sensor and communication technologies for spacecraft that can generate responses through detection, diagnosis, reasoning, and adapt to system faults in support
of Integrated Intelligent Vehicle Management (IIVM). These real-time responses allow the IIVM to modify the affected vehicle subsystem(s) prior to a catastrophic event.
Furthermore, this framework integrates technologies which can provide a continuous, intelligent, and adaptive health state of a vehicle and use this information to improve safety
and reduce costs of operations. Recent investments in avionics, health management, and controls have been directed towards IIVM. As this concept has matured, it has become
clear that IIVM requires the same sensors and processing capabilities as the real-time avionics functions to support diagnosis of subsystem problems. New sensors have been
proposed, in addition to augment the avionics sensors to support better system monitoring and diagnostics. As the designs have been considered, a synergy has been realized
where the real-time avionics can utilize sensors proposed for diagnostics and prognostics to make better real-time decisions in response to detected failures. IIVM provides for a
single system allowing modularity of functions and hardware across the vehicle. The framework that supports IIVM consists of 11 major on-board functions necessary to fully
manage a space vehicle maintaining crew safety and mission objectives. These systems include the following Guidance and Navigation Communications and Tracking Vehicle
Monitoring Information Transport and Integration Vehicle Diagnostics Vehicle Prognostics Vehicle Mission Planning, Automated Repair and Replacement Vehicle Control Human
Computer Interface and Onboard Verification and Validation. Furthermore, the presented framework provides complete vehicle management which not only allows for increased
crew safety and mission success through new intelligence capabilities, but also yields a mechanism for more efficient vehicle operations.
One of the FREESTAR experiments, the Low Power Transceiver (LPT) experiment is a low-power, lightweight software programmable transceiver prototype technology
demonstration that is being developed by NASA as a low-cost S-band spacecraft navigation and communication device. The LPT prototype receives Global Positioning System
(GPS) satellite signals for spacecraft navigation support and provides both forward and return, low-rate data communications links to the Merritt Island (MILA) and Dryden Flight
Research Facility (DFRC) ground stations and to the Tracking and Data Relay Satellite System (TDRSS). The experiment is designed to demonstrate the system's ability to do
simultaneous communications and navigation, as well as multi-mode communications and reconfiguration. LPT is managed by NASA's Goddard Space Flight Center and sponsored
by NASA HQ Code M. The LPT experiment consists of one thermally conductive box containing the electronics stack, three S-band antennas and one L-band antenna. The LPT
payload uses general Orbiter services, including power control, command, and telemetry provided through the HHC avionics. On-orbit, the LPT payload will be primarily operated via
direct communications between LPT and ground stations (MILA, WLPS, or DFRC) and or TDRSS, with backup command and telemetry capability provided via the hitchhiker
avionics and remote Payload Operations Control Center. During operations, LPT will utilize high S-band frequencies for communications. The LPT TDRSS (and GN) forward link
(uplink) frequency is 2106.40625 MHz and their TDRSS (and GN) return link (downlink) frequency is 2287.5 MHz (utilizing Left-handed Circular Polarization to work with the TDRSS
MA system). Two standard switch panel switches will be utilized to prohibit inadvertent operation of the antenna. An additional inhibit will be provided through the HH avionics power
relay to the LPT.
Advanced Avionics
Avionics and Astrionics – Guidance Navigation and Control
NTRS
Abstract
The objective of this research is to design an intelligent plug-n-play avionics system that provides a reconfigurable platform for supporting the guidance, navigation and control (GN
and ampC) requirements for different elements of the space exploration mission. The focus of this study is to look at the specific requirements for a spacecraft that needs to go from
earth to moon and back. In this regard we will identify the different GN and ampC problems in various phases of flight that need to be addressed for designing such a plug-n-play
avionics system. The Apollo and the Space Shuttle programs provide rich literature in terms of understanding some of the general GN and ampC requirements for a space vehicle.
The relevant literature is reviewed which helps in narrowing down the different GN and ampC algorithms that need to be supported along with their individual requirements.
FC
Ames Research
Center
Title
Spacecraft guidance, navigation, and
control requirements for an intelligent
plug-n-play avionics (PAPA)
architecture
Johnson Space
Center
Synchronized Position Hold, Engage,
Reorient, Experimental Satellites
(SPHERES)
SPHERES is a testbed for formation flying by satellites, the theories and calculations that coordinate the motion of multiple bodies maneuvering in microgravity. To achieve this
inside the ISS cabin, bowling-ball-sized spheres perform various maneuvers (or protocols), with one to three spheres operating simultaneously . The Synchronized Position Hold,
Engage, Reorient, Experimental Satellites (SPHERES) experiment will test relative attitude control and station-keeping between satellites, re-targeting and image plane filling
maneuvers, collision avoidance and fuel balancing algorithms, and an array of geometry estimators used in various missions. SPHERES consists of three self-contained satellites,
which are 18-sided polyhedrons that are 0.2 meter in diameter and weigh 3.5 kilograms. Each satellite contains an internal propulsion system, power, avionics, software,
communications, and metrology subsystems. The propulsion system uses CO2, which is expelled through the thrusters. SPHERES satellites are powered by AA batteries. The
metrology subsystem provides real-time position and attitude information. To simulate ground station-keeping, a laptop will be used to transmit navigational data and formation flying
algorithms. Once these data are uploaded, the satellites will perform autonomously and hold the formation until a new command is given.
Marshall Space
Flight Center
The Integrated Safety-Critical
Advanced Avionics Communication
and Control (ISAACC) System
Concept Infrastructure for ISHM
Johnson Space
Center
The Miniature Autonomous
Extravehicular Robotic Camera
(MiniAERCam) for Spacecraft
Inspection and Remote
Tightly Coupled Inertial Navigation
System Global Positioning System
(TCMIG)
Integrated System Health Management (ISHM) architectures for spacecraft will include hard real-time, critical subsystems and soft real-time monitoring subsystems. Interaction
between these subsystems will be necessary and an architecture supporting multiple criticality levels will be required. Demonstration hardware for the Integrated Safety-Critical
Advanced Avionics Communication and amp Control (ISAACC) system has been developed at NASA Marshall Space Flight Center. It is a modular system using a commercially
available time-triggered protocol, Tp C, that supports hard real-time distributed control systems independent of the data transmission medium. The protocol is implemented in
hardware and provides guaranteed low-latency messaging with inherent fault-tolerance and fault-containment. Interoperability between modules and systems of modules using the
TTPC is guaranteed through definition of messages and the precise message schedule implemented by the master-less Time Division Multiple Access (TDMA) communications
protocol. "Plug-and-play" capability for sensors and actuators provides automatically configurable modules supporting sensor recalibration and control algorithm re-tuning without
software modification. Modular components of controlled physical system(s) critical to control algorithm tuning, such as pumps or valve components in an engine, can be replaced or
upgraded as "plug and play" components without modification to the ISAACC module hardware or software. ISAACC modules can communicate with other vehicle subsystems
through time-triggered protocols or other communications protocols implemented over Ethernet, MIL-STD- 1553 and RS-485 422. Other communication bus physical layers and
protocols can be included as required. In this way, the ISAACC modules can be part of a system-of-systems in a vehicle with multi-tier subsystems of varying criticality. The goal of
the ISAACC architecture development is control and monitoring of safety critical systems of a manned spacecraft. These systems include spacecraft navigation and attitude control,
propulsion, automated docking, vehicle health management and life support. ISAACC can integrate local critical subsystem health management with subsystems performing long
term health monitoring. The ISAACC system and its relationship to ISHM will be presented.
AERCam is a nano-satellite class free-flying spacecraft with a full suite of avionics, propulsion, navigation, and communications. 1) 3 major development programs, one ending in
DTO of protoflight unit, other two ending in ground demonstrations with integrated hardware and software. 2) Incremental increase in capability to reduce crew workload, provide
better inspection capability. 3) Two crew evaluations and 4) Significant technology advancement.
Marshall Space
Flight Center
Many NASA applications planned for execution later this decade are seeking high performance, miniaturized, low power Inertial Management Units (IMU). Much research has gone
into Micro-Electro-Mechanical System (MEMS) over the past decade as a solution to these needs. While MEMS devices have proven to provide high accuracy acceleration
measurements, they have not yet proven to have the accuracy required by many NASA missions in rotational measurements. Therefore, a new solution has been formulated
integrating the best of all IMU technologies to address these mid-term needs in the form of a Tightly Coupled Micro Inertial Navigation System (INS) Global Positioning System
(GPS) (TCMIG). The TCMIG consists of an INS and a GPS tightly coupled by a Kalman filter executing on an embedded Field Programmable Gate Array (FPGA) processor. The
INS consists of a highly integrated Interferometric Fiber Optic Gyroscope (IFOG) and a MEMS accelerometer. The IFOG utilizes a tightly wound fiber coil to reduce volume and the
high level of integration and advanced optical components to reduce power. The MEMS accelerometer utilizes a newly developed deep etch process to increase the proof mass and
yield a highly accurate accelerometer. The GPS receiver consists of a low power miniaturized version of the Blackjack receiver. Such an IMU configuration is ideal to meet the midterm needs of the NASA Science Enterprises and the new launch vehicles being developed for the Space Launch Initiative (SLI).
Advanced Avionics
Avionics and Astrionics – Guidance Navigation and Control
FC
Goddard Space
Flight Center;
Marshall Space
Flight Cente
Title
Time Triggered Protocol (TTP) for
Integration Modular Avionics (IMA)
Jet Propulsion
Laboratory
Ultra-long-life spacecraft for long
duration space exploration missions
NTRS
Abstract
This viewgraph presentation is a review of the Time Triggered Protocol, designed to work with NASA's Integrated Safety-Critical Advanced Avionics Communication and Control
(ISAACC) system. ISAACC is the product of the Propulsion High-Impact Avionics Technologies (PHIAT) project at NASA Marshall Space Flight Center (MSFC) during FY03 to the
end of FY05. The goal is an avionics architecture suitable for control and monitoring of safety critical systems of manned spacecraft. It must be scalable to allow its use in robotic
vehicles or launch pad and propulsion test stand monitoring and control systems. The developed IMA should have a common power supply and rugged chassis for a set of modules,
many upgradeable software functions on one module (i.e. processing unit Reduced weight, straightforward update and system integration. It is also important that it have Partitioning
and a Memory Management Unit (MMU)
The predominant failure mode in an ultra longlife system is the wear-out of components. In order to survive long duration missions, current fault tolerant design techniques would
require excessive number of redundant components. This paper describes a more efficient fault tolerant avionics system architecture that requires much less redundant
components. This architecture employs generic function blocks that can be programmed to replace a wide variety of components in-flight. Hence, each individual generic block is
essentially equivalent to almost an entire redundant string of components in the conventional approach. In that way, the ultra long-life system can achieve much higher level of
reliability while carrying much less components. On the other hand, due to the programmability of the generic redundant blocks, the physical location of a specific component might
not be pre-determined. Therefore, wireless interconnection is employed to provide the necessary flexibility in connectivity. A testbed of this architecture is being developed at the Jet
Propulsion Laboratory.
Advanced Avionics
Avionics and Astrionics – On-Board Computing and Data Management
Company Name
MicroSat Systems, Inc.
Title
RECONFIGURABLE,
COMMERCIAL-BASED AVIONICS
FOR RESPONSIVE
MICROSATELLITE TARGETS
DOD SBIR Phase II
Quad Chart
MSI is developing a low cost, modular solution to mission unique, highly specialized, expensive avionics. The multi-node, ring topology enables swapping of the
processing functions between any node facilitating system redundancy. Each node function of ARMT can be specified independently through real-time,
software/firmware uploads, tailoring spacecraft system capability and price through the addition or removal of nodes.
Advanced Avionics
Avionics and Astrionics – On-Board Computing and Data Management
NTRS
Abstract
Future small and micro-missions, such as Mars Scouts and Deep Space probes, require a new look at highly integrated, re-configurable, low power avionics. This paper will
present our plans for developing a scalable, configurable, and highly integrated 32-bit embedded platform capable of implementing computationally intensive signal processing
and control algorithms in space flight instruments and systems. This platform is designed to service the need of both small and large spacecraft and planetary rovers that will
operate within moderate radiation environments. Some of the key characteristics of this platform are its small size, low power, high performance, and flexibility. This estimated
10 fold reduction in both size and power over state-of-the-art processing platforms will enable this new product to act as the core of a low-cost mobility system for a wide range
of missions.
FC
Jet Propulsion
Laboratory
Title
Advanced mobility avionics a
reconfigurable mirco-avionics platform for
the future needs of small planetary rovers
and micospacecraft
Marshall Space
Flight Center
Comparison of Communication
Architectures for Spacecraft Modular
Avionics Systems
This document is a survey of publicly available information concerning serial communication architectures used, or proposed to be used, in aeronautic and aerospace
applications. It focuses on serial communication architectures that are suitable for low-latency or real-time communication between physically distributed nodes in a system.
Candidates for the study have either extensive deployment in the field, or appear to be viable for near-term deployment. Eleven different serial communication architectures
are considered, and a brief description of each is given with the salient features summarized in a table in appendix A. This survey is a product of the Propulsion High Impact
Avionics Technology (PHIAT) Project at NASA Marshall Space Flight Center (MSFC). PHIAT was originally funded under the Next Generation Launch Technology (NGLT)
Program to develop avionics technologies for control of next generation reusable rocket engines. After the announcement of the Space Exploration Initiative, the scope of the
project was expanded to include vehicle systems control for human and robotics missions. As such, a section is included presenting the rationale used for selection of a timetriggered architecture for implementation of the avionics demonstration hardware developed by the project team
Marshall Space
Flight Center
Crew Launch Vehicle (CLV) Avionics and
Software Integration Overview
Jet Propulsion
Laboratory
Development of reliable electronic
packaging solutions for spacecraft
avionics miniaturization using embedded
passice devices
On January 14, 2004, the President of the United States announced a new plan to explore space and extend a human presence across our solar system. The National
Aeronautics and Space Administration (NASA) established the Exploration Systems Mission Directorate (ESMD) to develop and field a Constellation Architecture that will
bring the Space Exploration vision to fruition. The Constellation Architecture includes a human-rated Crew Launch Vehicle (CLV) segment, managed by the Marshall Space
Flight Center (MSFC), comprised of the First Stage (FS), Upper Stage (US), and Upper Stage Engine (USE) elements. The CLV s purpose is to provide safe and reliable crew
and cargo transportation into Low Earth Orbit (LEO), as well as insertion into trans-lunar trajectories. The architecture's Spacecraft segment includes, among other elements,
the Crew Exploration Vehicle (CEV), managed by the Johnson Space Flight Center (JSC), which is launched atop the CLV. MSFC is also responsible for CLV and CEV stack
integration. This paper provides an overview of the Avionics and Software integration approach (which includes the Integrated System Health Management (ISHM) functions),
both within the CLV, and across the CEV interface it addresses the requirements to be met, logistics of meeting those requirements, and the roles of the various groups. The
Avionics Integration and Vehicle Systems Test (ANST) Office was established at the MSFC with system engineering responsibilities for defining and developing the integrated
CLV Avionics and Software system. The AIVST Office has defined two Groups, the Avionics and Software Integration Group (AVSIG), and the Integrated System Simulation
and Test Integration Group (ISSTIG), and four Panels which will direct trade studies and analyses to ensure the CLV avionics and software meet CLV system and CEV
interface requirements. The four panels are 1) Avionics Integration Panel (AIP), 2) Software Integration Panel, 3) EEE Panel, and 4) Systems Simulation and Test Panel.
Membership on the groups and panels includes the MSFC representatives from the requisite engineering disciplines, the First Stage, the Upper Stage, the Upper Stage
Engine projects, and key personnel from other NASA centers. The four panels will take the results of trade studies and analyses and develop documentation in support of
Design Analysis Cycle Reviews and ultimately the System Requirements Review.
Miniaturization of electronic packages will play a key rule in future space avionics systems. Smaller avionics packages will reduce payloads while providing greater
functionality for information processing and mission instrumentation. Current surface mount technology discrete passive devices not only take up significant space but also
add weight. To that end, the use of embedded passive devices, such as capacitors, inductors and resistors will be instrumental in allowing electronics to be made smaller and
lighter. Embedded passive devices fabricated on silicon or like substrates using thin film technology, promise great savings in circuit volume, as well as potentially improving
electrical performance by decreasing parasitic losses. These devices exhibit a low physical profile and allow the circuit footprint to be reduced by stacking passive elements
within a substrate. Thin film technologies used to deposit embedded passive devices are improving and costs associated with the process are decreasing.
Advanced Avionics
Avionics and Astrionics – On-Board Computing and Data Management
NTRS
Abstract
Digital Avionics activities played an important role in the advancements made in civil aviation, military systems, and space applications. This document profiles advances
made in each of these areas by the aerospace industry, NASA centers, and the U.S. military. Emerging communication technologies covered in this document include Internet
connectivity onboard aircraft, wireless broadband communication for aircraft, and a mobile router for aircraft to communicate in multiple communication networks over the
course of a flight. Military technologies covered in this document include avionics for unmanned combat air vehicles and microsatellites, and head-up displays. Other
technologies covered in this document include an electronic flight bag for the Boeing 777, and surveillance systems for managing airport operations.
FC
Glenn Research
Center;
Marshall Space
Flight Center
Title
Digital Avionics
Jet Propulsion
Laboratory
Integrated Avionics System (IAS),
integrating 3-D technology on a
spacecraft panel
A novel system packaging approach incorporates solutions that provide broader environmental applications, more flexible system interconnectivity, scalability, and simplified
assembly test and integration schemes.
Jet Propulsion
Laboratory
Small computational node embedded
within a high-speed network fabric for
spacecraft flight applications
Goddard Space
Flight Center
The Space Technology 5 Avionics
System
"JPL is developing a highly integrated Micro Avionics Module aimed at the avionics needs of both small and large spacecraft and planetary rovers that will operate within a
moderate radiation environment. The Micro Avionics Module will be based upon the commercially available Xilinx Virtex-ll Pro field programmable gate array. The resulting
product will be a low power, highly scalable, microprocessor-based avionics package that can be used alone or within a network. For example, such an avionics module could
become the core C&DH computer of a Mars Scout-class spacecraft, or a stereo vision processor for a larger planetary rover. The key characteristics of this Micro Avionics
Module are:
• Small size. The module will be less than ten cubic inches in volume. Low power. Depending on how the module is configured, milli-watts to a few watts of power
consumption are anticipated.
• High computational power. One, two, or four PowerPC 405 processors running at up to 400 MHz are embedded within the Virtex-ll Pros gate array. The processors can be
voted against each other for increased fault tolerance or run independently for additional processing capability.
• Support for off-the-shelf operating systems like VxWorks and Linux.
• High 110 count. Hundreds of I/O pins are available. Each is programmable for voltage swing, impedance, and single-ended or differential usage.
• Large number of re-configurable FPGA gates. Up to eight million gates that can be used to implement dedicated functions like brushless motor controller(s), IEEE-1394
interface(s), software defined digital radio, etc.
• Architectural provisions such as a temperature adaptive power supply or self-controlled local heating so that the components can operate in ambient temperatures that are
outside their normal operating temperature.
• Networking capability. Implementing a network controller(s) within the FPGA portion of the Virtex-ll Pro is very straightforward and allow multiple modules to be operated in a
distributed manner across slow (e.g., 12C) and fast, low latency (e.g., gigabit ethernet) networks.
We will discuss the architecture and implementation details including the design methods employed to harden the Virtex-ll Pro for moderate radiation environments."
The Space Technology 5 (ST5) mission is a NASA New Millennium Program project that will validate new technologies for future space science missions and demonstrate the
feasibility of building launching and operating multiple, miniature spacecraft that can collect research-quality in-situ science measurements. The three satellites in the ST5
constellation will be launched into a sun-synchronous Earth orbit in early 2006. ST5 fits into the 25-kilogram and 24-watt class of very small but fully capable spacecraft. The
new technologies and design concepts for a compact power and command and data handling (C and ampDH) avionics system are presented. The 2-card ST5 avionics design
incorporates new technology components while being tightly constrained in mass, power and volume. In order to hold down the mass and volume, and quali and amp new
technologies for fUture use in space, high efficiency triple-junction solar cells and a lithium-ion battery were baselined into the power system design. The flight computer is colocated with the power system electronics in an integral spacecraft structural enclosure called the card cage assembly. The flight computer has a full set of uplink, downlink
and solid-state recording capabilities, and it implements a new CMOS Ultra-Low Power Radiation Tolerant logic technology. There were a number of challenges imposed by
the ST5 mission. Specifically, designing a micro-sat class spacecraft demanded that minimizing mass, volume and power dissipation would drive the overall design. The result
is a very streamlined approach, while striving to maintain a high level of capability, The mission's radiation requirements, along with the low voltage DC power distribution,
limited the selection of analog parts that can operate within these constraints. The challenge of qualifying new technology components for the space environment within a
short development schedule was another hurdle. The mission requirements also demanded magnetic cleanliness in order to reduce the effect of stray (spacecraft-generated)
magnetic fields on the science-grade magnetometer.
Advanced Avionics
Avionics and Astrionics – Pilot Support Systems
FC
Jet Propulsion Laboratory
Title
Micro-Inspector Avionics Module (MAM) A SelfContained Low Power, Reconfigurable Avionics
Platform for Small Spacecrafts and Instruments
Johnson Space Center
Mini AERCam Inspection Robot for Human Space
Missions
Jet Propulsion Laboratory
Mission operation for reconfigurable spacecraft
NTRS
Abstract
This paper describes development of a radiation tolerant, low power, reconfigurable avionics module aimed at meeting the avionics needs of the JPL
Micro-Inspector spacecraft. This module represents a complete avionics system, consisting of two PowerPC 405 CPUs embedded within a reconfigurable
FPGA fabric of over 8 Million logic gates, 64MB of EDAC protected Flash storage and 128MB of EDAC protected DDR SDRAM or SDRAM memories,
along with FPGA SEU mitigation logic, and all necessary power conversion. Processor SEU mitigation is achieved by running the two processors in a
lock-step and compare configuration. All of these building blocks are integrated into a double sided circuit board that takes as little as 6 square inches of
board space. This module can be embedded into a user system as part of a bigger circuit assembly or as a self contained module. This module is being
developed as part of a JPL led Micro-Inspector Program, funded by NASA ESMD aimed at producing a 10Kg micro spacecraft.
The Engineering Directorate of NASA Johnson Space Center has developed a nanosatellite-class free-flyer intended for future external inspection and
remote viewing of human spacecraft. The Miniature Autonomous Extravehicular Robotic Camera (Mini AERCam) technology demonstration unit has been
integrated into the approximate form and function of a flight system. The spherical Mini AERCam free flyer is 7.5 inches in diameter and weighs
approximately 10 pounds, yet it incorporates significant additional capabilities compared to the 35 pound, 14 inch AERCam Sprint that flew as a Shuttle
flight experiment in 1997. Mini AERCam hosts a full suite of miniaturized avionics, instrumentation, communications, navigation, imaging, power, and
propulsion subsystems, including digital video cameras and a high resolution still image camera. The vehicle is designed for either remotely piloted
operations or supervised autonomous operations including automatic stationkeeping and point-to-point maneuvering. Mini AERCam is designed to fulfill
the unique requirements and constraints associated with using a free flyer to perform external inspections and remote viewing of human spacecraft
operations. This paper describes the application of Mini AERCam for stand-alone spacecraft inspection, as well as for roles on teams of humans and
robots conducting future space exploration missions.
This paper will examine the issues and solutions to these challenges. First, a brief survey of the current FPGA technology and the process to reconfigure
these devices will be provided. Second, the avionics architectural support to facilitate the inflight reconfiguration will be discussed. Third, the mission
operation procedures and uplink downlink process to reconfigure the spacecraft in-orbit will be described.
Advanced Avionics
Avionics and Astrionics – Telemetry, Tracking, and Control
Company Name
Microcosm, Inc.
Title
MINIATURE STAR SENSOR USING “CAMERA-ONA-CHIP” CMOS ARRAYS
NASA SBIR Phase II
Field Center
Quad Chart
GSFC
In Phase II, the Microcosm team will develop a brassboard miniature star sensor, which will provide nanosatellites with high
bandwidth, moderate-to-high accuracy, low cost attitude determination. Phase II will complete the detailed sensor design,
addressing both a very low cost, moderate accuracy version, and a relatively low cost, high accuracy, and/or high reliability
version, to satisfy different customer requirements.
Advanced Avionics
Avionics and Astrionics – Telemetry, Tracking, and Control
Company Name
Applied Technology Associates
Title
AUTO-CORRECTING INERTIAL MEASUREMENT UNIT
DOD SBIR Phase II
Quad Chart
The ACIMU includes autonomous auto-correcting capabilities to overcome the inherent drift associated with all inertial
measurement units (IMUs). The joint inclusion of a global positioning system (GPS) receiver along with a state-of-the-art
autonomous star tracker makes available the necessary information required for the auto-correcting capability.
Advanced Avionics
Communications – Architectures and Networks
Company Name
CFD Research Corp
Title
Computer Aided Design Tools for Extreme
Environment Electronics
Field Center
JPL
NASA SBIR Phase I
Quad Chart
This project aims to provide Computer Aided Design (CAD) tools for radiation-tolerant, wide-temperature-range digital, analog, mixed-signal,
and radio-frequency electronic components suitable for operation in the extreme environments of the Moon, Mars, and other deep space
destinations. All such exploration systems will need reliable electronics able to operate in a wide temperature range (-230?C to +130 ?C) and
high radiation levels. There is very little knowledge of semiconductor device behavior in extreme low temperatures (currently ongoing
research) and no reliable models or design tools. CFDRC will develop first commercial-quality validated models and CAD tools for predicting
the electrical performance and reliability of electronic components in extreme low temperatures, with included radiation effects and reliability
analysis, using coupled semiconductor and thermal-mechanical simulation. This work will use and implement the newest data from the
ongoing NASA Exploration Systems and Research Technology (Code ES&RT) program, led by Prof. Cressler at Georgia Tech
(subcontractor in this proposal), involving JPL, BAE, Boeing, Vanderbilt, and others, aimed at developing electronics technology for mixedsignal circuit applications for lunar (to -230?C) applications. Reliable and validated CAD tools will help to predict electronics performance and
radiation response in the extreme temperatures, and reduce the amount of testing cost and time.
Advanced Avionics
Communications – Architectures and Networks
Company Name
MicroSat Systems, Inc.
Title
MODULAR, PLUG AND PLAY,
DISTRIBUTED AVIONICS
Field Center
MSFC
NASA SBIR Phase II
Quad Chart
The company proved the viability of an Ethernet version of its modular, plug and play (PnP) spacecraft avionics architecture,
which provides a near-term solution to modular, PnP avionics while distributing power and data management functions on a
single circuit. This allows for rapid interfacing to other satellite avionics, such as the GSFC Space Cube. Lower hardware
costs are only 40-60% of comparable centralized systems.
Advanced Avionics
Communications – Architectures and Networks
Company Name
Aeroastro, Inc.
Title
Simplified Spacecraft Multi-Tap Avionics Bus
(M-TAB)
Microsat Systems
Low Cost, Encrypted Communication Link for
MTS
DOD SBIR Phase I
Quad Chart
AeroAstro proposes to develop an innovative spacecraft communications and power bus that will solve the complexity and reliability issues identified on
the Space Tracking and Surveillance System (STSS) spacecraft. The multi-tap data and power bus (M-TAB) uses a single cable to connect all nodes of
the network in a chain which provides for dramatic improvements in: reliability, cost, mass, and integration and testing complexity. M-TAB will be
designed to provide redundant high-power and high-bandwidth data connections to all of the spacecraft's subsystems as well as providing an accurate
whole spacecraft timing source. AeroAstro intends a spiral development strategy for M-TAB, a near-term solution that will develop a bus architecture that
can be implemented quickly with little or no modification to legacy avionic systems, and a far-term solution that will be compatible with the avionics
standardization activities such as SPA-U. As part of the Phase I feasibility demonstration, AeroAstro intends to fabricate a laboratory testbed of the MTAB architecture. This testbed will be the launching point for a more flight-like prototype to be developed in Phase II, one that in form and function will
simulate the final M-TAB hardware and software.
This effort will develop an innovative, space-to-ground, encrypted comm system for a microsatellite target system used to calibrate ground and sea
based radar assets. For less than $200K, this system will provide the user the versatility to locate and operate a dedicated ground station anywhere on
the globe. It will full uplink and downlink encryption and transmit via a unified S-Band frequency.
Advanced Avionics
Communications – Architectures and Networks
FC
Goddard Space Flight Center
Title
A Proven Ground System Architecture for
Promoting Collaboration and Common Solutions at
NASA
NTRS
Abstract
NASA Goddard Space Flight Center's "GMSEC" ground system architecture was presented at GSAW2003 as a concept being studied. GMSEC would
utilize a publish subscribe middleware framework and standardized interfaces to allow custom and COTS ground system components to plug-and-play.
This capability, in turn, would reduce integration costs, allow for technology infusion over time, and encourage the development and sharing of common
components across missions and organizations. At GSAW2004, GMSEC was presented at a breakout session as a system working well in the NASA lab
and being applied as an integral piece of reengineering efforts for several GSFC missions. Today, GMSEC is supporting five satellites at GSFC and has
been selected by several future missions. Over 30 plug-and-play components are now available to missions using the GMSEC approach. Other
organizations, including Marshall Space Flight Center, Johns Hopkins University's Applied Physic Lab, and the Institute for Scientific Research are each
developing GMSEC-compatible components. Based on the success of GMSEC and efforts at other NASA Centers, the message bus approach is now
being evaluated as a NASA Agency-wide approach for many future missions involving multiple NASA Centers as we move towards the goals of NASA s
new Exploration Initiative. The presentation will explain the basic technical concepts of using a publish subscribe framework for mission operations
support (and its applicability to flight systems as well). Lessons learned on NASA's GMSEC program will allow the audience to better understand the
significant benefits of this architecture approach over the traditional "one-off" solution approach. The point of the presentation is to show the long-term
benefits of using a ground system architecture which incorporates many of the successful GMSEC concepts - message bus, mix of COTS and custom
software, standard interfaces, plug-and-play, etc. The implications for the development process will also be discussed.
Advanced Avionics
Communications – Autonomous Control and Monitoring
Company Name
Invocon, Inc.
Title
High Impact G-Loading Data Acquisition System
(HI-G-DAS)
DOD SBIR Phase I
Quad Chart
A system is proposed for high-performance wireless data acquisition in high-impact environments, specifically the launch environment of Naval
payloads. The High Impact G-Loading Data Acquisition System (HI-G-DAS) is a miniature, low power instrumentation system designed to
maximize flexibility by combining precision data gathering capability, user programmability, and long battery life. HI-G-DAS will use state of the art
techniques for component selection, PCB design, and modeling of critical elements to withstand acceleration on the order of 1000 g's. The robust
system will be capable of interfacing with multiple types of sensors through the use of proven modular techniques and channel-specific
programmability. Using wireless communication and battery charging techniques, the system can be fully protected from water ingress in harsh
sea conditions. Additionally, its automated underwater beacon will help to locate the device in the event that recovery is necessary. The use of
standard software & hardware interfaces as well as common wireless protocols will help to ensure that operation of the system is simple and
intuitive.
Advanced Avionics
Electronics – Highly-Reconfigurable
Company Name
Coherent Technical
Services, Inc
Title
Avionics for Scaled Remotely Operated
Vehicles
Field Center
LARC
NASA SBIR Phase I
Quad Chart
The use of UAVs has increased exponentially since 1995, and this growth is expected to continue. Many of these applications require extensive
Research and Development; however, the need to fund development of the UAV often competes with funding intended for the end-user application.
Therefore, off the shelf, low cost, easily configurable integrated avionics systems will significantly reduce the budget impact for UAVs yet will
support the wide range of applications for their use. CTSi and Virginia Commonwealth University are proposing the use of an integrated VCU
developed avionics package with a user configurable autopilot system that will meet the needs of a wide range of experimental test bed UAVs. The
system will include: 1. The ability for the safety pilot to take direct control of the aircraft using an on-board fail-safe control switch 2. A built-in
autopilot to provide return-to-home capability upon failure of the RF links, safety/ground pilot assistance in performing research maneuvers, and
limited upset recovery 3. An open-architecture hardware design enabling customer upgrade of sensors, actuators, and data links 4. An openarchitecture software design enabling push-button auto-coding of control algorithms direct from Simulink 5. A flexible architecture allowing
customer-developed control laws to be executed on ground-based computers via uplink and downlink telemetry or onboard the aircraft using an
optional Advanced Adaptive Flight Control Processor.
Advanced Avionics
Electronics – Highly-Reconfigurable
Company Name
Advanced Science and Novel
Title
SWITCHING FABRIC BASED ON MULTI-LEVEL
LVDS COMPATIBLE INTERCONNECT
American Semiconductor, Inc.
LOW-POWER, RAD-HARD RECONFIGURABLE, BIDIRECTIONAL FLEXFET LEVEL SHIFTER ReBiLS
FOR MULTIPLE GENERATION TECHNOLOGY
INTEGRATION FOR SPACE EXPLORATION
NASA SBIR Phase II
Field Center
Quad Chart
GSFC
Switching fabric (SF) is the key component of the next generation of back plane interconnects. High power
consumption of SFs limit their application in next generation nanosatellites. This technology will minimize
latency and reduce power consumption. It is radiation tolerant and easy to align SF based on a multi level
power efficient low voltage differential signal interface.
GSFC
Extended utilization of current and development of new space exploration systems require multiple voltage
levels be successfully integrated to interface technologies. The ASI ReBiLS will generate prototypes for
stand-alone and embedded bidirectional level shifters to provide efficient communication between existing
components at different voltages.
Advanced Avionics
Electronics – Highly-Reconfigurable
FC
Jet Propulsion
Laboratory
Title
Onboard plug and play standardization
effort
Goddard Space Flight
Center
SpaceWire Plug and Play
Goddard Space Flight
Center
Standards-and Component-Based
Mission Operations Architecture at
NASA's Goddard Space Flight Center
NTRS
Abstract
"This work introduces ""Onboard Plug and Play Standardization"". This work is a study of the role of plug and play services forspacecraft.
Idea: There exist use cases where plug and play services would give system engineers greater leverage in spacecraft development and automation.
A standard Plug and Play architecture is envisaged as mission complexity grows (Mars surface ops/Constellation missions etc)."
The ability to rapidly deploy inexpensive satellites to meet tactical goals has become an important goal for military space systems. In fact, Operationally Responsive Space
(ORS) has been in the spotlight at the highest levels. The Office of the Secretary of Defense (OSD) has identified that the critical next step is developing the bus standards
and modular interfaces. Historically, satellite components have been constructed based on bus standards and standardized interfaces. However, this has not been done to a
degree, which would allow the rapid deployment of a satellite. Advancements in plug-and-play (PnP) technologies for terrestrial applications can serve as a baseline model
for a PnP approach for satellite applications. Since SpaceWire (SpW) has become a de facto standard for satellite high-speed (greater than 200Mbp) on-board
communications, it has become important for SpW to adapt to this Plug and Play (PnP) environment. Because SpW is simply a bulk transport protocol and lacks built-in PnP
features, several changes are required to facilitate PnP with SpW. The first is for Host(s) to figure out what the network looks like, i.e., how pieces of the network, routers
and nodes, are connected together network mapping, and to receive notice of changes to the network. The second is for the components connected to the network to be
understood so that they can communicate. The first element, network topology mapping and amp change of status indication, is being defined (topic of this paper). The
second element describing how components are to communicate has been defined by ARFL with the electronic data sheets known as XTEDS. The first element, network
mapping, is recent activities performed by Air Force Research Lab (ARFL), Naval Research Lab (NRL), NASA and US industry (Honeywell, Clearwater, FL, and others).
This work has resulted in the development of a protocol that will perform the lower level functions of network mapping and Change Of Status (COS) indication required by
Plug 'n' Play over SpaceWire. This work will be presented to the SpaceWire working group for standardization under European Cooperation for Space Standardization
(ECSS) and to obtain a permanent Protocol ID (see SpaceWire Protocol ID What Does it Mean to You IEEE Aerospace Conference 2006). The portion of the Plug 'n' Play
protocol that will be described in this paper is how the Host(s) of a SpaceWire network map the network and detect additions and deletions of devices on a SpaceWire
network.
NASA Goddard Space Flight Center (GSFC) manages many of NASA s earth and space science satellite missions. A wide variety of commercial products and GSFCdeveloped software components are typically integrated into a unique system configuration for each mission. Independent development of the many mission operations
center systems has led to systems that are expensive to integrate, difficult to infuse with new capabilities developed for other programs, and cumbersome to maintain. This
traditional approach becomes even more problematic as NASA moves towards satellite constellations, new operations concepts, and even further budgets reductions. The
GSFC Mission Services Evolution Center (GMSEC) is creating a new architecture for future missions at GSFC. Instead of selecting the best-in-class components and
creating a standard control center system, GMSEC is developing component interface standards so that multiple products can plug-and-play into the configuration. Missions
can then select the best components based on the merits of the product and not simply based on recent integration history at NASA. The GMSEC system utilizes a publish
subscribe information bus and standard XML-based key message interfaces. Functional components can either match directly to the interface standard, or adapters can be
developed to match the product's interface to the GMSEC standard with out impacting the source product. Applications Program Interfaces (API's) are being developed to
isolate the underlying middleware from the applications software and to allow the middleware product to be switched if necessary. Interface Control Documents (ICDs)
between each pair of communicating components is replaced by a single message API specification document. New applications must simply match to the information bus
standards and need not worry about all of the other applications in the system. For legacy software, adapters can be developed to facilitate communications between the
application and the information bus. As the approach has matured, it has become apparent that it can provide innovative solutions to some of the multi-satellite challenges
facing GSFC.
Advanced Avionics
Electronics – Highly-Reconfigurable
FC
Johnson Space
Center
Title
Use of Field Programmable Gate Array
Technology in Future Space Avionics
NTRS
Abstract
Fulfilling NASA's new vision for space exploration requires the development of sustainable, flexible and fault tolerant spacecraft control systems. The traditional development
paradigm consists of the purchase or fabrication of hardware boards with fixed processor and or Digital Signal Processing (DSP) components interconnected via a
standardized bus system. This is followed by the purchase and or development of software. This paradigm has several disadvantages for the development of systems to
support NASA's new vision. Building a system to be fault tolerant increases the complexity and decreases the performance of included software. Standard bus design and
conventional implementation produces natural bottlenecks. Configuring hardware components in systems containing common processors and DSPs is difficult initially and
expensive or impossible to change later. The existence of Hardware Description Languages (HDLs), the recent increase in performance, density and radiation tolerance of
Field Programmable Gate Arrays (FPGAs), and Intellectual Property (IP) Cores provides the technology for reprogrammable Systems on a Chip (SOC). This technology
supports a paradigm better suited for NASA's vision. Hardware and software production are melded for more effective development they can both evolve together over time.
Designers incorporating this technology into future avionics can benefit from its flexibility. Systems can be designed with improved fault isolation and tolerance using
hardware instead of software. Also, these designs can be protected from obsolescence problems where maintenance is compromised via component and vendor
availability. To investigate the flexibility of this technology, the core of the Central Processing Unit and Input Output Processor of the Space Shuttle AP101S Computer were
prototyped in Verilog HDL and synthesized into an Altera Stratix FPGA.
Advanced Avionics
Electronics – Radiation-Hard/Resistant Electronics
Company Name
Microcosm, Inc.
Title
Plug-and-Play Star Sensor for Rapid
Spacecraft Integration
Field Center
ARC
NASA SBIR Phase I
Quad Chart
Microcosm, with Space Micro., and HRP Systems will design a plug-and-play (PnP) star sensor for small satellites. All three
companies are well experienced in developing PnP systems for the U. S. Air Force during the past 6 years. On an existing Phase
II Air Force SBIR program, Microcosm built a prototype star sensor called MicroMak. The sensor proposed here will focus on
PnP compatibility for NASA missions of interest. The PnP star sensor recurring cost target is $150 K to $200 K, with a mass
between 0.5 and 1 kg. Expected NASA mission applications necessitate a modified version of the baseline MicroMak sensor,
including: 1) Interfaces compatible with a new PnP avionics architecture, 2) radiation-hardened focal plane arrays (FPAs) and
processing electronics to enable longer mission life. The baseline MicroMak sensor was designed with inherent radiation-tolerant
features: complimentary metal oxide semiconductor (CMOS) FPAs with no direct space view, and all-reflective optical elements.
The new PnP star sensor will build on MicroMak heritage and provide a modular, PnP-compatible, long-life star sensor for NASA
missions, at low cost compared with traditional star sensors. A cost and mass reduction of a factor of 2 or more over traditional
sensors is expected.
Advanced Avionics
Information – Computer System Architectures
Company Name
Emergent Space Technologies, Inc.
Title
Ground Enterprise Management System
Field Center
ARC
NASA SBIR Phase I
Quad Chart
Spacecraft ground systems are on the cusp of achieving "plug-and-play" capability, i.e., they are approaching the state in which the
various components can be quickly integrated with near-automatic interoperability. When properly architected, such systems offer
advantages over their traditional counterparts, such as improved upgradeability and extensibility. They can also be more easily
automated and can provide better fault tolerance. These characteristics are important for all NASA spacecraft, from Exploration to
Science. However, plug-and-play systems also pose some interesting challenges that can undermine their effectiveness. For
example, they can be more complex in terms of information management and system administration. This is especially true when
automation is used to reduce the workload of the operators. In fact, as the number of components increases, as much "situational
awareness" of the ground system is needed the spacecraft. A software framework that supports plug-and-play integration, while
also providing information management and system coordination, is required. Emergent Space Technologies Inc. therefore
proposes to research and develop an Enterprise Management System for spacecraft ground systems. Called GEMS, it will provide
spacecraft controllers and crew members with "situational awareness" of the current state of the ground system as well as
understanding of how events and automated actions are affecting the system in real-time. The innovation lies in the development of
algorithms and software adapters that gather data from the various components of the ground system
to construct a data model that captures the system state and displays it to the controllers.
Advanced Avionics
Information – Computer System Architectures
Company Name
Seakr Engineering Inc.
Title
LOW COST, TAILOR ABLE AVIONICS FOR RAPID
RESPONSE SATELLITES
Microcosm, Inc.
APPLICATION OF SELF-CONFIGURING NETWORK FOR
PLUG-AND-PLAY DEVICE CONTROL
DOD SBIR Phase II
Quad Chart
The RCC architecture reduces risk, costs, and schedule for satellite missions by providing a reconfigurable space based platform
that could be used for a number of different spacecraft avionic applications. It provides a platform that can not only support rapid,
low cost plug-and-play integration into a spacecraft but because it is reconfigurable, it may also be used for many different
subsystems.
Microcosm proposes to develop and fly a prototype self-configuring, fault tolerant interconnect communication system, both
hardware and software, that can be used on next generation spacecraft for networking basic sensor systems as well as the
complete attitude control system (ACS). Development will be completed in Phase II, yielding a prototype network ready to migrate
to a technology demonstration spacecraft for ACS application.
Advanced Avionics
Information – Computer System Architectures
WIPO
Patent Number
WO07057189A1
Title
Modular Avionics System
Of An Aircraft
Assignee
Abstract
Airbus Deutschland In a modular avionics system comprising several cabinets that are arranged at various locations in an aircraft and that are interconnected in a network,
Gmbh
which cabinets are used for controlling or processing signals from and to sensors, actuators and other systems of the aircraft, it is proposed that the
system comprise parallel processors, for example transputers; the cabinets comprise at least two core processor modules (CPMl, CPM2) and at least
two input/output modules (IOM1, IOM2); the in- put/output modules (IOM1, IOM2) serve as interfaces to the systems to be con¬ trolled, and serve
for the control and intermediate storage of the data flowing into and out of the cabinet; each core processor module (CPMl , CPM2) communicates
independently with each IOM module and CPM module by way of links; and in each core processor a number of independent system programs works
under the control of an operating system. By being able to do without the backplane bus that is required in conventional systems the efficiency is
enhanced and changing applications is facilitated.
Advanced Avionics
Information – Software Tools for Distribution and Analysis and Simulation
Company Name
Continuum Dynamics, Inc.
Title
Serial In-Line Instrumentation Bus
for ROV Engineering Research
Field Center
LARC
NASA SBIR Phase I
Quad Chart
Advanced microcontrollers having digital signal processing features have enabled the capability to distribute on-board computation for remotely
operated vehicles (ROVs). Distributed processing can result in a lighter weight avionics suite with improved performance, by locating data conversion
units adjacent to the sensors and control actuators, and reducing EMI through minimization of the amount of interconnection wiring. The proposed
work will leverage CDI's and AMDI's substantial prior experience in the development and operation of flight control avionics for ROVs in the design of
a new system for supporting advanced research using these systems. The avionics suite to be developed consists of serially interconnected
distributed nodes that may be programmed through a Matlab graphical interface to perform control and sensing functions in support of custom
requirements from the research community. The flexibility of custom-configured distributed computing nodes for use in a research context ensures that
"just enough" instrumentation and control is provided for the specific test requirements at hand. Phase I will provide risk reduction by demonstrating
the operation of the subcomponent technologies, culminating in a simplified flight test of the avionics system. Phase II continuation will develop the
complete system to support testing activities at a NASA research center of interest.
Advanced Avionics
Materials –Semi-Conductors/Solid State Device Materials
FC
Jet Propulsion Laboratory
Title
Spacecraft avionics miniaturization using embedded passive
packaging technology
NTRS
Abstract
"Miniaturization of electronic packages will play a key role in future space avionics systems. Smaller avionics packages will reduce payloads
while providing greater functionality for information processing and mission instrumentation. Current surface mount technology discrete passive
devices not only take up significant space but also add weight. To that end, the use of embedded passive devices, such as capacitors,
inductors and resistors will be instrumental in allowing electronics to be made smaller and lighter. Embedded passives fabricated on silicon or
like substrates using thin film technology, promise great savings in circuit volume, as well as potentially improving electrical performance by
decreasing parasitics. These devices exhibit a low physical profile and allow the circuit footprint to be reduced by stacking passive elements
within a substrate. Thin film technologies used to deposit embedded passive devices are improving and costs associated with the process are
decreasing.
There are still many challenges with regard to this approach that must be overcome. In order to become a viable approach these devices need
to work in conjunction with other active devices such as bumped die (flip chip) that share the same substrate area. This dictates that the
embedded passive devices are resistant to the subsequent assembly processes associated with die attach (temperature, pressure). Bare die
will need to be mounted directly on top of one or more layers of embedded passive devices. Currently there is not an abundant amount of
information available on the reliability of these devices when subjected to the high temperatures of die attach or environmental thermal cycling
for space conditions. Device performance must be consistent over time and temperature with minimal parasitic loss.
Pretested and assembled silicon substrates with layers of embedded capacitors made with two different dielectric materials, Ta2Os (TaO) and
benzocyclobutene (BCB), were subjected to the die attach process and tested for performance in an ambient environment. These assemblies
were subjected to environmental thermal cycling from -55°C to 125°C. Other substrates were assembled and tested with both embedded
passive and surface mount active devices. Preliminary results indicate embedded passive capacitors and resistors can fulfill the performance
and reliability requirements of space flight on future missions. Testing results are encouraging for continued development of integrating
embedded passive devices to replace conventional electronic packaging methods. "
Advanced Avionics
Power and Energy – Power Management and Distribution
Company Name
Design Net Engineering LLC
Title
Power Management for PnP Spacecraft
DOD SBIR Phase I
Quad Chart
This research work will produce an innovative Power Management and Distribution (PMAD) system design that can support both heritage satellite
components and the new power compatible plug-and-play (PnP) methodology. The proposed PMAD system design will support a significant range of
power requirements for satellites with fewer power switches, fewer current monitors and simpler hierarchical management. In addition, a sophisticated
simulation tool will be developed to assist design engineers at the AFRL Testbed with component selection, confirm power system is sized correctly and
mission requirements are met. Anticipated benefits of PMAD system are reduced design time, simpler integration, rapid mission configuration, more
autonomous operation, and reduced mission operations as well as decreased cost and schedule. This study will define standards for both PMAD and
PnP power components to assure compatibility during integration, enhance mission and systems design tools, implement PnP power standards for
xTEDS, develop a PMAD “smart” architecture design that supports the PnP principle and the power management tools that allow autonomous on-orbit
operation, as well as demonstrate utility and efficacy of those tools with “day in the life” simulations and provide a mature design for a prototype flight-like
device for Phase II efforts.
Advanced Avionics
Power and Energy – Power Management and Distribution
FC
GRC
Title
Mission Applicability Assessment of
Integrated Power Components and
Systems
Marshall Space
Flight Center
Modular, Reconfigurable, HighEnergy Technology Development
Jet Propulsion
Laboratory
New Generation Power System for
Space Applications
NTRS
Abstract
The need for smaller lightweight autonomous power systems has recently increased with the increasing focus on micro- and nanosatellites. Small area high-efficiency thin film
batteries and solar cells are an attractive choice for such applications. The NASA Glenn Research Center, Johns Hopkins Applied Physics Laboratory, Lithium Power Technologies,
MicroSat Systems, and others, have been working on the development of autonomous monolithic packages combining these elements or what are called integrated power supplies
(IPS). These supplies can be combined with individual satellite components and are capable of providing continuous power even under intermittent illumination associated with a
spinning or Earth orbiting satellite. This paper discusses the space mission applicability, benefits, and current development efforts associated with integrated power supply
components and systems. The characteristics and several mission concepts for an IPS that combines thin-film photovoltaic power generation with thin-film lithium ion energy storage
are described. Based on this preliminary assessment, it is concluded that the most likely and beneficial application of an IPS will be for small 'nanosatellites' or in specialized
applications serving as a decentralized or as a distributed power source or uninterruptible power supply.
The Modular, Reconfigurable High-Energy (MRHE) Technology Demonstrator project was to have been a series of ground-based demonstrations to mature critical technologies
needed for in-space assembly of a highpower high-voltage modular spacecraft in low Earth orbit, enabling the development of future modular solar-powered exploration cargotransport vehicles and infrastructure. MRHE was a project in the High Energy Space Systems (HESS) Program, within NASA's Exploration Systems Research and Technology (ESR
and ampT) Program. NASA participants included Marshall Space Flight Center (MSFC), the Jet Propulsion Laboratory (JPL), and Glenn Research Center (GRC). Contractor
participants were the Boeing Phantom Works in Huntsville, AL, Lockheed Martin Advanced Technology Center in Palo Alto, CA, ENTECH, Inc. in Keller, TX, and the University of AL
Huntsville (UAH). MRHE's technical objectives were to mature (a) lightweight, efficient, high-voltage, radiation-resistant solar power generation (SPG) technologies (b) innovative,
lightweight, efficient thermal management systems (c) efficient, 100kW-class, high-voltage power delivery systems from an SPG to an electric thruster system (d) autonomous
rendezvous and docking technology for in-space assembly of modular, reconfigurable spacecraft (e) robotic assembly of modular space systems and (f) modular, reconfigurable
distributed avionics technologies. Maturation of these technologies was to be implemented through a series of increasingly-inclusive laboratory demonstrations that would have
integrated and demonstrated two systems-of-systems (a) the autonomous rendezvous and docking of modular spacecraft with deployable structures, robotic assembly, reconfiguration
both during assembly and (b) the development and integration of an advanced thermal heat pipe and a high-voltage power delivery system with a representative lightweight highvoltage SPG array. In addition, an integrated simulation testbed would have been developed containing software models representing the technologies being matured in the laboratory
demos. The testbed would have also included models for non-MRHE developed subsystems such as electric propulsion, so that end-to-end performance could have been assessed.
This paper presents an overview of the MRHE Phase I activities at MSFC and its contractor partners. One of the major Phase I accomplishments is the assembly demonstration in the
Lockheed Martin Advanced Technology Center (LMATC) Robot-Satellite facility, in which three robot-satellites successfully demonstrated rendezvous and amp docking, selfassembly, reconfiguration, adaptable GN and ampC, deployment, and interfaces between modules. Phase I technology maturation results from ENTECH include material
recommendations for radiation hardened Stretched Lens Array (SLA) concentrator lenses, and a design concept and test results for a hi-voltage PV receiver. UAH's accomplishments
include Supertube heatpipe test results, which support estimates of thermal conductivities at 30,000 times that of an equivalent silver rod. MSFC performed systems trades and
developed a preliminary concept design for a 100kW-class modular reconfigurable solar electric propulsion transport vehicle, and Boeing Phantom Works in Huntsville performed
assembly and rendezvous and docking trades. A concept animation video was produced by SAIC, wllich showed rendezvous and docking and SLA-square-rigger deployment in LEO.
The Deep Space Avionics (DSA) Project is developing a new generation of power system building blocks. Using application specific integrated circuits (ASICs) and power switching
modules a scalable power system can be constructed for use on multiple deep space missions including future missions to Mars, comets, Jupiter and its moons. The key
developments of the DSA power system effort are five power ASICs and a module for power switching. These components enable a modular and scalable design approach, which
can result in a wide variety of power system architectures to meet diverse mission requirements and environments. Each component is radiation hardened to one megarad total dose.
The power switching module can be used for power distribution to regular spacecraft loads, to propulsion valves and actuation of pyrotechnic devices. The number of switching
elementsper load, pyrotechnic firings and valve drivers can be scaled depending on mission needs. Telemetry data is available from the switch module via an I2C data bus. The DSA
power system components enable power management and distribution for a variety of power buses and power system architectures employing different types of energy storage and
power sources. This paper will describe each power ASIC#s key performance characteristics as well as recent prototype test results. The power switching module test results will be
discussed and will demonstrate its versatility as a multipurpose switch. Finally, the combination of these components will illustrate some of the possible power system architectures
achievable from small single string systems to large fully redundant systems.
Advanced Avionics
Power and Energy – Power Management and Distribution
FC
Jet Propulsion
Laboratory
Title
Power Actuation and Switching
Module Development
Jet Propulsion
Laboratory
Power Actuation and Switching
Module Test Results
NTRS
Abstract
The Deep Space Avionics (DSA) Project is developing a Power Actuation and Switching Module (PASM). This component enables a modular and scalable design approach for power
switching applications, which can result in a wide variety of power switching architectures using this simple building block. The PASM is designed to provide most of the necessary
power switching functions of spacecraft for various Deep Space missions including future missions to Mars, comets, Jupiter and its moons. It is fabricated using an ASIC process that
is tolerant of high radiation. The development includes two application specific integrated circuits (ASICs) and support circuitry all packaged using High Density Interconnect (HDI)
technology. It can be operated in series or parallel with other PASMs, It can be used as a high-side or low-side switch and it can drive thruster valves, pyrotechnic devices such as
NASA standard initiators, bus shunt resistors, and regular spacecraft component loads. Each PASM contains two independent switches with internal current limiting and over-current
trip-off functions to protect the power subsystem from load faults. During turnon and turnoff each switch can limit the rate of current change (di dt) to a value determined by the user.
Threeway majority-voted On Off commandability and full switch status telemetry (both analog and digital) are built into the module. This paper describes the development process
used to design, model, fabricate, and test these compact and versatile power switches. Preliminary test results from prototype HDI PASM hardware are also discussed.
The X2000 Power System Electronics (PSE) is a Jet Propulsion Laboratory (JPL) task to develop a new generation of power system building blocks for use on future deep-space
missions. The effort includes the development of electronic components and modules that can be used as building blocks in the design of generic spacecraft power systems. All
X2000 avionics components and modules are designed for use in centralized or distributed spacecraft architectures. The Power Actuation and Switching Module (PASM) has been
developed under the X2000 program. This component enables a modular and scalable design approach for power switching applications, which can result in a wide variety of power
switching architectures using this simple building block. The PASM is designed to provide most of the necessary power switching functions of spacecraft for various Deep Space
missions including future missions to Mars, comets, Jupiter and its moons. It is fabricated using an ASIC process that is tolerant of high radiation. The development included two
application specific integrated circuits (ASICs) and support circuitry all packaged using High Density Interconnect (HDI) technology. It can be operated in series or parallel with other
PASMs. It can be used as a high-side or low-side switch and it can drive thruster valves, pyrotechnic devices such as NASA standard initiators, bus shunt resistors, and regular
spacecraft component loads. Each PASM contains two independent switches with internal current limiting and over-current trip-off functions to protect the power subsystem from load
faults. During turnon and turnoff each switch can limit the rate of current change (di dt) to a value determined by the user. Three-way majority-voted On Off commandability and full
switch status telemetry (both analog and digital) are built into the module. This paper is a follow up to the one presented at he IECEC 2004 conference that will include the lessons
learned and test results from the development.
Advanced Avionics
Power and Energy – Power Management and Distribution
Company Name
AeroAstro, Inc.
Title
AUTONOMOUS SPACECRAFT POWER
SCHEDULING
NASA SBIR Phase II
Field Center
Quad Chart
ARC
Designed in collaboration with the Univ. of Colorado, the Autonomous Power Scheduler (APS) will dynamically
monitor, schedule and distribute power on microsatellites. The APS represents an innovative paradigm shift from
conventional designs, controlling primary power in an intelligent, semi- or fully autonomous way to maximize mission
capability and performance. With APS, appropriate power will be distributed when and where it is
most needed.
Advanced Avionics
Propulsion – Beamed Energy
U.S. Patent Applications
Patent Number
Title
Assignee
Abstract
US2007045474A1
System and method for
propellantless photon tether
formation flight
None
The invention is a system and method for propellantless, ultrahigh precision satellite formation flying based on ultrahigh precision intracavity laser
thrusters and tethers with an intersatellite distance accuracy of nanometers at maximum estimated distances of tens of kilometers. The repelling force
of the intracavity laser thruster and the attracting force of tether tension between satellites form the basic forces to stabilize matrix structures of
satellites. Users of the present invention can also use the laser thruster for ultrahigh precision laser interferometric metrology, resulting in
simplification and payload weight reduction in integrating the thruster system and the metrology system.
Advanced Avionics
Robotics – Integrated Robotic Concepts and Systems
FC
JSC
Title
AERCam Autonomy Intelligent Software
Architecture for Robotic Free Flying Nanosatellite
Inspection Vehicles
NTRS
Abstract
TThe NASA Johnson Space Center has developed a nanosatellite-class Free Flyer intended for future external inspection and remote viewing of
human spacecraft. The Miniature Autonomous Extravehicular Robotic Camera (Mini AERCam) technology demonstration unit has been integrated into
the approximate form and function of a flight system. The spherical Mini AERCam Free Flyer is 7.5 inches in diameter and weighs approximately 10
pounds, yet it incorporates significant additional capabilities compared to the 35-pound, 14-inch diameter AERCam Sprint that flew as a Shuttle flight
experiment in 1997. Mini AERCam hosts a full suite of miniaturized avionics, instrumentation, communications, navigation, power, propulsion, and
imaging subsystems, including digital video cameras and a high resolution still image camera. The vehicle is designed for either remotely piloted
operations or supervised autonomous operations, including automatic stationkeeping, point-to-point maneuvering, and waypoint tracking. The Mini
AERCam Free Flyer is accompanied by a sophisticated control station for command and control, as well as a docking system for automated
deployment, docking, and recharge at a parent spacecraft. Free Flyer functional testing has been conducted successfully on both an airbearing table
and in a six-degree-of-freedom closed-loop orbital simulation with avionics hardware in the loop. Mini AERCam aims to provide beneficial on-orbit views
that cannot be obtained from fixed cameras, cameras on robotic manipulators, or cameras carried by crewmembers during extravehicular activities
(EVA s). On Shuttle or International Space Station (ISS), for example, Mini AERCam could support external robotic operations by supplying orthogonal
views to the intravehicular activity (IVA) robotic operator, supply views of EVA operations to IVA and or ground crews monitoring the EVA, and carry out
independent visual inspections of areas of interest around the spacecraft. To enable these future benefits with minimal impact on IVA operators and
ground controllers, the Mini AERCam system architecture incorporates intelligent systems attributes that support various autonomous capabilities. 1) A
robust command sequencer enables task-level command scripting. Command scripting is employed for operations such as automatic inspection scans
over a region of interest, and operator-hands-off automated docking. 2) A system manager built on the same expert-system software as the command
sequencer provides detection and smart-response capability for potential system-level anomalies, like loss of communications between the Free Flyer
and control station. 3) An AERCam dynamics manager provides nominal and off-nominal management of guidance, navigation, and control (GN and
ampC) functions. It is employed for safe trajectory monitoring, contingency maneuvering, and related roles. This paper will describe these architectural
components of Mini AERCam autonomy, as well as the interaction of these elements with a human operator during supervised autonomous control.
Advanced Avionics
Sensors and Sources – Sensor Webs/Distributed Sensors
FC
JPL
Title
Efficient and Robust Data Collection Using Compact Micro
Hardware, Distributed Bus Architectures and Optimizing
Software
NTRS
Abstract
Future In-Space propulsion systems for exploration programs will invariably require data collection from a large number of sensors.
Consider the sensors needed for monitoring several vehicle systems states of health, including the collection of structural health data,
over a large area. This would include the fuel tanks, habitat structure, and science containment of systems required for Lunar, Mars, or
deep space exploration. Such a system would consist of several hundred or even thousands of sensors. Conventional avionics system
design will require these sensors to be connected to a few Remote Health Units (RHU), which are connected to robust, micro flight
computers through a serial bus. This results in a large mass of cabling and unacceptable weight. This paper first gives a survey of
several techniques that may reduce the cabling mass for sensors. These techniques can be categorized into four classes power line
communication, serial sensor buses, compound serial buses, and wireless network. The power line communication approach uses the
power line to carry both power and data, so that the conventional data lines can be eliminated. The serial sensor bus approach reduces
most of the cabling by connecting all the sensors with a single (or redundant) serial bus. Many standard buses for industrial control and
sensor buses can support several hundreds of nodes, however, have not been space qualified. Conventional avionics serial buses
such as the Mil-Std-1553B bus and IEEE 1394a are space qualified but can support only a limited number of nodes. The third
approach is to combine avionics buses to increase their addressability. The reliability, EMI EMC, and flight qualification issues of
wireless networks have to be addressed. Several wireless networks such as the IEEE 802.11 and Ultra Wide Band are surveyed in this
paper. The placement of sensors can also affect cable mass. Excessive sensors increase the number of cables unnecessarily.
Insufficient number of sensors may not provide adequate coverage of the system. This paper also discusses an optimal technique to
place and validate sensors.
Advanced Avionics
Structures – Modular Interconnects
Company Name
Spaceworks, Inc.
Title
Reconfigurable PnP Spacecraft
Structure
DOD SBIR Phase I
Quad Chart
Plug-and-play avionics for spacecraft are being developed by the Air Force to create systems that promise revolutionary improvements in responsiveness for military
users. To keep pace, similar developments are needed in other supporting systems. SpaceWorks proposes to develop plug-and-play structures for spacecraft that have
features that support these advanced avionics systems and that also contribute to the overall goals of reconfigurability, flexibility, modularity, and rapid assembly,
integration, and checkout. The approach involves a small number of spacecraft panel types that can be assembled rapidly to form many useful spacecraft
configurations. The panels have identical mechanical and electrical interfaces at discrete sites and between panels to enable rapid plug-and-play functionality for
payloads and components. The plug-and-play electrical network consisting of electronics, harnessing, and connectors is recessed within each panel to maximize
useable volume and footprint. Panel design and materials maximize the flexibility to incorporate multifunctional features such as thermal management devices,
integrated structural sensors, and layered radiation shielding. SpaceWorks will fabricate a set of these panels leading to a demonstration showing how the approach
supports plug-and-play avionics, provides a high degree of reconfigurability, and can meet a very demanding integration and checkout timeline.
Mini-Micro Thrusters, LOX/Hydrocarbon Propulsion, and Attitude Control Systems
Avionics and Astrionics – Attitude Determination and Control
Company Name
Aerophysics, Inc.
Title
Semi-autonomous Local Space Imaging System
DOD SBIR Phase I
Quad Chart
Aerophysics, together with subcontractor Raytheon Missile Systems, proposes to develop a Local Space Imaging System (LSIS) based on
the proven AIM-9X Sidewinder missile gimbaled seeker. The LSIS will combine an AIM-9X-derived imager with an autonomous attitude
determination system to provide pointing/tracking/imaging capability independent of host vehicle orientation and without interface with the
host vehicle attitude control system. The LSIS instrument will employ a low-power, compact star tracker, sun sensor, and three-axis
magnetometer to measure the instrument’s inertial attitude. An on-board command and data handling subsystem will coordinate imager
pointing and host vehicle data interface. The proposed development effort integrates proven technology in a low-mass, low-power, bolt-on
sensor for defensive space situational awareness.
Mini-Micro Thrusters, LOX/Hydrocarbon Propulsion, and Attitude Control Systems
Avionics and Astrionics – Attitude Determination and Control
Company Name
Dynamic Structures &
Materials, LLC
Title
FAST-ACTING, COMPACT, PIEZOELECTRIC
ACTUATOR FOR CONTROL OF MINI-INTERCEPTOR
DOD SBIR Phase II
Quad Chart
To perform the rapid maneuvers and precise trajectory control required for the MKV mission, it is desirable that thruster valve throttling actuators be
able to respond with authority in less than 10 milliseconds. An innovative compact piezoelectric actuator and control system has been developed to
control the main divert and attitude control thrusters with controlled response times in the sub-5 millisecond range.
Mini-Micro Thrusters, LOX/Hydrocarbon Propulsion, and Attitude Control Systems
Avionics and Astrionics – Attitude Determination and Control
U.S. Patent Applications
Patent Number
Title
Assignee
Abstract
US2002003193A1
Safing mode for high
momentum states in
body stabilized
spacecraft
None
US20020148930A1
Maneuver device for
artificial satellite
None
US2002020785A1
Spacecraft orbit
control using orbit
position feedback
None
US2002040950A1
System and method
for spacecraft attitude
control
None
A spacecraft with a reaction wheel system can be autonomously safed by setting the solar wings to continuous tracking, determining a slew rate vector
based on the total angular momentum, and slewing the spacecraft using the slew rate vector until commanded to stop autonomous safing. When the total
angular momentum of the spacecraft to large to be handled by rotisserie, then the spacecraft is reoriented to align a suitable rotation vector with the
system momentum. In a typical application, the spacecraft has a reaction wheel assembly with four wheels arranged to form a right regular pyramid.
Two reaction wheels on opposite edges of the pyramid form a first pair and the two remaining reaction wheels forming a second pair. The slew axis of
rotation is found by determining as a selected pair the first pair if either reaction wheel in the second pair is inoperative, otherwise determining as the
selected pair the second pair and determining as the slew axis of rotation the normalized projection of the axes of rotation of the selected pair onto the
base. The slew direction is determined by the sign of the total angular momentum component along the slew axis of rotation.
"The present invention provides a maneuver device for an artificial satellite, which causes small attitude error during maneuver and which requires a
shorter period of setting time for obtaining a target attitude.The maneuver device is provided with: a feed forward torque instruction signal generator 8
which outputs a feed forward torque instruction signal 11 based on a maneuver plan; a thruster 10 which outputs control torque based on the feed
forward torque instruction signal 11; and an attitude control signal calculator 6 to which an attitude angle and an angular velocity of the artificial satellite
as well as a target attitude angle and a target angular velocity are input and which outputs an attitude control signal 13. The maneuver device is further
provided with a disturbance compensating signal calculator to which the feed forward torque instruction signal 11 and a detected angular velocity signal
16 are input, and which generates and outputs a disturbance compensating signal 12. The maneuver device is yet further provided with a reaction wheel
7 which generates control torque based on the attitude control signal 13 and the disturbance compensating signal 12.
"
A spacecraft orbit control system and method for controlling the orbit of a spacecraft during orbit raising includes processing spacecraft position data to
meet the spacecraft attitude sensing needs for long duration, low thrust orbit raising bums. The actual spacecraft position as determined by a global
positioning system (GPS) receiver is compared with the desired spacecraft position to generate an error signal indicative of a spacecraft position error for
adjusting the attitude of the spacecraft as thrusters move the spacecraft. The attitude of the spacecraft is adjusted to eliminate the spacecraft position
error such that the actual spacecraft position corresponds with the desired spacecraft position and the spacecraft maintains a desired orbit during the orbit
raising.
A momentum management system for attitude control of a spacecraft includes a housing to be fixed to the spacecraft and a momentum wheel rotor in the
housing for storing angular momentum. A gimbal assembly mounts the rotor in the housing. The rotor is driven by a drive with its output coupled to the
rotor. A torque generation imparts torque to the rotor about axes orthogonal to the drive axis. The gimbal assembly includes a gimbal ring coupling the
drive output to the rotor. The gimbal ring in turn includes flexure joints connecting the gimbal ring to the drive and the rotor. The flexure joints are
configured to permit the rotor to tilt about two flexure axes orthogonal to the drive axis to incline the rotor axis through a range of angles from about 0
degrees to about 7 degrees with respect to the drive axis under the control of said torque generation device. The preferred flexure joint is formed from
two resilient, crossing webs. The webs have ring ends connected to the body of the gimbal ring and mounting ends connected to either the drive or the
rotor. The system includes a launch restraint system to limit movement of the rotor along the drive axis, including a stop mounted on the drive output
and a cage mounted on the rotor, surrounding the stop. Under high acceleration, the cage engages the stop to support the rotor, relieving excess stress on
the flexures of the gimbal suspension. The launch restraint system also includes deflection stops adjacent opposite sides of each web of each flexure for
limiting deflection of the webs.
Mini-Micro Thrusters, LOX/Hydrocarbon Propulsion, and Attitude Control Systems
Avionics and Astrionics – Attitude Determination and Control
U.S. Patent Applications
Patent Number
Title
Assignee
Abstract
US2002077732A1
6-degree-of-freedom
control apparatus for
spacecraft
None
US2002121573A1
Method of controlling Agence Spatiale
or stabilizing the
Europeenne
attitude of a vehicle in
space
A 6-degree-of-freedom control apparatus for a spacecraft includes a plurality of thrusters, thruster modulator, position/velocity detector, target
position/velocity generator, attitude/angular velocity detector, target attitude/angular velocity generator, and 6-degree-of-freedom controller. The
thrusters control three position axes and three attitude axes of a spacecraft by jet. The thruster modulator selectively drives the thrusters on the basis of a
thruster control signal. The position/velocity detector measures a position and velocity of the spacecraft. The target position/velocity generator generates
target position and velocity values of the spacecraft. The attitude/angular velocity detector measures an attitude and angular velocity of the spacecraft.
The target attitude/angular velocity generator generates target attitude and angular velocity values of the spacecraft. The 6-degree-of-freedom controller
generates the thruster control signal on the basis of a deviation between an output from the position/velocity detector and an output from the target
position/velocity generator and a deviation between an output from the attitude/angular velocity detector and an output from the target attitude/angular
velocity generator and outputs the signal to the thruster modulator.
A method of controlling the attitude of a vehicle in space relative to three reference axes, the method consisting: initially in adjusting the attitude of the
vehicle in conventional manner to desired angles on the basis of a two-axis detector which points towards a reference heavenly body; and subsequently
in adjusting the attitude of the vehicle about the third reference axis by using the speeds of rotation of the reaction wheels which reflect conservation of
the total angular momentum of the vehicle to perform an angle measurement about the third axis.
US2002139902A1
Low-thrust cryogenic
propulsion module
SNECMA
MOTEURS
The cryogenic propulsion module comprises a main cryogenic thruster (10), two attitude-controlling secondary thrusters (21, 22), tanks (31, 32, 33, 34)
for feeding cryogenic propellants, a device for intermittently pressurizing the tanks (31, 32, 33, 34), and a device for initiating firing of the main
cryogenic thruster (10) in intermittent manner while the tanks (31, 32, 33, 34) are intermittently pressurized. The device for intermittently pressurizing a
tank (31, 32, 33, 34) comprises a heat exchange circuit associated with a heat accumulator (61, 62) and a device (71, 72) for putting a predetermined
quantity of a propellant into circulation in the heat exchanger circuit. The module also comprises a device for heating the heat accumulator (61, 62) in
the periods between two consecutive firings.
US2003149529A1
Fault detection
pseudo gyro
The Aerospace
Corporation
A fault detection pseudo gyro emulates mechanical gyros by processing space system appendage measurement data and reaction wheel tachometer data
within reference and control systems of a satellite to compute the vehicular bus angular velocity rate data by accounting for external torques, the
momentum transfer between the satellite, the bus, and the appendages for providing accurate relative vehicular position and angular velocity rate data in
an attitude reference and control systems now having accurate vehicular angular velocity rate data while enabling fault detection of hardware gyros and
momentum sensors.
US2003164429A1
Satellite Harmonic
Torque Estimator
MCGOVERN
LAWRENCE|PRICE
XEN
A spacecraft embedded in a reference frame rotating relative to inertial space. The spacecraft generally includes actuators for maneuvering the
spacecraft with respect to the reference frame, an attitude measurement device that measures the pitch and roll attitude of the spacecraft with respect to
the reference frame, a control device adapted to keep the roll and pitch angles of the spacecraft close to the commanded roll and pitch angles, and a
harmonic torque estimator adapted to read the commanded angular velocity of the spacecraft relative to an inertial frame, read momentum wheel speeds,
read known predicted external torques and combines angular velocity, measured wheel speed and known external torque to produce an estimated
observable periodic torque.
Mini-Micro Thrusters, LOX/Hydrocarbon Propulsion, and Attitude Control Systems
Avionics and Astrionics – Attitude Determination and Control
U.S. Patent Applications
Patent Number
Title
Assignee
US2003195674A1
Low energy method
for changing the
inclinations of
orbiting
satellitesUSing weak
stability boundaries
and a computer
process for
implementing same
Wick Injection of
Liquids for Colloidal
Propulsion
Galaxy Development, A fuel efficient technique for changing the inclination, with respect to the Earth's equator, for a satellite includes first maneuvering the satellite to the
LLC
moon on a BCT (Ballistic Capture Transfer). At the moon, the satellite is in the so-called fuzzy boundary or weak stability boundary. A negligibly small
maneuver can then bring it back to the Earth on a reverse BCT to the desired Earth inclination. Another maneuver puts it into the new ellipse at the earth.
In the case of satellites launched from Vandenberg AFB into LEO in a circular orbit of an altitude of 700 km with an inclination of 34°, approximately 6
km/s is required to change the inclination to 90°. The previous flight time associated with this method was approximately 170 days. A modification of
this method also achieves a significant savings and unexpected benefits in energy as measured by Delta-V, where the flight time is also substantially
reduced to 88 or even 6 days.
US2003209005A1
Abstract
None
Propellant liquid is supplied to a Colloidal Thruster for Micro-Satellite vehicles in Space by capillarity induced flow through a wick element comprising
a permeable porous aggregate of fibers or particles of material that is wetted by the propellant liquid. An intense electric field at the tip of the wick
element dispersed the arriving liquid into a fine spray of charged droplets. Electrodes having appropriate design, location and potentials accelerate the
charge droplets to high velocity, thereby providing reactive thrust to the vehicle. In this method of propellant liquid introduction the flow rate and
exhaust velocity, and therefore the thrust level, are determined by the applied potential difference, thereby eliminating the need for pumps or pressurized
gas and flow controllers to provide the desired flow-rate for the propellant liquid.
US20040104308A1
Attitude
None
determination and
alignmentUSing
electro-optical sensors
and global navigation
satellites
US2004069905A1
High-efficiency REA
optimized
stationkeeping
"Systems and methods of controlling and correcting the attitude of a spacecraft or a spacecraft-based payload due to disturbances that cause attitude
changes determined by a spacecraft attitude measurement sensor and one or more additional attitude measurement sensors on the spacecraft. The
systems and methods measure the attitude of the spacecraft and measure relative attitude errors between the spacecraft and the payload. The
measurements are processed to generate control signals that selectively control the pointing direction of the spacecraft or payload. Embodiments are
disclosed wherein the additional attitude measurement sensor is disposed on a platform that is stable with respect to the spacecraft attitude measurement
sensor, and wherein it is disposed on a platform that is not stable with respect to the spacecraft attitude measurement sensor. Another embodiment is
disclosed having a plurality of spacecraft attitude measurement sensors, one each for the payload and for the spacecraft attitude measurement sensor.
"
A velocity change (ΔV) thruster is operated on a spacecraft, which unavoidably causes attitude error. A reaction wheel (RWA) corrects the attitude. At
the beginning of the thruster maneuver, the total attitude control momentum required to at least correct for the ΔV thruster attitude errors over the
duration of the entire maneuver is determined, and the RWA momentum may also be determined. Attitude control thrusters (REAs) are operated. The
REAs are operated to correct at least the net ΔV thruster induced attitude error, and preferably also to reset the RWA to its nominal momentum. The
maneuver may be stationkeeping.
None
Mini-Micro Thrusters, LOX/Hydrocarbon Propulsion, and Attitude Control Systems
Avionics and Astrionics – Attitude Determination and Control
U.S. Patent Applications
Patent Number
Title
Assignee
Abstract
US2004104308A1
Attitude Correction
Of Disturbances
Relative to a
Spacecraft Attitude
Sensor
None
Systems and methods of controlling and correcting the attitude of a spacecraft or a spacecraft-based payload due to disturbances that cause attitude
changes determined by a spacecraft attitude measurement sensor and one or more additional attitude measurement sensors on the spacecraft. The
systems and methods measure the attitude of the spacecraft and measure relative attitude errors between the spacecraft and the payload. The
measurements are processed to generate control signals that selectively control the pointing direction of the spacecraft or payload. Embodiments are
disclosed wherein the additional attitude measurement sensor is disposed on a platform that is stable with respect to the spacecraft attitude measurement
sensor, and wherein it is disposed on a platform that is not stable with respect to the spacecraft attitude measurement sensor. Another embodiment is
disclosed having a plurality of spacecraft attitude measurement sensors, one each for the payload and for the spacecraft attitude measurement sensor.
US2004140401A1
System and method
for controlling the
attitude of a flying
object
Hydrodynamic
propellantless
propulsion
NEC
CORPORATION
The attitude of a satellite is controlled by at least three reaction wheels having mutually different rotation axes and at least two control moment
gyroscopes having gimbal axes which extend in an equal direction. The reaction wheels and the control momentum gyros are controlled according to
calculated values. When the rotors of the gyroscopes are given rotations in opposite directions to each other, angular momentum vectors are produced in
respective directions normal to the directions of rotation of the rotors about their gimbal axes.
A propellantless hydrodynamic centrifugal thruster (122) comprising a hydrodynamic stator (102) with a hydrodynamic stator chamber (104) housing a
centrifugal thrust generator (106) with a plurality of radial chambers (112A through L), and a propulsion fluid (114). The radial chambers (112A through
L) are distributed in an annular array on one face of the centrifugal thrust generator (106). The centrifugal thrust generator (106) spins at the rotational
speed (WR) and employs the mass of fluid (114) to generate unbalanced centrifugal forces (Fc). The vector sums of all the unbalanced centrifugal force
(Fc) vector components generate a propellantless and unidirectional propulsion force (F).
Thruster for
propelling and
directing a vehicle
without interacting
with environment and
method for making
the same
None
US2004219007A1
US2005077425A1
None
A thruster for propelling and directing a vehicle without interacting with its environment without using propellant and particularly adapted for use in
space, comprising rotating means having a pair of a first and a second axes of rotation, connectable to the vehicle such that each of the first and second
axes of rotation extends opposite to each other with respect to a center of mass of the vehicle. The thruster is also provided with actuator means for
actuating a first and a second rotational movement of the rotating means respectively around each of the first and second axis of rotation. The first
rotational movement is actuated in a clockwise direction thereby generating a first reacting torque in a counter-clockwise direction that causes a pivotal
movement of the vehicle around the first axis of rotation in the counter-clockwise direction. The second rotational movement is actuated in a counterclockwise direction thereby generating a second reacting torque in a clockwise direction that causes a pivotal movement of the vehicle around the second
axis of rotation in the clockwise direction. The thruster is also provided with a control mechanism to control and coordinate the actuator means to impart
propulsion and direction to the vehicle. In a further embodiment, the thruster is enclosed. In a further embodiment, the thruster allows to spin the vehicle.
A method for propelling and directing a vehicle without interacting with its environment is also disclosed.
Mini-Micro Thrusters, LOX/Hydrocarbon Propulsion, and Attitude Control Systems
Avionics and Astrionics – Attitude Determination and Control
U.S. Patent Applications
Patent Number
Title
Assignee
Abstract
US2005109138A1
Inertial propulsion
drive
None
US2005133670A1
Unified Sensor-Based None
Attitude
Determination and
Control for Spacecraft
Operations
Method and System
None
for Optimizing
Torque in a Cmg
Array
An inertial thrust drive (10) comprising a centrifugal thrust generator (12) that comprises a first motor (14); with a weighted arm (16) comprising a radial
arm (18) and a weight (20); a platform (22), a second motor (24); the entire assembly mounted on a thrust mount (26). The motor (14) rotates the
weighted arm (16) in a counterclockwise rotational direction (30) to generate unbalanced centrifugal forces in its plane of rotation. The centrifugal thrust
generator (12) is supported by the platform (22). The platform (22) rotates the thrust generator (12) in a clockwise direction of rotation (32) opposite to
the arm (16) rotational direction (30). Both, the first motor (14) and the second motor (24) rotate about a common central axis (34). To generate a
directional propulsion force (36), the weighted arm (16) generates unbalanced centrifugal forces in its plane of rotation; and the platform (22) rotates the
thrust generator (12) in the opposite direction to maintain the arm (16) pointing in the same direction. The synergy of superimposing the rotational
energy of the platform (22) on the thrust generator (12) generates a directional propulsion force (36). The propulsion force (36) vector is useful as a
source of thrust for propellantless propulsion.
Systems and method of attitude determination and control for spacecraft include the use of a unified set of sensors for all phases of space flight. For
example, the same set of sensors may be used for on-station operations and transfer orbit operations. The use of a unified set of sensors reduces
complexity, spacecraft cost, and spacecraft weight.
US2006022091A1
US2006049314A1
US2006049315A1
US2007023579A1
Quantized controlmoment gyroscope
array
Precision attitude
control system for
gimaled thruster
Unified attitude
control for spacecraft
transfer orbit
operations
None
A momentum-control system for a spacecraft is disclosed. The momentum-control system comprises an attitude-control system. The attitude-control
system receives data concerning a desired spacecraft maneuver and determines a torque command to complete the desired spacecraft maneuver. A
momentum actuator control processor coupled to the attitude-control system receives the torque command. The momentum actuator control processor
calculates a gimbal rate command comprising a range-space gimbal rate required to produce the torque command and a null-space gimbal rate required
to maximize the ability to provide torque in the direction of a current torque. At least four control-moment gyros are coupled to the momentum control
actuator control processor. Each of the control-moment gyros receives and executes the gimbal rate to produce the desired maneuver.
A momentum actuator for steering a spacecraft is disclosed. The momentum actuator comprises a rotor, a gimbal upon which the rotor is mounted, and a
stepper motor coupled to the gimbal and operable to rotate the rotor about the gimbal in a series of steps. In one embodiment of the present invention the
spin rate of the rotor can be varied to provide torque.
Lockheed Martin
A system for providing attitude control with respect to a spacecraft is provided. The system includes a reaction wheel control module configured to
Corporation
control a number of reaction wheel assemblies associated with the spacecraft in order to control attitude, and a maneuver control module configured to
use a number of gimbaled Hall Current thrusters (HCTs) to control the total momentum associated with the spacecraft during an orbit transfer. The total
momentum includes the momentum associated with the reaction wheel assemblies and the angular momentum of the spacecraft. Using the gimbaled
HCTs to control the momentum associated with the reaction wheel assemblies during the orbit transfer results in minimal HCT gimbal stepping.
The Boeing Company A method of controlling attitude of a spacecraft during a transfer orbit operation is provided. The method includes providing a slow spin rate,
determining the attitude of the spacecraft using a unified sensor set, and controlling the attitude of the spacecraft using a unified control law. The use of a
unified set of sensors and a unified control law reduces spacecraft complexity, cost, and weight.
Mini-Micro Thrusters, LOX/Hydrocarbon Propulsion, and Attitude Control Systems
Avionics and Astrionics – Attitude Determination and Control
U.S. Patent Applications
Patent Number
Title
Assignee
Abstract
US2007023580A1
High-torque, low
power reaction wheel
array and method
Honeywell
International, Inc.
US2007124032A1
Method and system
for controlling sets of
collinear control
moment gyroscopes
Honeywell
International, Inc.
Methods and apparatus are provided for reaction wheel (RW) assemblies for spacecraft. The apparatus comprises a M/G coupled to an inertia wheel and
a controller coupled to the M/G that, in response to commands it receives, couples power to or from an M/G of another RW assembly over a shared
transfer connection (V<SUB>XFR</SUB>), so that one M/G acts as a generator powering another M/G acting as a motor. The controller compares
generator voltage V<SUB>GEN </SUB>to the BEMF of the motor on V<SUB>XFR </SUB>and reconfigures a multi-winding M/G or steps-up the
generated voltage to maintain V<SUB>GEN</SUB>>V<SUB>XFR </SUB>to maximize direct energy transfer from one RW to the other as long as
possible. When V<SUB>GEN </SUB>declines sufficiently, the controller can couple the motor to the spacecraft power bus and/or the generator to a
power dump so as to continue to provide the commanded torque if needed. Operation is automatic.
A control system of a spacecraft for controlling two or more sets of collinear control moment gyroscopes (CMGs) comprises an attitude control system.
The attitude control system is configured to receive a command to adjust an orientation of the spacecraft, determine an offset for a momentum disk for
each of the two or more sets of CMGs that maximizes torque, determine a momentum needed from the two or more sets of CMGs to adjust the
orientation of the spacecraft, and calculate a total torque needed by taking the derivative of the momentum. The control system further comprises a
momentum actuator control processor coupled to the attitude control system, the momentum actuator control processor configured to calculate a required
gimbal movement for each of the CMGs in each of the two or more sets of collinear CMGs from total torque.
Mini-Micro Thrusters, LOX/Hydrocarbon Propulsion, and Attitude Control Systems
Avionics and Astrionics – Attitude Determination and Control
USPTO
Patent
Number
US6340137
Title
Assignee
Abstract
Moment control unit for
spacecraft attitude control
Honeywell International
Inc.
A moment control device for positioning a spacecraft which employs a plurality of spinning bodies operable to impart a desired torque to a space
craft, the bodies being constructed in a unitary combination and the unitary combination being mounted to the spacecraft to be positioned.
US6341249
Autonomous unified on-board XING GUANG
orbit and attitude control
QIAN|PARVEZ
system for satellites
SHABBIR AHMED
US6356814
Spacecraft magnetic torquer
feedback system
Microcosm, Inc.
US6454217
Method and apparatus for rate
integration supplement for
attitude referencing with
quaternion differencing
Space Systems/Loral,
Inc.
An apparatus and method of unified orbit and attitude control for acquisition and maintenance techniques of multiple satellites in a formation based
on GPS input, utilizing Modern Feedback Control for providing precise autonomous on-board orbit and attitude control. This control system can
place and maintain any satellite in its designated location in a formation, while simultaneously providing the capability to attain and maintain the
attitude of any of the satellites in the formation with respect to the reference "head-of-fleet' satellite. Utilizing the two different options of GPS
signal, code pseudo range for orbit determination and control and phase pseudorange for attitude determination and control, the relative orbit and
attitude state vectors of all the satellites in the formation is determined and modern advanced multivariable feedback control techniques, for
example, linear quadratic Gaussian/loop transfer recovery controllers for orbit control and Sliding Controller or Lyapunov Controller for attitude
control are used to provide a unified orbit and attitude control. The control of acquisition and maintenance for multiple spacecraft formation flying is
a tracking problem, which can be converted into a regulator problem using the relative orbit and relative attitude kinematics and dynamics.
A magnetic torquer (10) for spacecraft attitude control, providing a torque that is controllable over an extended range greater than a normally used
linear range, and thereby providing a higher ratio of maximum torque to device weight. In one form of the torquer, a feedback loop including a
sensor (20), a signal subtraction circuit (24), a controller (28) and an adjustable power supply (16), continuously determines a corrected command
signal that results in the generation of a desired magnetic moment and a corresponding torque, over the extended range, regardless of non-linearity
across the extended range. In an alternate form of the torquer, the corrected command signal is generated from a mathematical model (56) of the
magnetic torquer (10). Use of the invention also minimizes the effect of any residual magnetic moment that might by present when the actuating
current is reduced to zero.
A control system for providing attitude control in spacecraft. The control system comprising a primary attitude reference system, a secondary
attitude reference system, and a hyper-complex number differencing system. The hyper-complex number differencing system is connectable to the
primary attitude reference system and the secondary attitude reference system.
Mini-Micro Thrusters, LOX/Hydrocarbon Propulsion, and Attitude Control Systems
Avionics and Astrionics – Attitude Determination and Control
USPTO
Patent
Number
US6463365
Title
Assignee
Abstract
System and method for
controlling the attitude of a
space craft
Raytheon Company
US6499699
Satellite attitude control
system and method
Alcatel
A system (30) for adjusting the orientation of a spacecraft adapted for use with a satellite (10). The system (30) includes a first control circuit (32,
38, 40) for canceling any momentum of the spacecraft via a counter-rotating spacecraft bus (16, 18). A second controller (32, 42, 44, 46, 48) orients
the spacecraft via the application of internal spacecraft forces. In a specific embodiment, the spacecraft bus (16, 18) serves a dual use as storage
section and includes a mass (16) having a moment of inertia on the same order as the moment of inertia of the satellite (10). The satellite (10)
includes a bus section (16) and a payload section (14). The mass (16) includes the bus section (16). The first control circuit (32, 38, 40) runs
software to selectively spin the mass (16) to cancel the momentum of the satellite (10). The software computes an actuator control signal, via a
computer (32), that drives a first actuator (38) that spins the mass (16). The first control circuit (32, 38, 40) further includes a circuit for determining
the inertial angular rate of the satellite (10) that includes a gyroscope sensor package (34) in communication with the computer (32). The gyroscope
sensor package (34) provides a rate signal to the computer (32) that is representative of the momentum of the satellite (10). The computer (32) runs
software for generating the actuator control signal in response to the receipt of the rate signal from the gyroscope sensor package (34). The second
controller (32, 42, 44, 46, 48) includes a first reaction wheel (20) having an axis of rotation (26) approximately perpendicular to an axis of rotation
(28) of a second reaction wheel (22). The first and second reaction wheels (20, 22) are rigidly mounted to the spacecraft bus (18, 16) and are free to
spin about their respective axis. The first and second reaction wheels (20, 22) are selectively spun via first and second actuators (44, 48),
respectively, in response to the receipt of first and second steering control signals, respectively.
A satellite attitude control system includes a programmed processor system which includes a gyroscopic actuator first control stage for changing the
attitude of the satellite and a reaction wheel second control stage for assuring that pointing of the satellite is accurate and stable. The method is
intended to be used for a satellite including the two control stages indicated above, which it uses selectively for the operations indicated above.
US6536713
Method of controlling or
stabilizing the attitude of a
vehicle in space
Agence Spatiale
Europeenne
A method of controlling the attitude of a vehicle in space relative to three reference axes, the method consisting:initially in adjusting the attitude of
the vehicle in conventional manner to desired angles on the basis of a two-axis detector which points towards a reference heavenly body;
andsubsequently in adjusting the attitude of the vehicle about the third reference axis by using the speeds of rotation of the reaction wheels which
reflect conservation of the total angular momentum of the vehicle to perform an angle measurement about the third axis.
US6600976
Gyroless control system for
zero-momentum three-axis
stabilized spacecraft
Lockheed Martin
Corporation
US6648274
Virtual reaction wheel array
BAILEY, DAVID
A.|JOHNSON
NORMAN E.
A method for maintaining three axis control of a geosynchronous spacecraft without body rate measurements using reaction wheel assemblies. Earth
sensor assembly angle measurements are utilized for high-bandwidth roll and pitch control. A positive pitch momentum bias is stored in the reaction
wheel assemblies. A gyroscopic feedforward torque is applied to rotate reaction wheel assembly momentum in a yaw/roll plane of the spacecraft at
orbit rate. A dynamic mode that couples yaw and roll axes and that results from applying the gyroscopic feedforward torque and the high-bandwidth
roll control is damped based on earth sensor assembly roll measurements.
An attitude control system (ACS) for a spacecraft includes an attitude control assembly (ACA) interface, a control moment gyro (CMG) array and a
reaction wheel assembly (RWA) control unit. The ACA interface converts torque commands received from the RWA control unit into CMG gimbal
rates for the CMG array. The ACA interface receives CMG gimbal angles from the CMGs and converts the CMG gimbals angles into RWA speeds,
which are then provided to the RWA control unit.
Mini-Micro Thrusters, LOX/Hydrocarbon Propulsion, and Attitude Control Systems
Avionics and Astrionics – Attitude Determination and Control
USPTO
Patent
Number
US6702234
US6732977
US6745984
Title
Assignee
Abstract
Fault tolerant attitude control
system for zero momentum
spacecraft
Lockheed Martin
Corporation
A fault tolerant attitude control system for a zero momentum spacecraft. A zero momentum attitude control system is operable to control spacecraft
attitude utilizing data received from an earth sensor, a sun sensor, and the inertial measurement unit. A gyroless attitude control system is operable
to control spacecraft attitude without receiving data from the inertial measurement unit. A redundancy management system is operable to monitor an
inertial measurement unit to detect faults and to reconfigure the inertial measurement unit if a fault is detected and operable to determine when an
inertial measurement unit fault is resolved. A controller is operable to automatically switch the spacecraft from the zero momentum attitude control
to the gyroless attitude control when a fault in the inertial measurement unit is detected and is operable to automatically switch the spacecraft from
the gyroless attitude control to the zero momentum control upon resolution of the fault.
System for on-orbit correction Lockheed Martin
of spacecraft payload pointing Corporation
errors
Method of controlling the
Astrium Sas
attitude and stabilization of a
satellite in low orbit
AA method for correcting spacecraft pointing errors. An orbit frequency distortion angle between a rate sensor and at least one attitude sensor is
estimated. An estimated attitude of the spacecraft is adjusted to correct for the distortion angle.
For controlling the attitude of a satellite placed on a low earth orbit, components of a vector Bm of the earth's magnetic field along three
measurement axes of a frame of reference bound with the satellite (typically by means of a three-axis magnetometer) are measured. The orientation
of the earth's magnetic field in the frame of reference is computed and a derivative Bm of the vector is also computed. Magneto-couplers carried by
the satellite are energized to create a torque for spinning the satellite at an angular frequency ωc about a determined spin axis of the satellite, where
ωc is greater than an orbital angular frequency 2ω0 of the satellite.
A reorientation controller for a satellite includes a slew rate command generator that generates a slew rate command signal ({right arrow over
(omega)}r-cmd) in response to an attitude error signal. The attitude error signal corresponds to the difference between an initial attitude and a target
attitude. The slew rate command generator may introduce a phase lead (thetaL) into the slew rate command signal ({right arrow over (omega)}rcmd). The controller may perform a spin phase synchronization when the target attitude is unsynchronized in spin phase with the initial attitude. The
satellite may perform a reorientation maneuver in response to the slew rate command signal ({right arrow over (omega)}r-cmd).
US6860451
Spacecraft spin axis
reorientation method
The Boeing Company
US7014150
Method and system for
optimizing torque in a CMG
array
Honeywell International
Inc.
A momentum-control system for a spacecraft is disclosed. The momentum-control system comprises an attitude-control system. The attitude-control
system receives data concerning a desired spacecraft maneuver and determines a torque command to complete the desired spacecraft maneuver. A
momentum actuator control processor coupled to the attitude-control system receives the torque command. The momentum actuator control
processor calculates a gimbal rate command comprising a range-space gimbal rate required to produce the torque command and a null-space gimbal
rate required to maximize the ability to provide torque in the direction of a current torque. At least four control-moment gyros are coupled to the
momentum control actuator control processor. Each of the control-moment gyros receives and executes the gimbal rate to produce the desired
maneuver.
US7057734
Integrated reaction wheel
assembly and fiber optic gyro
Honeywell International
Inc.
The present invention relates to a reaction wheel assembly and fiber optic gyro device. The device includes a reaction wheel assembly having a
reaction wheel assembly housing, a fiber optic gyro coil integrated with the reaction wheel assembly housing, and a fiber optic gyro electronics
integrated with the reaction wheel assembly housing. The fiber optic gyro coil may be wound around the reaction wheel assembly housing. The gyro
coil may also be located within the reaction wheel assembly housing.
Mini-Micro Thrusters, LOX/Hydrocarbon Propulsion, and Attitude Control Systems
Avionics and Astrionics – Attitude Determination and Control
USPTO
Patent
Number
US7097140
Title
Assignee
Abstract
System and method for
spacecraft attitude control
Bristol Aerospace
Limited
US7136752
Method and apparatus for onboard autonomous pair
catalog generation
The Boeing Company
US7219014
Method and apparatus for
real-time star exclusion from
a database
The Boeing Company
A momentum management system for attitude control of a spacecraft includes a housing to be fixed to the spacecraft and a momentum wheel rotor
in the housing for storing angular momentum. A gimbal assembly mounts the rotor in the housing. The rotor is driven by a drive with its output
coupled to the rotor. A torque generation imparts torque to the rotor about axes orthogonal to the drive axis. The gimbal assembly includes a gimbal
ring coupling the drive output to the rotor. The gimbal ring in turn includes flexure joints connecting the gimbal ring to the drive and the rotor. The
flexure joints are configured to permit the rotor to tilt about two flexure axes orthogonal to the drive axis to incline the rotor axis through a range of
angles from about 0 degrees to about 7 degrees with respect to the drive axis under the control of said torque generation device. The preferred
flexure joint is formed from two resilient, crossing webs. The webs have ring ends connected to the body of the gimbal ring and mounting ends
connected to either the drive or the rotor. The system includes a launch restraint system to limit movement of the rotor along the drive axis,
including a stop mounted on the drive output and a cage mounted on the rotor, surrounding the stop. Under high acceleration, the cage engages the
stop to support the rotor, relieving excess stress on the flexures of the gimbal suspension. The launch restraint system also includes deflection stops
adjacent opposite sides of each web of each flexure for limiting deflection of the webs.
A system ( 18 ) includes: a) A vehicle ( 12 ) includes an attitude or angular velocity control system ( 38 ), a plurality of star trackers or star sensors (
22 ) each having a field of view ( 28 ); b) a memory ( 30 ) having a star catalog ( 32 ), an allocated area for a star pair catalog ( 58 ) and a reference
table ( 56 ) stored therein; and c) a processor ( 24 ) coupled to the attitude or angular velocity control system ( 38 ), the star trackers or star sensors (
22 ), and the memory ( 30 ). The processor ( 24 ) populates the star pair catalog ( 58 ), using the method described herein. The processor ( 24 ) then
periodically determines the vehicle inertial attitude or angular velocity or sensor alignment, based, in part, on the star pair catalog ( 58 ) and
reference table ( 56 ). The novel ability of the software to autonomously populate the star pair catalog ( 58 ) allows users to avoid uploading a large
amount of data, and the problems associated with such an upload.
A vehicle ( 12 ) including a control system ( 18 ) is used for controlling vehicle attitude or angular velocity ( 38 ). The processor ( 24 ) is coupled to
a star sensor or tracker ( 22 ) and a memory ( 30 ) that may include a star catalog ( 32 ), and an exclusion list ( 36 ). The exclusion list ( 36 ), a list of
stars to be temporarily excluded from consideration when determining attitude or angular velocity or relative alignment of star sensors or trackers, is
calculated on-board. Such a calculation prevents the necessity for a costly, periodic, ground calculation and upload of such data. By manipulating
the star catalog, or sub-catalogs derived from said catalog, based upon the exclusion list ( 36 ), measurements of such excluded stars are prevented
from corrupting the attitude or angular velocity or alignment estimates formulated on board.
US7228231
Multiple stay-out zones for
ground-based bright object
exclusion
The Boeing Company
A vehicle ( 12 ) including a control system ( 18 ) is used for controlling vehicle attitude or angular velocity ( 38 ). The processor ( 24 ) is coupled to
a star sensor or tracker ( 22 ) and a memory ( 30 ) that may include a star catalog ( 32 ), and an exclusion list ( 36 ). The exclusion list ( 36 ), a list of
stars to be temporarily excluded from consideration when determining attitude or angular velocity or relative alignment of star sensors or trackers, is
calculated on board. Such a calculation prevents the necessity for a costly, periodic, ground calculation and upload of such data. By manipulating
the star catalog, or sub-catalogs derived from said catalog, based upon the exclusion list ( 36 ), measurements of such excluded stars are prevented
from corrupting the attitude or angular velocity or alignment estimates formulated on board. The system uses multiple stayout zones for excluding
stars from the exclusion list. A central exclusion zone excludes all stars while a second or more exclusion zones allow some stars to be used in the
attitude determination.
Mini-Micro Thrusters, LOX/Hydrocarbon Propulsion, and Attitude Control Systems
Avionics and Astrionics – Attitude Determination and Control
USPTO
Patent
Number
US7249531
Title
Assignee
Control moment gyro for
Eads Astrium SAS
attitude control of a spacecraft
Abstract
A control moment gyro comprising a wheel support structure which is mounted on a base by a motor, which is used to rotate the support around a
first axis and bearing a wheel arrangement for driving the wheel spinner at a high speed around a second axis perpendicuIar to the first. The support
structure includes a tubular part which rotates on the base and a flange which is disposed on only one side of the wheel and which is fixed both to
the wheel and the tubular part. The diameter of the wheel and the point at which the wheel is fixed to the flange are proportioned such that the wheel
penetrates the tubular part.
Mini-Micro Thrusters, LOX/Hydrocarbon Propulsion, and Attitude Control Systems
Avionics and Astrionics – Attitude Determination and Control
WIPO
Patent Number
WO03086861A1
Title
WO04101360A2
Momentum Control
System and Method
Honeywell International
Inc.
WO05021379A1
Inertia Wheel for Space
Vehicle
Eads AstriumSas
WO05037646A2
Precision Attitude Control
System for Gimbaled
Thruster
Lockheed Martin
Corporation, Patel,
Moonish, R., Goodzeit,
Neil, E.
WO05045366A1
Method of Controlling the
Attitude of Satellites,
Particularly Agile
Satellites with a Reduced
Number of Gyrodynes
Eads AstriumSas
WO05052513A1
Reaction Wheel Assembly Honeywell International
and Fiber Optic Gyro
Inc.
Virtual Reaction Wheel
ArrayUSing Control
Moment Gyros
Assignee
Honeywell International
Inc.
Abstract
An attitude control system (ACS) for a spacecraft includes an attitude control assembly (ACA) interface (316), a control moment gyro (CMG)
array (314) and a reaction wheel assembly (RWA) control unit (202). The ACA interface (316) converts torque commands received from the RWA
control unit (202) into CMG gimbal rates for the CMG array (314). The ACA interface receives CMG gimbal angles from the CMGs and converts
the CMG gimbals angles into RWA speeds, which are then provided to the RWA control unit (202).
A momentum control system and method is provided that provides attitude control for a vehicle while minimizing the negative effects of the
momentum control system. The momentum control system and method include at least one more reaction wheel than the degrees of freedom under
control. For example, in a vehicle designed to rotate in all three directions, at least four reaction wheels would be provided. The additional reaction
wheel provides an additional control parameter that can be used to minimize the cost of the momentum control system's performance. The cost of
the momentum control system that can be minimized includes the effects of vibration, power consumption, and undesirable changes in rotational
direction, among others.
The invention concerns an inertia wheel comprising at least one inertia mass (1) rotatably mounted through a bearing (2) to at least one ballbearing whereof the stationary part is designed to be fixed on the space vehicle. The inertia mass (1) incorporates at least one rotating ring (5) of
the bearing (2), and said rotating ring (5) is tightly secured to the inertia mass (1). The invention is useful for making inertia wheels, reaction
wheels and satellite gyrodynes.
A system for providing attitude control with respect to a spacecraft is provided. The system includes a reaction wheel control module configured to
control a number of reaction wheel assemblies associated with the spacecraft in order to control attitude, and a maneuver control module
configured to use a number of gimbaled Hall Current thrusters (HCTs) to control the total momentum associated with the spacecraft during an orbit
transfer. The total momentum includes the momentum associated with the reaction wheel assemblies and the angular momentum of the spacecraft.
Using the gimbaled HCTs to control the momentum associated with the reaction wheel assemblies during the orbit transfer results in minimal HCT
gimbal stepping.
The invention relates to a method of controlling the attitude of a satellite (1) comprising two gyrodynes (3, 4) and a third main actuator (2) which
delivers torques at least along the Z axis. The inventive method consists in: fixing the gimbal axes, A1 and A2, of the gyrodynes (3, 4) parallel to
Z; setting a non-zero bias (e) between the angular momentum vectors (Formula I) of the gyrodynes; using the measurements provided by the
sensors on board the satellite to estimate the kinematic and dynamic variables necessary in order to control the attitude of the satellite (1);
calculating set variables in order to realise the objectives assigned to the satellite (1) attitude control system; and using the deviations between the
estimated variables and the set variables to calculate control orders and to send same to the main actuators (2, 3, 4) in order to control the changing
deviations over time, the control orders transmitted to the gyrodynes (3, 4) comprising orders which are used to vary the orientation of the gimbal
axis thereof.
The present invention relates to reaction wheel assembly and fiber optic gyro device. The device includes a reaction wheel assembly having a
reaction wheel assembly housing, a fiber optic gyro coil integrated with the reaction wheel assembly housing, and a fiber optic gyro electronics
integrated with the reaction wheel assembly housing. The fiber optic gyro coil may be wound around the reaction wheel assembly housing. The
gyro coil may also be located within the reaction wheel assembly housing.
Mini-Micro Thrusters, LOX/Hydrocarbon Propulsion, and Attitude Control Systems
Avionics and Astrionics – Attitude Determination and Control
WIPO
Patent Number
WO05075938A1
Title
WO06020314A1
Method and System for
Optimizing Torque in a
Cmg Array
Honeywell International
Inc.
WO06085996A1
Method and System for
Cmg Array Singularity
Avoidance
Honeywell International
Inc.
WO06137843A2
Quantized ControlMoment Gyroscope Array
Honeywell International
Inc.
Bearingless Gyroscope
Assignee
Paul Scherrer Institut
Abstract
A bearingless gyroscope is disclosed, comprising: a housing (12) having an inner wall (16) defining an internal spherical cavity (50) and an outer
wall (18) defining a hermetically sealed circumference; a rotor (10) arranged within and completely encapsulated by said cavity (50); and a gas
(48) arranged within said cavity (50).
A momentum-control system (100) for a spacecraft is disclosed. The momentum-control system (100) comprises an attitude-control system (102).
The attitude-control system (102) receives data concerning a desired spacecraft maneuver and determines a torque command to complete the
desired spacecraft maneuver (302). A momentum actuator control processor (104) coupled to the attitude-control system (102) receives the torque
command. The momentum actuator control processor (1040 calculates a gimbal rate command comprising a range-space gimbal rate required to
produce the torque command (304) and a null-space gimbal rate required to maximize the ability to provide torque in the direction of a current
torque (312). At least four control-moment gyros (106) are coupled to the momentum control actuator control processor. Each of the controlmoment gyros (106) receives and executes the gimbal rate to produce the desired maneuver.
A momentum control system (100) for producing commands to move CMGs (106) within a CMG array (106) is disclosed. The momentum
control system (100) comprising an attitude control system (102) configured to receive data representative of a desired maneuver. In response to
the data representative of the desired maneuver, the attitude control system (102) determines a torque command to complete the desired maneuver
(202). The momentum control system also includes a momentum actuator control processor (104) coupled to receive the torque command from the
attitude control system (102) and operable, in response there to, to calculate a gimbal rate command (212) comprising a range-space gimbal rate
(208) and a null-space gimbal rate (210). In one embodiment, the null-space gimbal rate is calculated based on the projection of a performance
gradient onto the null space of the vector space formed by the control momentum gyroscopes (210).
A momentum actuator (106) for steering a spacecraft is disclosed. The momentum actuator (106) comprises a rotor (110), a gimbal (113) upon
which the rotor is mounted, and a stepper motor (112) coupled to the gimbal (113) and operable to rotate the rotor (110) about the gimbal (113) in a
series of steps. In one embodiment of the present invention the spin rate of the rotor (110) can be varied to provide additional torque.
Mini-Micro Thrusters, LOX/Hydrocarbon Propulsion, and Attitude Control Systems
Avionics and Astrionics – Attitude Determination and Control
FC
Goddard Space Flight
Center
Title
A Shaftless Magnetically
Levitated Multifunctional
Spacecraft Flywheel Storage
System
Jet Propulsion
Laboratory
An attitude control system for
the SIM test bed 3 (STB3)
Jet Propulsion
Laboratory
An Overview of the Formation
and Attitude Control System for
the Terrestrial Planet Finder
Formation Flying Interferometer
Goddard Space Flight
Center
Attitude Control System Design
for the Solar Dynamics
Observatory
Goddard Space Flight
Center
Control Modes of the ST7
Disturbance Reduction System
Flight Validation Experiment
NTRS
Abstract
Presently many types of spacecraft use a Spacecraft Attitude Control System (ACS) with momentum wheels for steering and electrochemical batteries to provide electrical power for
the eclipse period of the spacecraft orbit. Future spacecraft will use Flywheels for combined use in ACS and Energy Storage. This can be done by using multiple wheels and varying
the differential speed for ACS and varying the average speed for energy storage and recovery. Technology in these areas has improved since the 1990s so it is now feasible for
flywheel systems to emerge from the laboratory for spacecraft use. This paper describes a new flywheel system that can be used for both ACS and energy storage. Some of the
possible advantages of a flywheel system are lower total mass and volume, higher efficiency, less thermal impact, improved satellite integration schedule and complexity, simplified
satellite orbital operations, longer life with lower risk, less pointing jitter, and greater capability for high-rate slews. In short, they have the potential to enable new types of missions
and provide lower cost. Two basic types of flywheel configurations are the Flywheel Energy Storage System (FESS) and the Integrated Power and Attitude Control System (IPACS).
The Space Interferometry Mission (SIM) interferometers are required to reject slow (sub-Hz) motions of the spacecraft by feeding the attitude information from the bright-star
tracking interferometers to the dim-star science interferometer, so that the photon-starved dim-star interferometer star tracking system can hold the science object in the field of view
of the interferometer, without losing track of the interferometer fringes. The design is also required to reject translation-induced path-length errors in the science and the guide
interferometers. The SIM Interferometry Test Bed 3 (STB3) is the spaceborne-stellar-interferometer simulator for SIM, built and operating at JPL. Its construction details and
performance are described elsewhere in this conference. The test bed consists of an interferometer system built on a large optical table, and a star simulator built on another large
optical table placed directly across it. The optical tables float on independent, air-filled suspension legs simulating the SIM spacecraft and the distant stars it is to observe. In order to
demonstrate the performance requirements, a novel attitude control system (ACS) has been built and installed on the STB3. In this paper, the details of the design, construction and
performance of the attitude control system are presented. The attitude control system has been used to meet certain SIM requirements. An example of this performance test is also
included.
The Terrestrial Planet Finder formation flying Interferometer (TPF-I) will be a five-spacecraft, precision formation operating near a Sun-Earth Lagrange point. As part of technology
development for TPF-I, a formation and attitude control system (FACS) is being developed that achieves the precision and functionality associated with the TPF-I formation. This
FACS will be demonstrated in a distributed, real-time simulation environment. In this paper we present an overview of the FACS and discuss in detail its constituent formation
estimation, guidance and control architectures and algorithms. Since the FACS is currently being integrated into a high-fidelity simulation environment, component simulations
demonstrating algorithm performance are presented.
The Solar Dynamics Observatory mission, part of the Living With a Star program, will place a geosynchronous satellite in orbit to observe the Sun and relay data to a dedicated
ground station at all times. SDO remains Sun- pointing throughout most of its mission for the instruments to take measurements of the Sun. The SDO attitude control system is a
single-fault tolerant design. Its fully redundant attitude sensor complement includes 16 coarse Sun sensors, a digital Sun sensor, 3 two-axis inertial reference units, 2 star trackers,
and 4 guide telescopes. Attitude actuation is performed using 4 reaction wheels and 8 thrusters, and a single main engine nominally provides velocity-change thrust. The attitude
control software has five nominal control modes-3 wheel-based modes and 2 thruster-based modes. A wheel-based Safehold running in the attitude control electronics box improves
the robustness of the system as a whole. All six modes are designed on the same basic proportional-integral-derivative attitude error structure, with more robust modes setting their
integral gains to zero. The paper details the mode designs and their uses.
The Space Technology 7 experiment will perform an on-orbit system-level validation of two specific Disturbance Reduction System technologies a gravitational reference sensor
employing a free-floating test mass and a set of micro-Newton colloidal thrusters. The ST7 Disturbance Reduction System is designed to maintain the spacecraft s position with
respect to a free-floating test mass to less than 10 nm the square root of Hz over the frequency range of 1 to 30 mHz. This requirement will help ensure that the residual
accelerations on the test masses (beyond gravitational acceleration) will be below the ST7 goal of 300 (1 f 3 mHz(sup 2)) pm s(sup 2) the square root of Hz. This paper presents the
overall design and analysis of the spacecraft drag-free and attitude controllers being designed by NASA s Goddard Space Flight Center. These controllers close the loop between
the GRS and the micro-Newton colloidal thrusters. The ST7 DRS comprises three control systems the attitude control system to maintain a sun-pointing attitude, the drag free
control to center the spacecraft about the test masses, and the test mass suspension control. There are five control modes in the operation of the ST7-DRS, starting from the
attitude-only mode and leading to the challenging science mode. The design and analysis of each of the control modes are presented. An 18-DOF model is developed to capture the
essential dynamics of the ST7-DRS package. It includes all rigid-body dynamics of the spacecraft and two test masses (three translations and three rotations for the spacecraft and
each of the test masses). Actuation and measurement noise and major disturbance sources acting on the spacecraft and test masses are modeled.
Mini-Micro Thrusters, LOX/Hydrocarbon Propulsion, and Attitude Control Systems
Avionics and Astrionics – Attitude Determination and Control
FC
Goddard Space Flight
Center
Title
Controller design for the ST7
disturbance reduction system
Goddard Space Flight
Center; Jet Propulsion
Laboratory
Controller Design for the ST7
Disturbance Reduction System
Goddard Space Flight
Center
Design and Analysis of the ST7
Disturbance Reduction System
(DRS) Spacecraft Controller
Langley Research
Center
Dynamics and Control of
Attitude, Power, and
Momentum for a Spacecraft
Using Flywheels and Control
Moment Gyroscopes
Glenn Research Center
Flywheel Integrated Power and
Attitude Control System
Demonstrated
Glenn Research Center
Integrated Power and Attitude
Control System Demonstrated
With Flywheels G2 and D1
NTRS
Abstract
The Space Technology 7 experiment will perform an on-orbit system-level validation of two specific Disturbance Reduction System technologies a gravitational reference sensor
employing a free-floating test mass and a set of micronewton colloidal thrusters. The Disturbance Reduction System is designed to maintain a spacecraft's position with respect to
the free-floating test mass to less than 10 nm square root of Hz, over the frequency range 10(sup -3) Hz to 10 (sup -2) Hz. This paper presents the design and analysis of the
coupled drag-free and attitude control system that closes the loop between the gravitational reference sensor and the micronewton thrusters while incorporating star tracker data at
low frequencies. The effects of actuation and measurement noise and disturbances on the spacecraft and test masses are evaluated in a seven-degree-of-freedom planar model
incorporating two translational and one rotational degrees of freedom for the spacecraft and two translational degrees of freedom for each test mass.
The Space Technology 7 experiment will perform an on-orbit system-level validation of two specific Disturbance Reduction System technologies a gravitational reference sensor
employing a free-floating test mass and a set of micro-Newton colloidal thrusters. The Disturbance Reduction System is designed to maintain a spacecraft's position with respect to
the free-floating test mass to less than 10 nm square root of Hz, over the frequency range 10(exp -3) Hz to 10(exp -2) Hz. This paper presents the design and analysis of the
coupled drag-free and attitude control system that closes the loop between the gravitational reference sensor and the micro-Newton thrusters while incorporating star tracker data at
low frequencies. The effects of actuation and measurement noise and disturbances on the spacecraft and test masses are evaluated in a seven-degree-of-freedom planar model
incorporating two translational and one rotational degrees of freedom for the spacecraft and two translational degrees of freedom for each test mass.
The Space Technology 7 experiment will perform an on-orbit system-level validation of two specific Disturbance Reduction System technologies a gravitational reference sensor
employing a free-floating test mass and a set of micronewton colloidal thrusters. The Disturbance Reduction System is designed to maintain a spacecraft's position with respect to
the free-floating test mass to less than 10 nm square root of Hz, over the frequency range 10(exp -3) Hz to 10(exp -2) Hz. This paper presents the design and analysis of the
coupled drag-free and attitude control system that closes the loop between the gravitational reference sensor and the micronewton thrusters while incorporating star tracker data at
low frequencies. The effects of actuation and measurement noise and disturbances on the spacecraft and test masses are evaluated in a seven-degree-of-freedom planar model
incorporating two translational and one rotational degrees of freedom for the spacecraft and two translational degrees of freedom for each test mass.
Several laws are designed for simultaneous control of the orientation of an Earth-pointing spacecraft, the energy stored by counter-rotating flywheels, and the angular momentum of
the flywheels and control moment gyroscopes used together as an integrated set of actuators for attitude control. General, nonlinear equations of motion are presented in vectordyadic form, and used to obtain approximate expressions which are then linearized in preparation for design of control laws that include feedback of flywheel kinetic energy error as
a means of compensating for damping exerted by rotor bearings. Two flywheel steering laws are developed such that torque commanded by an attitude control law is achieved while
energy is stored or discharged at the required rate. Using the International Space Station as an example, numerical simulations are performed to demonstrate control about a torque
equilibrium attitude, and illustrate the benefits of kinetic energy error feedback. Control laws for attitude hold are also developed, and used to show the amount of propellant that can
be saved when flywheels assist the CMGs. Nonlinear control laws for large-angle slew maneuvers perform well, but excessive momentum is required to reorient a vehicle like the
International Space Station.
This year, NASA Glenn Research Center's Flywheel Development Team experimentally demonstrated the Integrated Power and Attitude Control System, a full-power, high-speed,
two-flywheel system, simultaneously regulating a power bus and providing a commanded output torque. Operation and power-mode transitions were demonstrated up to 2000 W in
charge and 1100 W in discharge, while the output torque was simultaneously regulated between - 0.8 N-m.
On September 14, 2004, NASA Glenn Research Center's Flywheel Development Team experimentally demonstrated a full-power, high-speed, two-flywheel system, simultaneously
regulating a power bus and providing a commanded output torque. Operation- and power-mode transitions were demonstrated up to 2000 W in charge and 1100 W in discharge,
while the output torque was simultaneously regulated between plus or minus 0.8 N-m. The G2 and D1 flywheels--magnetically levitated carbon-fiber wheels with permanent magnet
motors--were used for the experiment. The units were mounted on an air bearing table in Glenn's High Energy Flywheel Facility. The operational speed range for these tests was
between 20,000 and 60,000 rpm. The bus voltage was regulated at 125 V during charge and discharge, and charge-discharge and discharge-charge transitions were demonstrated
by changing the amount of power that the power supply provided between 300 and 0 W. In a satellite system, this would be the equivalent of changing the amount of energy that the
solar array provides to the spacecraft. In addition to regulating the bus voltage, we simultaneously controlled the net torque produced by the two flywheel modules. Both modules
were mounted on an air table that was restrained by a load cell. The load cell measured the force on the table, and the torque produced by the two flywheels on the table could be
calculated from that measurement. This method was used to measure the torque produced by the modules, yielding net torques from -0.8 to 0.8 N-m. This was the first Glenn
demonstration of the Integrated Power and Attitude Control System (IPACS) at high power levels and speeds.
Mini-Micro Thrusters, LOX/Hydrocarbon Propulsion, and Attitude Control Systems
Avionics and Astrionics – Attitude Determination and Control
NTRS
Abstract
A computer simulation of a flywheel energy storage single axis attitude control system is described. The simulation models hardware which will be experimentally tested in the
future. This hardware consists of two counter rotating flywheels mounted to an air table. The air table allows one axis of rotational motion. An inertia DC bus coordinator is set forth
that allows the two control problems, bus regulation and attitude control, to be separated. Simulation results are presented with a previously derived flywheel bus regulator and a
simple PID attitude controller.
FC
Glenn Research Center
Title
Single Axis Attitude Control and
DC Bus Regulation With Two
Flywheels
Goddard Space Flight
Center; Marshall Space
Flight Center
Solar Sail Control Actuator
Concepts
The thrust produced by a solar sail is a direct function of its attitude. Thus, solar sail thrust vector control is a key technology that must be developed for sailcraft to become a viable
form of deep-space transportation. The solar sail community has been studying various sail Attitude Control System (ACS) actuator designs for near Earth orbit as well as deep
space missions. These actuators include vanes, spreader bars, two-axis gimbals, floating locking gimbals with wheels, and translating masses. This paper documents the various
concepts and performs an assessment at the highest level. This paper will only compare the various ACS actuator concepts as they stand at the publication time. This is not an
endorsement of any particular concept. As concepts mature, the assessments will change.
Goddard Space Flight
Center
The Effects of Thrust
Uncertainty and Attitude
Knowledge Errors on the MMS
Formation Maintenance
Maneuver
Goddard Space Flight
Center
The Microwave Anisotropy
Probe (MAP) Mission
Glenn Research Center
Pulsed Plasma Thruster (PPT)
Technology Earth Observing-1
PPT Operational and Advanced
Components Being Developed
Goddard Space Flight
Center
Integrated Orbit and Attitude
Control for a Nanosatellite with
Power Constraints
The Magnetospheric Multiscale (MMS) Mission will use a formation of four spinning spacecraft to study the Earth s magnetosphere. The science objectives of MMS require a nearregular tetrahedron formation to be maintained with side lengths ranging from ten kilometers to several thousand kilometers at orbit apogee. To reduce spacecraft complexity and
cost, the current mission concept assumes MMS can achieve its formation goals through open-loop orbit control from the ground, rather than in-flight, closed-loop formation control
that has been the subject of recent study. Significant analysis has been performed to provide optimal reference orbit and relative orbit designs. However, the feasibility of achieving
these orbits, and maintaining them for an extended period of time in the presence of real world errors and perturbations has not been investigated. In particular, attitude knowledge
and control errors, which may have a negligible effect on orbit control for conventional missions with spinning spacecraft, can contribute significant errors to the MMS orbits. In this
work, a 6 degree-of-freedom (DOF) simulation is developed and used to analyze the effects of realistic errors on formation maintenance maneuver accuracy. Several realistic
considerations including a finite-burn model, attitude perturbations, and thrust uncertainty are studied. The primary objective is to quantify the effects of realistic attitude and orbit
control, knowledge, and actuator errors on the formation geometry by observing representative maneuver errors of a single spacecraft.
The Microwave Anisotropy Probe mission is designed to produce a map of the cosmic microwave background radiation over the entire celestial sphere by executing a fast spin and
a slow precession of its spin axis about the Sun line to obtain a highly interconnected set of measurements. The spacecraft attitude is sensed and controlled using an inertial
reference unit, two star trackers, a digital sun sensor, twelve coarse sun sensors, three reaction wheel assemblies, and a propulsion system. This paper presents an overview of the
design of the attitude control system to carry out this mission and presents some early flight experience.
In 2002 the pulsed plasma thruster (PPT) mounted on the Earth Observing-1 spacecraft was operated successfully in orbit. The two-axis thruster system is fully incorporated in the
attitude determination and control system and is being used to automatically counteract disturbances in the pitch axis of the spacecraft. The first tests conducted in space
demonstrated the full range of PPT operation, followed by calibration of control torques from the PPT in the attitude control system. Then the spacecraft was placed in PPT control
mode. To date, it has operated for about 30 hr. The PPT successfully controlled pitch momentum during wheel de-spin, solar array acceleration and deceleration during array
rewind, and environmental torques in nominal operating conditions. Images collected with the Advanced Landsat Imager during PPT operation have demonstrated that there was no
degradation in comparison to full momentum wheel control. In addition, other experiments have been performed to interrogate the effects of PPT operation on communication
packages and light reflection from spacecraft surfaces. Future experiments will investigate the possibility of orbit-raising maneuvers, spacecraft roll, and concurrent operation with
the Hyperion imager. Future applications envisioned for pulsed plasma thrusters include longer life, higher precision, multiaxis thruster configurations for three-axis attitude control
systems or high-precision, formationflying systems. Advanced components, such as a "dry" mica-foil capacitor, a wear-resistant spark plug, and a multichannel power processing
unit have been developed under contract with Unison Industries, General Dynamics, and C.U. Aerospace. Over the last year, evaluation tests have been conducted to determine
power processing unit efficiency, atmospheric functionality, vacuum functionality, thruster performance evaluation, thermal performance, and component life.
Small satellites tend to be power-limited, so that actuators used to control the orbit and attitude must compete with each other as well as with other subsystems for limited electrical
power. The Virginia Tech nanosatellite project, HokieSat, must use its limited power resources to operate pulsed-plasma thrusters for orbit control and magnetic torque coils for
attitude control, while also providing power to a GPS receiver, a crosslink transceiver, and other subsystems. The orbit and attitude control strategies were developed independently.
The attitude control system is based on an application of LQR to an averaged system of equations, whereas the orbit control is based on orbit element feedback. In this paper we
describe the strategy for integrating these two control systems and present simulation results to verify the strategy.
Mini-Micro Thrusters, LOX/Hydrocarbon Propulsion, and Attitude Control Systems
Avionics and Astrionics – Guidance Navigation and Control
Company Name
Microcosm, Inc.
Title
Generic Adaptive Approaches for Orbit and Attitude
Determination on Earth Pointing Spacecraft
DOD SBIR Phase I
Quad Chart
Many critical space assets include imaging payloads that keep vigilant watch over the Earth’s surface. The navigation (NAV) and attitude
determination and control (ADCS) approaches traditionally used to control these assets rely on the use of GPS (Global Positioning System) and
a variety of ADCS sensors as primary inputs for control. However, GPS and ADCS sensors are at some risk of loss, outage, or degradation.
Microcosm, with partner HRP Systems, proposes to develop innovative NAV and ADCS approaches, derived from Plug and Play (PnP) avionics
and software development concepts, to provide primary and back-up NAV/ADCS operations for these critical Earth pointing space assets. The
team will analyze discrete NAV and ADCS mode changes based on sensor performance and availability, as well as the development of a flexible
Kalman Filter that seamlessly transitions as the complement and/or quality of inputs changes. These frameworks allow for a broad array of
traditional NAV and ADCS sensors, as well as the integration of “synthetic sensor inputs” which could come from specialized payloads,
communication devices, or ground interaction. The team will draw from their involvement in the PnP ADCS/NAV activities of AFRL’s Responsive
Space Testbed (RST) and PnPSat to develop the proposed capabilities.
Mini-Micro Thrusters, LOX/Hydrocarbon Propulsion, and Attitude Control Systems
Avionics and Astrionics – Guidance Navigation and Control
USPTO
Patent Number
Title
Assignee
Abstract
US6859170
Extended kalman filter for
autonomous satellite navigation
system
The Johns Hopkins
University
An autonomous navigation system for an orbital platform incorporating a global positioning system based navigation device optimized for
low-Earth orbit and medium-Earth orbit applications including a 12 channel, GPS tracking application-specific integrated circuit (15)
operating in concert with a computer system (90) implementing an extended Kalman filter and orbit propagator which autonomously generates
estimates of position, velocity and time to enable planning, prediction and execution of event-based commanding of mission operations.
Mini-Micro Thrusters, LOX/Hydrocarbon Propulsion, and Attitude Control Systems
Avionics and Astrionics – Guidance Navigation and Control
NTRS
Abstract
The NASA New Millennium Program Space Technology 7 (ST7) project will validate technology for precision spacecraft control. The Disturbance Reduction System (DRS) will
be part of the European Space Agency's LISA Pathfinder project. The DRS will control the position of the spacecraft relative to a reference to an accuracy of one nanometer
over time scales of several thousand seconds. To perform the control, the spacecraft will use a new colloid thruster technology. The thrusters will operate over the range of 5 to
30 micro-Newtons with precision of 0.1 micro- Newton. The thrust will be generated by using a high electric field to extract charged droplets of a conducting colloid fluid and
accelerating them with a precisely adjustable voltage. The control reference will be provided by the European LISA Technology Package, which will include two nearly
freefloating test masses. The test mass positions and orientations will be measured using a capacitance bridge. The test mass position and attitude will be adjustable using
electrostatically applied forces and torques. The DRS will control the spacecraft position with respect to one test mass while minimizing disturbances on the second test mass.
The dynamic control system will cover eighteen degrees of freedom six for each of the test masses and six for the spacecraft. After launch in late 2009 to a low Earth orbit, the
LISA Pathfinder spacecraft will be maneuvered to a halo orbit about the Earth-Sun L1 Lagrange point for operations.
FC
JPL
Title
Space Technology 7 Disturbance
Reduction System - precision control
flight Validation
GSFC/JPL
ST7-DRS A Step Towards Drag-free
and High-precision Formation Control
The Space Technology 7 Disturbance Reduction System (ST7-DRS) is an in-space technology demonstration within NASA's New Millennium Program. ST7-DRS is designed to
validate system-level technologies that are required for future gravity missions (including the planned LISA gravitational-wave observatory) and for future formation-flying
interferometer missions (including the planned MAXIM black-hole imager). ST7-DRS is based around a freely-floating test mass contained within a spacecraft structure that will
shield this test mass from all external forces (aside from gravity). The spacecraft position will be continuously controlled, such that the spacecraft, itself, will remain centered
about this test mass, essentially flying in formation with it. Colloidal micro-thrusters will be used to control the spacecraft s position to within a few nanometers, over time scales
of tens to thousands of seconds. In order to detect the residual acceleration noise on the main test mass, a second test mass will be flown alongside the first, within the same
physical spacecraft structure. This test mass will serve as a cross-reference for the first, and will also be used as a reference for the spacecraft's attitude control. The
spacecraft's attitude will be controlled to an accuracy of a few milli-arc-seconds, also utilizing the colloidal micro-thrusters. ST7-DRS will consist of an instrument package
(containing the test masses) and a set of micro-thrusters, which will be attached to the European Space Agency s SMART-2 spacecraft, set to launch in November 2007.
GSFC
The Space Technology-7 Disturbance
Reduction Systems
The Space Technology 7 Disturbance Reduction System (DRS) is an in-space technology demonstration designed to validate technologies that are required for future missions
such as the Laser Interferometer Space Antenna (LISA) and the Micro-Arcsecond X-ray Imaging Mission (MAXIM). The primary sensors that will be used by DRS are two
Gravitational Reference Sensors (GRSs) being developed by Stanford University. DRS will control the spacecraft so that it flies about one of the freely-floating Gravitational
Reference Sensor test masses, keeping it centered within its housing. The other GRS serves as a cross-reference for the first as well as being used as a reference for .the
spacecraft s attitude control. Colloidal MicroNewton Thrusters being developed by the Busek Co. will be used to control the spacecraft's position and attitude using a six degreeof-freedom Dynamic Control System being developed by Goddard Space Flight Center. A laser interferometer being built by the Jet Propulsion Laboratory will be used to help
validate the results of the experiment. The DRS will be launched in 2008 on the European Space Agency (ESA) LISA Pathfinder spacecraft along with a similar ESA experiment,
the LISA Test Package.
JPL
The ST7-DRS Mission Status and
Plans
ST-7 is developing enabling drag-free control technology with low noise micro-thrusters and drag-free control algorithms. The flight computer and amp dynamic control software
are complete, and colloid micro-newton thrusters are beginning ground-based performance testing.
Jet Propulsion
Laboratory
Electromagnetic Formation Flight
(EMFF) for Sparse Aperture Arrays
Traditional methods of actuating spacecraft in sparse aperture arrays use propellant as a reaction mass. For formation flying systems, propellant becomes a critical consumable
which can be quickly exhausted while maintaining relative orientation. Additional problems posed by propellant include optical contamination, plume impingement, thermal
emission, and vibration excitation. For these missions where control of relative degrees of freedom is important, we consider using a system of electromagnets, in concert with
reaction wheels, to replace the consumables. Electromagnetic Formation Flight sparse apertures, powered by solar energy, are designed differently from traditional propulsion
systems, which are based on V. This paper investigates the design of sparse apertures both inside and outside the Earth's gravity field.
Mini-Micro Thrusters, LOX/Hydrocarbon Propulsion, and Attitude Control Systems
Avionics and Astrionics – Telemetry Tracking and Control
U.S. Patent Applications
Patent Number
Title
Assignee
Abstract
US2004026571A1
Apparatus and Methods for InSpace Satellite Operations
None
Apparatus and methods for performing satellite proximity operations such as inspection, recovery and life extension of a target satellite through
operation of a "Satellite Inspection Recovery and Extension" ("SIRE") spacecraft which can be operated in the following modes (teleoperated,
automatic, and autonomous). The SIRE concept further consists of those methods and techniques used to perform certain (on-orbit) operations
including, but not limited to, the inspection, servicing, recovery, and lifetime extension of satellites, spacecraft, space systems, space platforms, and
other vehicles and objects in space, collectively defined as "target satellites". The three basic types of SIRE proximity missions are defined as
"Lifetime Extension", "Recovery", and "Utility". A remote cockpit system is provided to permit human control of the SIRE spacecraft during
proximity operations.
Mini-Micro Thrusters, LOX/Hydrocarbon Propulsion, and Attitude Control Systems
Information – Software Development Environments
Company Name
Emergent Space Technologies,
Inc
Title
System Engineering Process Realization Toolkit
Field Center
GSFC
NASA SBIR Phase I
Quad Chart
NASA faces many systems engineering challenges as it seeks to conduct exploration and science missions concurrently. One
such challenge is implementing a repeatable systems engineering process to missions that vary greatly in scale and complexity.
The same process applied to lunar and Mars exploration missions, for example, do not necessarily scale well to Earth science
missions. Existing systems engineering tools do not adequately address this problem. They tend to be inflexible implementations of
a "one size fits all" approach to process, and provide little or no interoperability with other tools used in the system engineer's
workflow. Emergent Space Technologies, Inc. proposes to develop the Systems Engineering Process Realization Toolkit (SEPRO). SE-PRO will broaden the capabilities for capturing, communicating, and implementing the systems engineering process,
regardless of the size of the mission based on the ability to customize the Capability Maturity Model Integration (CMMI) approach
to process implementation. The innovation lies in the intuitive graphical user interface that allows systems engineers to combine
processes like CMMI, NASA Procedural Requirements (NPR's), and Goddard Procedural Requirements (GPR's), into a single
process that is modifiable by the missions to specify the workflows, team assignments, and output products according to their
needs.
Mini-Micro Thrusters, LOX/Hydrocarbon Propulsion, and Attitude Control Systems
Information – Software Tools for Distributed Analysis and Simulation
USPTO
Patent
Number
US6341749
Title
Assignee
Abstract
Method of simultaneously reducing
inclination and eccentricity for
geostationary orbit transfer
Hughes Electronics Corporation
A method and apparatus is provided for calculating an estimate of thrust vectors and burn times for an optimal two-burn orbit
transfer from an inclined, eccentric initial orbit to a geostationary final orbit. A non-linear root finding algorithm is used to
calculate the thrust vectors and burn times for the optimal two-burn orbit transfer. Thrust vectors and burn times are then
computed for an optimal multi-segment orbit transfer from the initial orbit to the final orbit.
Mini-Micro Thrusters, LOX/Hydrocarbon Propulsion, and Attitude Control Systems
Materials – Ceramics
Company Name
Trex Enterprises Corp.
Title
Volume, Near Net Shape Manufacturing of
Chemical Vapor Composite Divert & Attitude
Control Systems (DACS) Nozzles
DOD SBIR Phase I
Quad Chart
Divert & Attitude Control System nozzles will be fabricated using Trex’s chemical vapor composite (CVC) silicon carbide material. Because the CVC
process results in high purity SiC, the excellent high temperature performance expected of the material is maintained. For example, some high
temperature tests show erosion free performance in CVC SiC at temperatures as high as 4000oF. Furthermore, due to its high hardness and chemical
resistance, CVC SiC is highly resistant to both mechanical and chemical erosion from DACS propellants. The Phase 1 program will produce net shape
CVC SiC DACS nozzles of varying nozzle dimensions, including wall thickness of material and throat diameter. Trex will also develop fiber-reinforced
CVC SiC nozzles with greater fracture toughness and greater survivability under extreme firing temperatures and pressures. Trex will thereby
demonstrate cost-effect manufacturing of robust, erosion-resistant DACS nozzles using its CVC process.
Mini-Micro Thrusters, LOX/Hydrocarbon Propulsion, and Attitude Control Systems
Materials – Ceramics
Company Name
Smaht Ceramics, Inc
Title
ADVANCED, REINFORCED-NZP
COMPOSITES FOR BALLISTIC MISSILE
COMPONENTS & STRUCTURES
DOD SBIR Phase II
Quad Chart
Phase I has demonstrated significantly improved mechanical attributes as compared to unreinforced NZP™, especially at higher temperatures.
The primary goal for Phase II is to successfully complete technological development and optimization as it applies to requirements for Aerojet’s
TDACS and ACS missile components.
Mini-Micro Thrusters, LOX/Hydrocarbon Propulsion, and Attitude Control Systems
Materials – Metallics
Company Name
Plasma Processes, Inc.
Title
INNOVATIVE TUNGSTEN
ALLOYS FOR ADVANCED
PROPULSION SYSTEMS
Field Center
MSFC
NASA SBIR Phase II
Quad Chart
Innovative process for fabricating net shape, tungsten-rhenium-hafnium carbide alloy components are proposed. Development of these
materials will allow the production of components with unique properties and reduce the size, weight, and cost of propulsion systems.
Mini-Micro Thrusters, LOX/Hydrocarbon Propulsion, and Attitude Control Systems
Power and Energy – MHD and Related Conversion
Company Name
Colorado Power Electronics, Inc.
Title
THREE PHASE RESONANT DC POWER
CONVERTER FOR ION THRUSTERS
Field Center
KSC
NASA SBIR Phase II
Quad Chart
The goal of low stored energy has been demonstrated while producing a record-breaking efficiency of 95.5%. This new
converter claims this remarkable accomplishment not at just one full power point but at two distinct points. The converter
was able to deliver its full rated power of 1kW at 2850V and 5000V load voltages without any physical adjusts.
Mini-Micro Thrusters, LOX/Hydrocarbon Propulsion, and Attitude Control Systems
Propulsion
Company Name
Orion Propulsion, Inc.
Title
OXYGEN-METHANE THRUSTER
Field Center
MFSC
NASA SBIR Phase II
Quad Chart
Two main innovations will be developed in the Phase II effort that are fundamentally associated with our gaseous
oxygen/gaseous methane RCS thruster. The first innovation is that of an integrated torch igniter/injector which, provides simple
and reliable ignition for the thruster. The second innovation is the thruster's capability to operate with both gaseous and liquid
propellants.
Mini-Micro Thrusters, LOX/Hydrocarbon Propulsion, and Attitude Control Systems
Propulsion – Beamed Energy
Company Name
Physics, Materials & Applied Math
Research
Title
SPACE PROPULSION THROUGH ULTRASHORT
PULSE LASER ABLATION
Exquadrum, Inc. and University of
Alabama in Huntsville
TECHNOLOGY TO ENABLE RAPID APPLICATION OF
LASER PROPULSION
DOD SBIR Phase II
Quad Chart
This test matrix can identify the thrust that can be obtained in space, using both individual pulses and pulse trains from nanosecond
and/or Joule-level ultra short pulse lasers. This information will identify missions that can be supported by laser-propulsion techniques
either from ground- or space-based sources.
The objective of the proposed research and development effort is to conduct the basic research that is necessary to populate a matrix
of validated design inputs for future flight systems. This technology will help to enable a responsive, low-cost launch system for microsatellites, based on the combination of a high efficiency multi-stage electromagnetic gun and laser beam propulsion.
Mini-Micro Thrusters, LOX/Hydrocarbon Propulsion, and Attitude Control Systems
Propulsion – Beamed Energy
U.S. Patent Applications
Patent Number
Title
Assignee
Abstract
US7118075
System and method for
attitude control and station
keeping
None
A system and method for supplying thrust to a structure, such as a satellite or spacecraft, for the purposes of station keeping and attitude
control of the structure in low-gravity (orbital) and zero-gravity environments. The system includes devices for emitting energy beams and
targets impacted by the energy beams to cause ablation of the targets. The beam-emitting devices and targets are adapted to cooperate and
cause the structure to selectively undergo translational and/or rotational motion in reaction to the motion of material ablated from the targets.
The position, alignment, and/or attitude of the structure can thereby be controlled in a zero or low-gravity environment by selectively
emitting the energy beams at the targets.
Mini-Micro Thrusters, LOX/Hydrocarbon Propulsion, and Attitude Control Systems
Propulsion – Beamed Energy
WIPO
Patent Number
WO05096317A1
Title
Electro-Thermal Nanoparticle
Generator
Assignee
Kansas State
University Research
Foundation
Abstract
A thermo-chemical nanoparticle generator for the controlled production of nanoparticles, and a method of controllably producing
nanoparticles are provided. The combustion of a carbon based propellant under the controlled variation of combustion parameters
(temperature, pressure and stoichiometry) allows for the production of ionized nanoparticles of a defined mass. In addition, with the
assistance of a high voltage generator a combined ionized gas-/solid particle plasma stream can be produced, which can accelerated in an
acceleration tube to high emission velocities allowing for applications of the invention ranging from spacecraft propulsion to plasma
welding.
Mini-Micro Thrusters, LOX/Hydrocarbon Propulsion, and Attitude Control Systems
Propulsion – Beamed Energy
FC
Jet Propulsion Laboratory
Title
Desorption-assisted sun diver
missions
JPL
Inductive energy storage driven
vacuum arc thruster
MSFC
Survey of Beamed Energy Propulsion
Concepts by the MSFC Space
Environmental Effects Team
NTRS
Abstract
Solar-driven sails which can also accelerate by "boil off" of coated materials offer new high-velocity missions. These can take advantage of high temperature characteristics of
the sail by using the large solar flux at perihelion. For the near term use of beamed power, beam illumination at approxkW cm2 in LEO can simulate conditions any solar
grazer mission will experience to within 0.01 A.U. Sublimation (or desorption) thrust from LEO into interplanetary orbit can omit the several-year orbits conventional solar sails
need to reach approx0.1 AU. A second "burn" at perihelion, the highest available orbital velocity in the inner solar system, and thus optimum point for a delta-V, then yields
high velocities of approx50 km s for and gt40 A.U. missions. The mission begins with deployment in Low Earth Orbit by conventional rocket. Then the launch begins, driven by
a microwave beam (and much smaller solar photon thrust) from nearby in orbit. Beam heating makes a "paint" (polymer layer 1) desorp from the sail. Under this enhanced
thrust, in repeated shots at perihelion in steepening elliptical orbits, the sail attains approx15 km s velocity, canceling most of its solar orbital velocity, and so can fall edge-on
toward the sun immediately. (This is far faster than using solar pressure to spiral down, which takes years.) It approaches the sun edge-on, to minimize radiation pressure on it
in the inward fall. At perihelion the spacecraft rotates to face the sun. Under intense sunlight approx20 times Earth insolation, the sail desorps away polymer 2, getting a
approx50 km s boost at its maximum (infall) velocity. It then sails away as a conventional, reflecting solar sail, with the final Aluminum layer revealed. Its final speed is after
leaving the solar potential well is approx10 AU year. Within approx5 years, it sails beyond Pluto, giving high velocity mapping of the outer solar system, the heliopause and
interstellar medium. copyright 2002 American Institute of Physics.
A new type of vacuum arc thruster in combination with an innovative power processing unit (PPU) has been developed that promises to be a high efficiency (approx15), low
mass (approx100 g) propulsion system for micro- and nanosatellites. This thruster accelerates a plasma that consists almost exclusively of ions of the cathode material and
has been operated with a wide variety of cathodes. The streaming velocity of the plasma exhaust varies with cathode material, from a low of 11 km s for Ti up to 30 km s for
Al, with a corresponding range of specific impulse from 1100 s for Ta to 3000 s for Al. Initiation of the arc requires only a few hundred volts due to an innovative "triggerless"
approach in which a conductive layer between the cathode and the anode produces the initial charge carriers needed for plasma production. The initial starting voltage spike
as well as the energy to operate the vacuum arc are generated by a low mass ( and lt300 g) inductive energy storage PPU which is controlled using 5 V level signals. The
thrust-to-power ratio has been estimated to reach up to approximately20 muN W. The vacuum arc thruster was tested at the Jet Propulsion Laboratory using W as cathode
material. Experimental results are within 65 of the estimated values. copyright 2002 American Institute of Physics.
This is a survey paper of work that was performed by the Space Environmental Effects Team at NASA's Marshall Space Flight Center in the area of laser energy propulsion
concepts. Two techniques for laser energy propulsion were investigated. The first was ablative propulsion, which used a pulsed ruby laser impacting on single layer coatings
and films. The purpose of this investigation was to determine the laser power density that produced an optimum coupling coefficient for each type of material tested. A
commercial off-the-shelf multilayer film was also investigated for possible applications in ablative micro-thrusters, and its optimum coupling coefficient was determined. The
second technique measured the purely photonic force provided by a 300W CW YAG laser. In initial studies, the photon force resulting from the momentum of incident photons
was measured directly using a vacuum compatible microbalance and these results were compared to theory. Follow-on work used the same CW laser to excite a stable
optical cavity for the purpose of amplifying the available force from incident photons. copyright 2003 American Institute of Physics
Mini-Micro Thrusters, LOX/Hydrocarbon Propulsion, and Attitude Control Systems
Propulsion – Chemical
NASA SBIR Phase I
Quad Chart
Science and Technology Applications, LLC's (STA) vision for a versatile space propulsion system is a highly throttleable, high
performance, and cost effective Liquid Oxygen/Hydrogen engine utilizing innovative design and manufacturing processes to
simultaneously meet NASA's Space Exploration requirements. To that end, an innovative injector element concept is proposed consisting
of axial flow coaxial injectors with pintle center post. This element concept is expected to provide deep throttle capability while maintaining
performance over the entire range. Cold flow tests will be performed to characterize mixing and atomization distribution.
Company Name
Science and Technology
Applications, LLC
Title
Innovative Deep Throttling, High Performance
Injector Concept
Field Center
MSFC
WASK Engineering, Inc.
Transpiration Cooled Thrust Chamber
Technology
GRC
NASA has determined that it requires extremely durable, high-performance, low cost engines to meet future multi-use in-space, non-toxic,
cryogenic propulsion requirements such as orbit transfer, descent, ascent and pulsing attitude control. Transpiration-cooling technology
has long been considered a candidate for long-life thrust chambers but has never been deployed on a domestic rocket engine. In this
program WASK Engineering, Inc. demonstrates methane transpiration cooling of an oxygen/methane thrust chamber at 260 psia chamber
pressure and a range of mixture ratios up to 3.2 O/F in a 65 lbf engine assembly. Key tasks are the design and fabrication of a
transpiration-cooled chamber spool section that integrates into existing hardware from an on-going USAF program and then hot fire testing
it in the existing test stand. Post-test data analyses are used to anchor and refine thermal and performance algorithms in transpiration
cooling models that then validate, or invalidate, transpiration cooled thrust chambers for this set of requirements.
AeroAstro, Inc.
Resonating Nitrous Oxide Thruster
MSFC
AeroAstro proposes decomposing nitrous oxide (N2O) as an alternative propellant to existing spacecraft propellants. Decomposing N2O
can be used as either a high Isp, hot-gas monopropellant or as a low Isp, cold gas for ACS thrusters. AeroAstro further proposes to use an
innovative technique to achieve N2O decomposition: gasdynamic resonance. Gasdynamic resonance will elevate the N2O to the
activation temperatures required for exothermic decomposition, allowing monopropellant operation without the difficulties of a catalyst.
One of the challenges of long-duration space exploration systems is finding a propellant for microspacecraft that is safe, reliable, robust,
and performs better than current propulsion systems. N2O can replace both hot-gas propellants such as hydrazine and cold-gas ACS
systems such as nitrogen or isobutane. N2O is non-toxic, has a low freezing point (-91oC), and stores as a liquid. N2O is also a byproduct
of the catalysis of ammonia, a main effluent of waste-water recycling systems for long-duration manned space missions. The anticipated
results of this effort are data demonstrating the operating parameters of resonating N2O, and a dual-mode thruster design capable of both
hot-gas and cold-gas operation. Phase II activity will evolve the design of the dual-mode thruster and demonstrate operation over a range
of conditions.
Mini-Micro Thrusters, LOX/Hydrocarbon Propulsion, and Attitude Control Systems
Propulsion – Chemical
Company Name
CFD Research Corporation
Title
HIGH ENERGY, LOW
TEMPERATURE GELLED BIPROPELLANT FORMULATION
FOR LONG-DURATION IN-SPACE
PROPULSION
EERGC Corp.
LOADED GELLED
BIPROPELLANTS FOR OPTIMIZED
PERFORMANCE
WASK Engineering, Inc.
TRANSPIRATION COOLED
THRUST CHAMBER
TECHNOLOGY
Field Center
MSFC
NASA SBIR Phase II
Quad Chart
The use of gelled propellants for deep space planetary missions may enable adoption of high performance (Isp-vac>360 sec) propellant
combinations that do not require power-intensive heating and stirring cycles before firings, and whose handling and safety characteristics are
close to stated goals of "green" propellants. Phase II will culminate in a hot-fire demonstration of a GLP/GMON-30 rocket chamber, to be
performed at AMRDEC facilities.
The focus of this program is the development and validation of formulations, and development methodologies, for optimizing highperformance particulate-loaded bipropellant gels to maximize specific thrust, taking into account not only composition but the effect of particle
size and properties on maximum achievable combustion performance.
GRC
NASA has determined that it requires extremely durable, high-performance, low cost engines to meet future multi-use in-space, non-toxic,
cryogenic propulsion requirements such as orbit transfer, descent, ascent and pulsing attitude control. Transpiration-cooling technology has
long been considered a candidate for long-life thrust chambers, but has never been deployed on a domestic rocket engine.
Mini-Micro Thrusters, LOX/Hydrocarbon Propulsion, and Attitude Control Systems
Propulsion – Chemical
Company Name
Combustion
Propulsion & Ballistic
Techno
Title
Advanced Divert and Attitude Control (DACS) system for
the Multiple (Miniature) Kill Vehicles (MKV)
XCOR Aerospace, Inc.
Advanced Regeneratively Cooled LOX/methane Rocket
Engine with Innovative Injector Design
DOD SBIR Phase I
Quad Chart
One possible method to intercept intercontinental ballistic missiles which deploy countermeasures is to use a "shotgun" approach, in which numerous
small, cheap kill vehicles are deployed simultaneously in the hopes of scoring multiple hits. In order to be effective, such vehicles must be compact,
inexpensive, reliable, and responsive. Miniaturizing a propulsion system is difficult because energetic materials may respond unpredictably at small scales
in an intense combustion environment. We propose an innovative compact propulsion system, which will circumvent these concerns by using a liquid
oxidizer as a coolant to prevent thermal effects from propagating through a pulse-operation-type motor. This system will utilize several self-contained gelled
propellant charges. The oxidizing coolant will be injected into the motor from a bank of pre-pressurized tanks, maximizing the simplicity and
manufacturability of the system; no pumps will be needed, and each propellant charge is modular and isolated from the others. The complete design will
incorporate high energy density, high thrust, quick time response, high reliability, and relatively low expense. The probability of success for this project is
high, since many energetic propellants have been developed by CPBT Corporation in previous SBIR studies. Commercialization of the new propulsion
system has been planned for various applications.
XCOR proposes an innovative LOX/methane engine with regenerative cooling and novel, low cost injector, which provides a relatively high specific impulse
engine with safer, environmentally friendly propellants, and reliable, responsive operations. This engine is designed for use with self pressurizing
propellants, and has no associated pumps or propellant tank pressurization system. XCOR has experience developing rocket motors with many of these
features and it is now time to combine these techniques. We have used a variety of fuel and oxidizer combinations with consistently safe and reliable
startup and shutdown during over 13,000 seconds of engine test and flight operations. XCOR will also test a new configuration of injector. The proposed
engine design has four specific advantages. First, it uses non-toxic LOX/methane propellants with inherently high specific impulse. Second, it has a new
injector design that is inherently low cost to manufacture. Third, it is optimized to operate with self pressurizing propellants, which eliminate the need for
either pumps or a propellant tank pressurization system. Fourth, use of LOX/methane means that the main engines, OMS and RCS can all use the same
propellants, thus simplifying the overall system and reducing weight.
Mini-Micro Thrusters, LOX/Hydrocarbon Propulsion, and Attitude Control Systems
Propulsion – Chemical
Company Name
Orbital Research Inc.
Title
AN ACTIVE THRUST VECTORING (ATV) CONTROL
SYSTEM FOR TACTICAL MISSILE STEERING
SpaceDev
SMALL SHUTTLE-COMPATIBLE PROPULSION
MODULE
DOD SBIR Phase II
Quad Chart
The program will focus on the application of low-power, light-weight, mechanical control actuators to achieve thrust vector control with
limited thrust losses. Phase II will focus on the development of an optimal Active Thrust Vectoring control system for 3D nozzles to fit
inside the 7” diameter missile hardware to ensure compatibility with the current air-to-air missiles.
Phase II will pursue the development of a full-scale prototype MTV Propulsion Module incorporating a hybrid motor that will
demonstrate the functional characteristics and technology of an MTV hybrid motor. Our proposed Propulsion Module is a scalable,
affordable and modular design that utilizes safe, storable propellants.
Mini-Micro Thrusters, LOX/Hydrocarbon Propulsion, and Attitude Control Systems
Propulsion – Chemical
USPTO
Patent Number
Title
Assignee
Abstract
US6949152
Hypergolic azide fuels with hydrogen
peroxide
The Boeing Company
Hypergolic fuel propulsion systems contain a fuel composition and an oxidizer composition. The fuel composition contains an
azide compound that has at least one tertiary nitrogen and at least one azide functional group, as well as a catalyst that contains at
least one transition metal compound. The fuel composition optionally further contains a hydrocarbon fuel. The oxidizer
composition contains hydrogen peroxide. The transition metal catalyst is preferably selected from the group consisting of
compounds of cobalt and manganese. The invention provides a method for propelling a vehicle by providing a fuel mixture that
spontaneously ignites with the oxidizer composition in an engine to provide thrust.
Mini-Micro Thrusters, LOX/Hydrocarbon Propulsion, and Attitude Control Systems
Propulsion – Chemical
FC
MSFC
Title
Demonstration of a Non-Toxic Reaction
Control Engine
Marshall Space
Flight Center
Evaluation of Impinging Stream Vortex
Chamber Concepts for Liquid Rocket
Engine Applications
NTRS
Abstract
Three non-toxic demonstration reaction control engines (RCE) were successfully tested at the Aerojet Sacramento facility under a technology contract sponsored by the National
Aeronautics and Space Administration's (NASA) Marshall Space Flight Center (MSFC). The goals of the NASA MSFC contract (NAS8-01109) were to develop and expand the
technical maturity of a non-toxic, on-orbit auxiliary propulsion system (APS) thruster under the auspices of the Exploration Systems Mission Directorate. The demonstration
engine utilized Liquid Oxygen (LOX) and Ethanol as propellants to produce 870 lbf thrust. The Aerojet RCE's were successfully acceptance tested over a broad range of
operating conditions. Steady state tests evaluated engine response to varying chamber pressures and mixture ratios. In addition to the steady state tests, a variety of pulsing
tests were conducted over a wide range of electrical pulse widths (EPW). Each EPW condition was also tested over a range of percent duty cycles (DC), and bit impulse and
pulsing specific impulse were determined for each of these conditions. Subsequent to acceptance testing at Aerojet, these three engines were delivered to the NASA White
Sands Test Facility (WSTF) in April 2005 for incorporation into a cryogenic Auxiliary Propulsion System Test Bed (APSTB). The APSTB is a test article that will be utilized in an
altitude test cell to simulate anticipated mission applications. The objectives of this APSTB testing included evaluation of engine performance over an extended duty cycle map
of propellant pressure and temperature, as well as engine and system performance at typical mission duty cycles over extended periods of time. This paper provides acceptance
test results and a status of the engine performance as part of the system level testing.
To pursue technology developments for future launch vehicles, NASA Marshall Space Flight Center (MSFC) is examining vortex chamber concepts for liquid rocket engine
applications. Past studies indicated that the vortex chamber schemes potentially have a number of advantages over conventional chamber methods. Due to the nature of the
vortex flow, relatively cooler propellant streams tend to flow along the chamber wall. Hence, the thruster chamber can be operated without the need of any cooling techniques.
This vortex flow also creates strong turbulence, which promotes the propellant mixing process. Consequently, the subject chamber concepts not only offer system simplicity, but
also enhance the combustion performance. Test results have shown that chamber performance is markedly high even at a low chamber length-to-diameter ratio (LD). This
incentive can be translated to a convenience in the thrust chamber packaging. Variations of the vortex chamber concepts have been introduced in the past few decades. These
investigations include an ongoing work at Orbital Technologies Corporation (ORBITEC). By injecting the oxidizer tangentially at the chamber convergence and fuel axially at the
chamber head end, Knuth et al. were able to keep the wall relatively cold. A recent investigation of the low L D vortex chamber concept for gel propellants was conducted by
Michaels. He used both triplet (two oxidizer orifices and one fuel orifice) and unlike impinging schemes to inject propellants tangentially along the chamber wall. Michaels called
the subject injection scheme an Impinging Stream Vortex Chamber (ISVC). His preliminary tests showed that high performance, with an Isp efficiency of 9295, can be obtained.
MSFC and the U. S. Army are jointly investigating an application of the ISVC concept for the cryogenic oxygen hydrocarbon propellant system. This vortex chamber concept is
currently tested with gel propellants at AMCOM at Redstone Arsenal, Alabama. A version of this concept for the liquid oxygen (LOX) hydrocarbon fuel (RP-1) system has been
derived from the one for the gel propellant. An unlike impinging injector was employed to deliver the propellants to the chamber. MSFC is also conducting an alternative injection
scheme, called the chasing injector, associated with this vortex chamber concept. In this injection technique, both propellant jets and their impingement point are in the same
chamber cross-sectional plane. Long duration tests (approximately up to 15 seconds) will be conducted on the ISVC to study the thermal effects. This paper will report the
progress of the subject efforts at NASA Marshall Space Flight Center. Thrust chamber performance and thermal wall compatibility will be evaluated. The chamber pressures,
wall temperatures, and thrust will be measured as appropriate. The test data will be used to validate CFD models, which, in turn, will be used to design the optimum vortex
chambers. Measurements in the previous tests showed that the chamber pressures vary significantly with radius. This is due to the existence of the vortices in the chamber flow
field. Hence, the combustion efficiency may not be easily determined from chamber pressure. For this project, measured thrust data will be collected. The performance
comparison will be in terms of specific impulse efficiencies. In addition to the thrust measurements, several pressure and temperature readings at various locations on the
chamber head faceplate and the chamber wall will be made. The first injector and chamber were designed and fabricated based on the available data and experience gained
during gel propellant system tests by the U.S. Army. The alternate injector for the ISVC was also fabricated. Hot-fire tests of the vortex chamber are about to start and are
expected to complete in February of 2003 at the TS115 facility of MSFC.
Mini-Micro Thrusters, LOX/Hydrocarbon Propulsion, and Attitude Control Systems
Propulsion – Chemical
FC
Marshall Space
Flight Center
Title
Evaluation of Impinging Stream Vortex
Chamber Concepts for Liquid Rocket
Engine Applications
Glenn Research
Center
Microfabricated Liquid Rocket Motors
Marshall Space
Flight Center
NASA's In-Space Propulsion Program
Marshall Space
Flight Center
NASA's In-Space Propulsion
Technology Program A Step Toward
Interstellar Exploration
NTRS
Abstract
NASA Marshall Space Flight Center (MSFC) and the U. S. Army are jointly investigating vortex chamber concepts for cryogenic oxygen hydrocarbon fuel rocket engine
applications. One concept, the Impinging Stream Vortex Chamber Concept (ISVC), has been tested with gel propellants at AMCOM at Redstone Arsenal, Alabama. A version of
this concept for the liquid oxygen (LOX) hydrocarbon fuel (RP-1) propellant system is derived from the one for the gel propellant. An unlike impinging injector is employed to
deliver the propellants to the chamber. MSFC has also designed two alternative injection schemes, called the chasing injectors, associated with this vortex chamber concept. In
these injection techniques, both propellant jets and their impingement point are in the same chamber cross-sectional plane. One injector has a similar orifice size with the
original unlike impinging injector. The second chasing injector has small injection orifices. The team has achieved their objectives of demonstrating the self-cooled chamber wall
benefits of ISVC and of providing the test data for validating computational fluids dynamics (CFD) models. These models, in turn, will be used to design the optimum vortex
chambers in the future.
Under NASA Glenn Research Center sponsorship, MIT has developed the concept of micromachined, bipropellant, liquid rocket engines. This is potentially a breakthrough
technology changing the cost-performance tradeoffs for small propulsion systems, enabling new applications, and redefining the meaning of the term low-cost-access-to-space.
With this NASA support, a liquid-cooled, gaseous propellant version of the thrust chamber and nozzle was designed, built, and tested as a first step. DARPA is currently funding
MIT to demonstrate turbopumps and controls. The work performed herein was the second year of a proposed three-year effort to develop the technology and demonstrate very
high power density, regeneratively cooled, liquid bipropellant rocket engine thrust chamber and nozzles. When combined with the DARPA turbopumps and controls, this work
would enable the design and demonstration of a complete rocket propulsion system. The original MIT-NASA concept used liquid oxygen-ethanol propellants. The military
applications important to DARPA imply that storable liquid propellants are needed. Thus, MIT examined various storable propellant combinations including N2O4 and hydrazine,
and H2O2 and various hydrocarbons. The latter are preferred since they do not have the toxicity of N2O4 and hydrazine. In reflection of the newfound interest in H2O2, it is
once again in production and available commercially. A critical issue for the microrocket engine concept is cooling of the walls in a regenerative design. This is even more
important at microscale than for large engines due to cube-square scaling considerations. Furthermore, the coolant behavior of rocket propellants has not been characterized at
microscale. Therefore, MIT designed and constructed an apparatus expressly for this purpose. The report details measurements of two candidate microrocket fuels, JP-7 and
JP-10.
In order to implement the ambitious science and exploration missions planned over the next several decades, improvements in in-space transportation and propulsion
technologies must be achieved. For robotic exploration and science missions, increased efficiencies of future propulsion systems are critical to reduce overall life-cycle costs.
Future missions will require 2 to 3 times more total change in velocity over their mission lives than the NASA Solar Electric Technology Application Readiness (NSTAR)
demonstration on the Deep Space 1 mission. New opportunities to explore beyond the outer planets and to the stars will require unparalleled technology advancement and
innovation. NASA's In-Space Propulsion (ISP) Program is investing in technologies to meet these needs. The ISP technology portfolio includes many advanced propulsion
systems. From the next generation ion propulsion system operating in the 5-10 kW range, to advanced cryogenic propulsion, substantial advances in spacecraft propulsion
performance are anticipated. Some of the most promising technologies for achieving these goals use the environment of space itself for energy and propulsion and are
generically called, propellantless because they do not require on-board fuel to achieve thrust. Propellantless propulsion technologies include scientific innovations such as solar
and plasma sails, electrodynamic and momentum transfer tethers, and aeroassist and aerocapture. An overview of both propellantless and propellant-based advanced
propulsion technologies, and NASA s plans for advancing them, will be provided.
NASA's In-Space Propulsion Technology Program is investing in technologies that have the potential to revolutionize the robotic exploration of deep space. For robotic
exploration and science missions, increased efficiencies of future propulsion systems are critical to reduce overall life-cycle costs and, in some cases, enable missions
previously considered impossible. Continued reliance on conventional chemical propulsion alone will not enable the robust exploration of deep space. The maximum theoretical
efficiencies have almost been reached and are insufficient to meet needs for many ambitious science missions currently being considered. By developing the capability to
support mid-term robotic mission needs, the program is laying the technological foundation for travel to nearby interstellar space. The In-Space Propulsion Technology Program
s technology portfolio includes many advanced propulsion systems. From the next-generation ion propulsion systems operating in the 5-10 kW range, to solar sail propulsion,
substantial advances in spacecraft propulsion performance are anticipated. Some of the most promising technologies for achieving these goals use the environment of space
itself for energy and propulsion and are generically called "propellantless" because they do not require onboard fuel to achieve thrust. Propellantless propulsion technologies
include scientific innovations, such as solar sails, electrodynamic and momentum transfer tethers, and aerocapture. This paper will provide an overview of those propellantless
and propellant-based advanced propulsion technologies that will most significantly advance our exploration of deep space.
Mini-Micro Thrusters, LOX/Hydrocarbon Propulsion, and Attitude Control Systems
Propulsion – Chemical
FC
Marshall Space
Flight Center
Title
NASA's In-Space Propulsion
Technology Program Overview and
Status
Glenn Research
Center
On-Board Chemical Propulsion
Technology
MSFC
Test Results for a Non-toxic, Dual
Thrust Reaction Control Engine
NTRS
Abstract
NASA's In-Space Propulsion Technology Program is investing in technologies that have the potential to revolutionize the robotic exploration of deep space. For robotic
exploration and science missions, increased efficiencies of future propulsion systems are critical to reduce overall life-cycle costs and, in some cases, enable missions
previously considered impossible. Continued reliance on conventional chemical propulsion alone will not enable the robust exploration of deep space - the maximum theoretical
efficiencies have almost been reached and they are insufficient to meet needs for many ambitious science missions currently being considered. The In-Space Propulsion
Technology Program s technology portfolio includes many advanced propulsion systems. From the next generation ion propulsion system operating in the 5 - 10 kW range, to
advanced cryogenic propulsion, substantial advances in spacecraft propulsion performance are anticipated. Some of the most promising technologies for achieving these goals
use the environment of space itself for energy and propulsion and are generically called, 'propellantless' because they do not require onboard fuel to achieve thrust.
Propellantless propulsion technologies include scientific innovations such as solar sails, electrodynamic and momentum transfer tethers, aeroassist, and aerocapture. This
paper will provide an overview of both propellantless and propellant-based advanced propulsion technologies, and NASA s plans for advancing them as part of the 60M per year
In-Space Propulsion Technology Program.
On-board propulsion functions include orbit insertion, orbit maintenance, constellation maintenance, precision positioning, in-space maneuvering, de-orbiting, vehicle reaction
control, planetary retro, and planetary descent ascent. This paper discusses on-board chemical propulsion technology, including bipropellants, monopropellants, and
micropropulsion. Bipropellant propulsion has focused on maximizing the performance of Earth storable propellants by using high-temperature, oxidation-resistant chamber
materials. The performance of bipropellant systems can be increased further, by operating at elevated chamber pressures and or using higher energy oxidizers. Both options
present system level difficulties for spacecraft, however. Monopropellant research has focused on mixtures composed of an aqueous solution of hydroxl ammonium nitrate
(HAN) and a fuel component. HAN-based monopropellants, unlike hydrazine, do not present a vapor hazard and do not require extraordinary procedures for storage, handling,
and disposal. HAN-based monopropellants generically have higher densities and lower freezing points than the state-of-art hydrazine and can higher performance, depending
on the formulation. High-performance HAN-based monopropellants, however, have aggressive, high-temperature combustion environments and require advances in catalyst
materials or suitable non-catalytic ignition options. The objective of the micropropulsion technology area is to develop low-cost, high-utility propulsion systems for the range of
miniature spacecraft and precision propulsion applications.
A non-toxic, dual thrust reaction control engine (RCE) was successfully tested over a broad range of operating conditions at the Aerojet Sacramento facility. The RCE utilized
LOX Ethanol propellants and was tested in steady state and pulsing modes at 25-lbf thrust (vernier) and at 870-lbf thrust (primary). Steady state vernier tests vaned chamber
pressure (Pc) from 0.78 to 5.96 psia, and mixture ratio (MR) from 0.73 to 1.82, while primary steady state tests vaned Pc from 103 to 179 psia and MR from 1.33 to 1.76. Pulsing
tests explored EPW from 0.080 to 10 seconds and DC from 5 to 50 percent at both thrust levels. Vernier testing accumulated a total of 6,670 seconds of firing time, and 7,215
pulses, and primary testing accumulated a total of 2,060 seconds of firing time and 3,646 pulses.
Mini-Micro Thrusters, LOX/Hydrocarbon Propulsion, and Attitude Control Systems
Propulsion – Electromagnetic Thrusters
Company Name
Photonic Associates, LLC
Title
Feasibility of a 5mN Laser-driven Minithruster
Field Center
GRC
Phoenix Nuclear Labs
Non-ambipolar Electron Source
GRC
NASA SBIR Phase I
Quad Chart
We have developed a next-generation thruster under a Phase II SBIR which we believe can meet NASA requirements after some
modifications and improvements. It is the first practical example of chemically-augmented electric propulsion, using efficient laser diodes
focused on an exothermic fuel tape to make a jet. We own the patents on this technology. Advantages of our thruster technology are large
thrust/power ratio (up to 1.35mN/W) and thrust density, small minimum impulse bit (10nN-s) and instantaneous thrust control. Other
advantages are absence of magnetic fields, high voltage, toxic chemicals, fuel and/or oxidizer storage tanks, heaters or valves, and the fact
that the source of concentrated energy is physically removed from the thrust converter, so that only the fuel, not some engine component,
wears or ablates during operation. Engine lifetime will be limited only by the amount of fuel onboard, not by the 200k year lifetime of the
diode lasers which generate the ms pulses. Problems which we want to address in this Phase I effort are inadequate ablative layer
thickness control which has led to excessive rms thrust noise, a footprint which is larger than we would like and plume contaminants
generated by carbon doping used for laser absorption in the present fuel tape. Our goal is 1% rms thrust noise and an order-of-magnitude
reduction in contaminants deposited by the thruster plume.
A device to produce electron beams from magnetized plasma created with rf fields combined with electron extraction by electron sheaths is
proposed. The source can provide electrons for neutralizing positive ion beams emerging from ion thrusters or as a generic electron
source. With hollow cathode sources currently employed to provide neutralizing electrons, operation is limited in time and/or current density
by cathode deterioration. RF electron sources provide an alternative approach that does not consume electrode material. The current from
this Non-ambipolar Electron Source (NES) exceeds the current normally extracted from conventional rf plasma sources by a factor of
(mi/me)1/2 where mi and me are the ion and electron mass. Ions are lost to a negatively biased conducting cylinder with area Ai chosen to
be Ai ≥ (mi/me)1/2 *Ae where Ae is the electron extraction area. Slots in the conducting cylinder allow the cylinder to serve as a Faraday
shield to reduce capacitive coupling from the antenna to the plasma. Proposed phase 1 design improvements should result in electron
currents comparable to hollow cathode sources with lower neutral gas flow in the inductive discharge phase and higher currents with
helicon operation. Phase 2 will develop prototype sources suitable for spacecraft testing.
Mini-Micro Thrusters, LOX/Hydrocarbon Propulsion, and Attitude Control Systems
Propulsion – Electromagnetic Thrusters
NASA SBIR Phase II
Quad Chart
A power supply concept capable of operation over 25:1 and 64:1 impedance ranges at full power has been successfully
demonstrated in the Phase I effort at efficiencies of 96% and above. This is because one power supply design could meet
the needs of many different single-throttle-point devices without the need to re-design and re-qualify hardware.
Company Name
Colorado Power Electronics, Inc.
Title
WIDE OUTPUT RANGE POWER
PROCESSING UNIT FOR ELECTRIC
PROPULSION
Field Center
GRC
Colorado Power Electronics Inc.
HIGH EFFICIENCY THREE PHASE
RESONANT CONVERSION FOR
STANDARDIZED ARCHITECTURE POWER
SYSTEM APPLICATIONS
MSFC
A low-cost, standardized-architecture power system is proposed for NASA electric propulsion applications. Three
approaches are combined to develop a system that will meet current and future NASA needs and exceed currently
available power processor unit (PPU) performance in terms of electrical efficiency, specific mass, and cost.
Micromechatronics, Inc.
PULSED PLASMA THRUSTER PIEZOIGNITER FOR SMALL SATELLITE
GRC
The proposal addressed the development of a novel discharge initiation (DI) system for application in small satellites
(<40kg). The novel system, IGNIT-SONER, uses a high frequency (~60kHz) multispark rather than a single DC-capacitive
discharge to achieve ignition conditions. This allows optimization of power consumption as well as increasing the reliability
of the system.
Reisz Engineers and University of
Michigan
EFFECT OF AMBIPOLAR POTENTIAL ON
THE PROPULSIVE PERFORMANCE OF THE
GDM PLASMA THRUSTER
MSFC
The Gasdynamic Mirror (GDM) thruster is an electric propulsion device, without electrodes, that will magnetically confine
plasma with such density and temperature as to make the ion-ion collision mean free path much shorter than its length, but
the macroscopic time scale is smaller than the characteristic time for electron-ion equilibrium.
Xintek Inc. (formerly Applied
Nanotechnologies, Inc.)
CARBON NANOTUBE BASED ELECTRIC
PROPULSION THRUSTER WITH LOW
POWER CONSUMPTION
ARC
The CNT based FEEP ion source will be developed innovative in several aspects: integration of CNTs into the ion emission
anode, buildup of the edge anode structure by the combination of the metal tip emitter operation of different thrust unit.
The power consumption of ion thrust is expected to be reduced by a factor of 5 and more.
Mini-Micro Thrusters, LOX/Hydrocarbon Propulsion, and Attitude Control Systems
Propulsion – Electromagnetic Thrusters
Company Name
Busek Co., Inc.
Title
ESPA On-orbit Maneuvering System (OMS)
Electrodynamic Applications, Inc.
Helicon Hall Thruster
DOD SBIR Phase I
Quad Chart
Busek Co. Inc. proposes to develop a low-cost, multi-mode propulsion module that can be added to an EELV Secondary Payload Adapter
(ESPA), turning it into a self-propelled free flying spacecraft (S/C). This Orbit Maneuvering System (OMS) is enabled by the availability of
Busek’s high Isp electric propulsion technology. The baseline system includes four BHT-200 Hall Effect Thrusters (HETs) currently flying on
TacSat2, a COTS hydrazine resistojet, and eight flight qualified hydrazine monopropellant thrusters for attitude control. This 800 W propulsion
system can deliver deltaV >1000 m/s to a 500 kg S/C. Most or all of the deltaV is provided by the HET. Responsive, high thrust deltaV<300 m/s
is provided by the resistojet and ACS thrusters. The proposed OMS is modular and scalable. A 1200 W version using a single BHT-1000
thruster could deliver deltaV >1000 m/s to a 1200 kg system. In Phase I, the propulsion system and OMS architecture will be defined and the
selected
system will be developed into a preliminary design. To assist, Busek will subcontract CSA Engineering, which builds the ESPA ring, and include
MicroSat Systems Inc. in its integrated product team. In Phase II, MSI will be an explicit subcontractor to Busek.
The Hall thruster has great potential in satisfying many of the spacecraft propulsion needs of the United States Air Force for the next several
decades due to its combination of high specific impulse, high thrust efficiency, and high thrust density. The USAF has recently concluded that
bi-modal Hall thrusters (i.e., high thrust-to-power and high specific impulse) are attractive for future missions such as LEO-to-GEO orbital
transfer vehicles (OTVs). To that end, the USAF has recently issued a contract to an Aerojet-led team to develop a 20-kW xenon-propellant Hall
thruster that adheres to IHPRPT Phase III goals. Our proposal seeks to develop a new, two-stage Hall thruster that can achieve IHPRPT Phase
III goals by using a helicon ionization, first-stage coupled to a state-of-the-art Hall accelerator stage. Use of the helicon ionization source, which
is one of the world's most efficient, will allow high Hall thruster efficiency at low discharge voltages by reducing the ion production cost.
Moreover,
given the helicon source affinity for noble gases, it may be possible to develop an efficient Hall thruster that operates with propellants that are
cheaper than xenon, such as argon.
Mini-Micro Thrusters, LOX/Hydrocarbon Propulsion, and Attitude Control Systems
Propulsion – Electromagnetic Thrusters
DOD SBIR Phase II
Quad Chart
Develop an advanced high total impulse (It), low wet mass (Mw) micro pulsed plasma thruster (PPT) system; an advanced MPACS
package with ten times larger It/Mw over SOTA thereby exceeding the IHPRPT goals for electromagnetic propulsion. Key areas for
innovation: clustering of Teflon propellant and electronics mass reduction.
Company Name
Busek Company, Inc.
Title
HIGH PERFORMANCE MICRO PPT
Busek Company, Inc. and Worcester
Polytechnic Institute
COMPACT INDUCED CURRENT HALL
THRUSTER
This Phase II STTR program will investigate a new kind of electric thruster, an inductively-driven Hall thruster. In contrast to conventional
Hall thrusters, this device needs no cathode. We expect it to be more efficient and smaller. It is also quite different from existing inductive
thrusters, because it has a magnetic core and a bias field, has much longer–duration pulses, runs at lower ionization and can be made
much smaller.
Busek Company, Inc.
HIGH CURRENT CATHODE DEVELOPMENT
High power dual mode Hall effect thruster system is an enabling technology for the orbital transport of DoD and commercial space assets.
A dual mode Hall thruster is a propulsion system that is capable of efficient operation in a high thrust to power mode as well as a high
specific impulse mode.
Busek Company, Inc.
EXTENDED LIFETIME LOW POWER HALL
THRUSTERS
The focus of the Phase II effort is to extend the lifetime of the BHT-200 to 3000 hours making it attractive to a broader range of users. In
the Phase II, Busek will continue with design modifications, primarily to the neutral flow distribution and magnetic field profiles that
minimize the impingement of ions onto the insulator surfaces.
Electrodynamic Applications, Inc. and
University of Michigan
THE USE OF BORON NITRIDE FOR
IMPROVED COLD-CATHODE ELECTRON
FIELD EMISSION TECHNOLOGY
This proposal outlines the development of a field emitter cathode utilizing the chemically inert, tough, low work-function material Boron
Nitride (BN) and corresponding testing of a low power Hall thruster to better understand cathode integration. The desirable characteristics
of leading electron emission materials such as molybdenum tips and Carbon Nanotubes (CNT) are well known.
Exquadrum, Inc
DUAL MODE PROPULSION MODULE
TECHNOLOGY FOR SPACE CONTROL
The Dual Mode Propulsion Module Technology for Space Control concept utilizes an innovative approach to rocketry in which a single
thruster can function as a solid rocket motor or as an electric thruster. The resulting system is highly flexible and hence able to perform a
wide variety of space control missions. In this project, a full-scale thruster will be fabricated and experimentally demonstrated.
Starfire Industries LLC
HIGH POWER HALL THRUSTER MODEL
DEVELOPMENT
The significance of this Phase II/III approach is to gain full predictive modeling capability for future high-power Hall thruster configuration
design. “Traditional simulation of high-magnitude crossed field devices employ semi-empirical correction factors to fit experimental data
that greatly limits parametric study and predictive modeling, requiring extensive and expensive ground testing for quantitative data
approximated by non-trivial devices, such as Hall-Effect thrusters.
Mini-Micro Thrusters, LOX/Hydrocarbon Propulsion, and Attitude Control Systems
Propulsion – Electromagnetic Thrusters
WIPO
Patent Number
WO05003557A1
Title
WO05028310A2
Laser Propulsion Thruster
Assignee
Design Net Engineering, Llc
Abstract
A hybrid electric-laser propulsion (HELP) thruster. A propellant has self-regenerative surface morphology. A laser ablates the
propellant to create an ionized exhaust plasma that is non-interfering with a trajectory path of expelled ions. An electromagnetic
field generator generates an electromagnetic field that defines a thrust vector for the exhaust plasma. Multiple HELP thrusters
may be ganged together, and controlled, according to mission criteria.
Spacecraft Thruster
Elwing Llc
A thruster has a chamber (6) defined within a tube (2). The tube has a longitudinal axis which defines an axis (4) of thrust; an
injector (8) injects ionizable gas within the tube, at one end of the chamber. A magnetic field generator with two coils (12, 14)
generates a magnetic field parallel to the axis; the magnetic field has two maxima along the axis (4); an electromagnetic field
generator has a first resonant cavity (16) between the two coils generating a microwave ionizing field at the electron cyclotron
resonance in the chamber (6), between the two maxima of the magnetic field. The electromagnetic field generator has a second
resonant cavity (18) on the other side of the second coil (14). The second resonant cavity (18) generates a ponderomotive
accelerating field accelerating the ionized gas. The thruster ionizes the gas by electron cyclotron resonance, and subsequently
accelerates both electrons and ions by the magnetized ponderomotive force.
Mini-Micro Thrusters, LOX/Hydrocarbon Propulsion, and Attitude Control Systems
Propulsion – Electromagnetic Thrusters
FC
Goddard Space Flight
Center
Title
Integrated Orbit and Attitude
Control for a Nanosatellite with
Power Constraints
Glenn Research Center
Microwave Discharge and
Neutralization Plasma
Production Developed as an
Alternative to Direct-Current
Hollow Cathode Plasmas for
Ion Propulsion
Marshall Space Flight
Center
Plasma Measurements in an
Integrated-System FARAD
Thruster
NTRS
Abstract
HokieSat is a NASA Goddard sponsored spacecraft currently being built by students at Virginia Tech. HokieSat is part of the Ionospheric Observation Nanosatellite Formation
(ION-F) project. The project involves spacecraft built by three schools Virginia Tech (VT), Utah State University (USU), and University of Washington (UW). The three spacecraft
are similar in design and will perform formation flying demonstrations, and make ionospheric measurements. HokieSat uses Pulsed Plasma Thrusters (PPTs) to maintain its
position in the formation. There are two pairs of PPTs on HokieSat their position on HokieSat's hexagonal cross-section is shown. Thrusters T(sub 2) and T(sub 3) provide
translation control, and Thrusters TI and T4 can provide yaw steering. Any thruster can be fired individually. However because they share a capacitor, thrusters T(sub 1) and
T(sub 2) or thrusters T(sub 3) and T(sub 4) cannot be fired simultaneously. Thrusters T(sub 2) T(sub 3) can be fired simultaneously, as well as thrusters T(sub 1) and T(sub 4).
Each thruster provides an impulse-bit of 56 micronN-s and fires at a rate of 1 Hz. For translation control thrusters T2 and T3 are fired together providing an impulse-bit of 112
micronN-s. All four thrusters are positioned slightly above the center of mass, and therefore exert a torque on the spacecraft. Because there are no thrusters in the zenith-nadir
directions, and the communication system requires that the spacecraft remain nadir-pointing, there is no way to thrust in the radial direction. The attitude of HokieSat is controlled
by 3 orthogonal magnetic torque coils. Attitude control is achieved by forcing a current through the torque coils, which interacts with the Earth's magnetic field and creates a
torque. Due to magnetic field interactions between the coils and PPTs, the two actuator systems cannot be used simultaneously, and any attitude or orbit control must be
performed in a piecewise fashion. Power limitations place an additional constraint on the HokieSat control subsystem. When the spacecraft is in eclipse, the power subsystem can
provide only enough power to operate vital spacecraft functions.
Gridded ion propulsion technology holds great promise for enabling future robotic exploration missions with large delta-v requirements. Such missions are made possible by the
high-specific-impulse capability of gridded ion thrusters. However, such missions will require that the engines operate continuously for up to 10 years Thruster lifetime definition,
improvement, and validation are, therefore, very important. Microwave plasma production, in contrast to conventional hollow-cathode-based ion-thruster-discharge plasma
production, can literally eliminate the discharge cathode failure mechanism. Because microwave electron cyclotron resonance (ECR) plasma production is electrodeless, there is
nothing to wear out. In addition, the microwave power source itself has a lifetime measured in hundreds of thousands of hours. High-power ECR plasma production research
conducted at the NASA Glenn Research Center has culminated in the design and testing of a large-area (90- by 40-cm) plasma source. The source has been operated up to 2000
W at a frequency of 2.45 GHz and 2500 W at a microwave frequency of 5.85 GHz. Beam extraction at 2.45 GHz at powers up to 16 kW also has been demonstrated with this
flexible device. Recently, a series of plasma measurements were conducted on the ECR plasma source. Plasma uniformity at the exit plane of the source along the thruster center
was measured to be greater than 90 percent. Transverse uniformity was also greater than 90 percent. Current densities measured at the exit plane suggest that the source can
satisfy the beam current requirements originally laid out in the 2002 NASA announcement calling for the development of a high-power, high-specific-impulse electric propulsion
system. Low measured plasma potentials and the absence of doubly charged xenon, as indicated by emission spectra, suggest a long-life plasma source. The ECR plasma
source effort establishes a credible path for the resolution of a key ion thruster failure mode, cathode life.
Pulsed inductive plasma accelerators are spacecraft propulsion devices in which energy is stored in a capacitor and then discharged through an inductive coil. The device is
electrodeless, inducing a current sheet in a plasma located near the face of the coil. The propellant is accelerated and expelled at a high exhaust velocity (order of 10 km/s)
through the interaction of the plasma current and the induced magnetic field. The Faraday Accelerator with RF-Assisted Discharge (FARAD) thruster[1,2] is a type of pulsed
inductive plasma accelerator in which the plasma is preionized by a mechanism separate from that used to form the current sheet and accelerate the gas. Employing a separate
preionization mechanism allows for the formation of an inductive current sheet at much lower discharge energies and voltages than those used in previous pulsed inductive
accelerators like the Pulsed Inductive Thruster (PIT). A benchtop FARAD thruster was designed following guidelines and similarity performance parameters presented in Refs.
[3,4]. This design is described in detail in Ref. [5]. In this paper, we present the temporally and spatially resolved measurements of the preionized plasma and inductivelyaccelerated current sheet in the FARAD thruster operating with a Vector Inversion Generator (VIG) to preionize the gas and a Bernardes and Merryman circuit topology to provide
inductive acceleration. The acceleration stage operates on the order of 100 J/pulse. Fast-framing photography will be used to produce a time-resolved, global view of the evolving
current sheet. Local diagnostics used include a fast ionization gauge capable of mapping the gas distribution prior to plasma initiation; direct measurement of the induced
magnetic field using B-dot probes, induced azimuthal current measurement using a mini-Rogowski coil, and direct probing of the number density and electron temperature using
triple probes.
Mini-Micro Thrusters, LOX/Hydrocarbon Propulsion, and Attitude Control Systems
Propulsion – Electrostatic Thrusters
Company Name
Busek Co., Inc.
Title
High Current Cathode Development
DOD SBIR Phase I
Quad Chart
High power dual mode electric propulsion systems are an enabling technology for the orbit transport of DoD space assets. In the
Phase I program Busek Co. will design, fabricate and test a high current center mounted cathode. The cathode will be designed
to deliver electron current to a 20 kW thruster capable of both high thrust to power and high Isp operation. The proposed high
current cathode will be scaled from well established hollow cathode technology, but incorporate design innovations that address
challenges with discharge current scaling and lifetime requirements. In Phase I we will fabricate a prototype cathode and
demonstrate high sustained current operation. Additional testing with our existing 20 kW thruster will be performed to
characterize integrated performance. Our subcontractor, Boeing EDD will model the cathode design and operation using a
proprietary model and make first order lifetime predictions. In Phase II, an advanced engineering model prototype will be
fabricated. An extensive test program with our 20 kW thruster will be performed to verify thruster/cathode operation over the dual
mode thruster operating envelope. The Hall thruster system, of which the cathode is an integral component will demonstrate the
capability to exceed Phase III IHPRPT goals for electrostatic propulsion.
Mini-Micro Thrusters, LOX/Hydrocarbon Propulsion, and Attitude Control Systems
Propulsion – Electrostatic Thrusters
Company Name
Busek, Inc.
Title
SIMPLE COLLOID THRUSTER FOR GENERATING
PRECISE IMPULSE THRUST
DOD SBIR Phase II
Quad Chart
Busek will explore the high-thrust electro spray source, characterizing its performance and exploring variants for further improvements in
performance. Concurrent development of the next-generation prototype of propellant storage, valve, feed system and electronics will
occur. Expected delivery of an integrated, EM-like thruster system including propellant storage, valve, and feed system assembly.
Connecticut Analytical Corporation and
Yale & MIT Universities
TWO-DIMENSIONAL MICRO-COLLOID THRUSTER
FABRICATION
Connecticut Analytical Corporation has successfully demonstrated the use of Holey fibers as a 2-D array. Holey fibers offer a selfregulating, passive hydrostatic feed source for colloidal propulsion use, based on a "wick-jet" concept.
Mini-Micro Thrusters, LOX/Hydrocarbon Propulsion, and Attitude Control Systems
Propulsion – Electrostatic Thrusters
U.S. Patent Applications
Patent Number
Title
Assignee
Abstract
US2004010355A1
Multi-function reaction wheel assemblies for
controlling spacecraft attitude
None
A plurality of sensors provide spacecraft attitude signals to reaction wheel assemblies which provide converted signals to a
flight controller that controls the reaction wheel assemblies. The reaction wheel assemblies include a secondary power
supply to power the sensors, and power can be regenerated from the reaction wheels. A microcomputer in the reaction
wheel assemblies can control the reaction wheel assemblies in place of the flight controller.
US2005018198A1
Integrated reaction wheel assembly and fiber
optic gyro
None
The present invention relates to a reaction wheel assembly and fiber optic gyro device. The device includes a reaction
wheel assembly having a reaction wheel assembly housing, a fiber optic gyro coil integrated with the reaction wheel
assembly housing, and a fiber optic gyro electronics integrated with the reaction wheel assembly housing. The fiber optic
gyro coil may be wound around the reaction wheel assembly housing. The gyro coil may also be located within the reaction
wheel assembly housing.
US2006168936A1
Dual mode hybrid electric thruster
The Boeing Company
An invention is provided for a dual mode hybrid electric thruster propulsion system. The dual mode hybrid electric thruster
provides low thrust using ion/plasma exhaust only, and high thrust by mixing a neutral molecular gas with the ion/plasma
exhaust. The dual mode hybrid electric thruster includes a main propellant duct defining a main propellant intake. Coupled
to the main propellant duct is an accelerating element, which includes an ionization chamber and an exhaust output. In
addition, a momentum exchange zone is located at the exhaust output of the accelerating element. The momentum
exchange zone includes a secondary propellant duct defining a secondary propellant intake, which provides a secondary
propellant into the momentum exchange zone. In high-thrust mode, the momentum exchange zone exchanges momentum
between high velocity ions from the accelerating element with atoms of the secondary propellant.
Mini-Micro Thrusters, LOX/Hydrocarbon Propulsion, and Attitude Control Systems
Propulsion – Electrostatic Thrusters
USPTO
Patent Number
Title
Assignee
Abstract
US7294969
Two-Stage Hall Effect Plasma
Accelerator Including Plasma Source
Driven By High-Frequency Discharge
Japan Aerospace
Exploration Agency
Disclosed is a high-frequency discharge plasma generation-based two-stage Hall-effect plasma accelerator, which comprises an
annular acceleration channel having a gas inlet port, a high-frequency wave supply section, an anode, a cathode, a neutralizing
electron generation portion and a magnetic-field generation element, wherein: gas introduced from the gas inlet port into the
annular acceleration channel is ionized by a high-frequency wave supplied from the high-frequency wave supply section, to
generate plasma; a positive ion includes in the generated plasma is accelerated by an acceleration voltage applied between the
anode and cathode, and ejected outside; and an electron included in the generated plasma is restricted in its movement in the axial
direction of the annular acceleration channel by an interaction with a magnetic field. The two-stage Hall-effect plasma accelerator
is designed to control a degree of ion acceleration in accordance with the acceleration voltage serving as an acceleration control
parameter, and control an amount of plasma generation in accordance with the high-frequency wave output serving as a plasmageneration control parameter. The two-stage Hall-effect plasma accelerator of the present invention can control the ion acceleration
and the plasma generation in a highly independent manner.
Mini-Micro Thrusters, LOX/Hydrocarbon Propulsion, and Attitude Control Systems
Propulsion – Electrostatic Thrusters
WIPO
Patent Number
WO02101235A2
Title
Linear Gridless Ion
Thruster
Assignee
The Regents of the University of
Michigan
Abstract
A linear gridless ion thruster (LGIT) is provided to serve as an ion source for spacecraft propulsion or plasma processing. The LGIT is
composed of two stages: (1) an ionization stage composed of a hollow cathode, anode, and cusp magnetic field circuit to ionize the propellant
gas; and (2) an acceleration stage composed of a downstream cathode, upstream anode, and a radial magnetic field circuit to accelerate ions
created in the ionization stage. The LGIT replaces grids used in conventional ion thrusters (Kaufman guns) to accelerate ions with Hallcurrent electrons as in the case with conventional Hall thrusters.
Mini-Micro Thrusters, LOX/Hydrocarbon Propulsion, and Attitude Control Systems
Propulsion – Electrostatic Thrusters
FC
GSFC
Title
The NASA GSFC MEMS Colloidal Thruster
NTRS
Abstract
A number of upcoming missions require different thrust levels on the same spacecraft. A highly scaleable and efficient propulsion system would allow
substantial mass savings. One type of thruster that can throttle from high to low thrust while maintaining a high specific impulse is a Micro-Electro-Mechanical
System (MEMS) colloidal thruster. The NASA GSFC MEMS colloidal thruster has solved the problem of electrical breakdown to permit the integration of the
electrode on top of the emitter by a novel MEMS fabrication technique. Devices have been successfully fabricated and the insulation properties have been
tested to show they can support the required electric field. A computational finite element model was created and used to verify the voltage required to
successfully operate the thruster. An experimental setup has been prepared to test the devices with both optical and Time-Of-Flight diagnostics.
Mini-Micro Thrusters, LOX/Hydrocarbon Propulsion, and Attitude Control Systems
Propulsion – Feed System Components
Company Name
IN Space, LLC
Title
Innovative Swirl Injector for LOX and Hydrocarbon
Propellants
Field Center
GRC
NASA SBIR Phase I
Quad Chart
Gases trapped in the propellant feed lines of space-based rocket engines due to cryogenic propellant boil-off or pressurant
ingestion can result in poor combustion efficiencies, combustion instabilities, or long startup transients. To assist NASA in the use
of the high performing liquid oxygen propellant combinations in space engines, IN Space proposes to investigate the feasibility of
an innovative swirl injector design for liquid oxygen and hydrocarbon propellants to achieve high combustion efficiencies, stable
operation, and short and smooth startup transients despite potential two-phase oxidizer flow. Additionally anticipated benefits of the
injector include low inert mass and low manufacturing costs. IN Space plans to carry out the feasibility assessment of the injector
design by conducting broad parametric test fire evaluations of a notional LOX/hydrocarbon workhorse thruster based on present
NASA needs to assess the effects of several design considerations on the combustion efficiency, static combustion stability, and
startup transient duration performance merits. A preliminary flightweight injector design will also be generated in order to compare
the estimated injector mass with similar injector designs.
Mini-Micro Thrusters, LOX/Hydrocarbon Propulsion, and Attitude Control Systems
Propulsion – Feed System Components
Company Name
W.E. Research LLC and Stanford University
Title
SOLID DIAMOND INSULATORS FOR HALL
THRUSTERS
DOD SBIR Phase II
Quad Chart
This sensor package will consist of the Propulsion Instrumentation Electronics Package (PIE), which will be used to set, control and
sample 30 different sensors. The PIE is smaller, lighter, and more robust than typical central processing units; it uses field programmable
gate arrays to store and pass commands, set biases, and collect data from the sensors before forwarding it to the spacecraft bus.
Mini-Micro Thrusters, LOX/Hydrocarbon Propulsion, and Attitude Control Systems
Propulsion – Feed System Components
U.S. Patent Applications
Patent Number
Title
Assignee
Abstract
US2002148930A1
Maneuver device for artificial satellite
None
The present invention provides a maneuver device for an artificial satellite, which causes small attitude error during maneuver and
which requires a shorter period of setting time for obtaining a target attitude.The maneuver device is provided with: a feed forward
torque instruction signal generator 8 which outputs a feed forward torque instruction signal 11 based on a maneuver plan; a thruster
10 which outputs control torque based on the feed forward torque instruction signal 11; and an attitude control signal calculator 6 to
which an attitude angle and an angular velocity of the artificial satellite as well as a target attitude angle and a target angular velocity
are input and which outputs an attitude control signal 13. The maneuver device is further provided with a disturbance compensating
signal calculator to which the feed forward torque instruction signal 11 and a detected angular velocity signal 16 are input, and which
generates and outputs a disturbance compensating signal 12. The maneuver device is yet further provided with a reaction wheel 7
which generates control torque based on the attitude control signal 13 and the disturbance compensating signal 12.
Mini-Micro Thrusters, LOX/Hydrocarbon Propulsion, and Attitude Control Systems
Propulsion – Feed System Components
WIPO
Patent Number
WO06056710A1
Title
WO07065915A1
Resistojet
Assignee
EadsSpace
TransportationSas
Abstract
A resistojet (1) including a heating region (20) through which a propellant fluid (4) is intended to flow in order to be heated prior to
the discharge thereof, wherein said heating region (20) receives its power supply from an electrical power source (22). According to
the invention, said electrical power source (22) comprises photovoltaic cells (24) mounted to a heat exchanger (10) through which
said propellant fluid (4) is intended to flow before it reaches the heating region (20). The invention is useful in the field of spacecraft
propulsion.
Electronegative Plasma Motor
Ecole Polytechnique
The subject of the invention is a plasma motor comprising the extraction of a stream of positive ions, characterized in that it
comprises: a single ionization stage (1); feed means (6), for feeding said ionization stage with an ionizable electronegative gas;
means for creating an electric field so as to ionize the gas in the ionization stage; and first extraction means, for extracting a stream of
negative ions, and second extraction means, for extracting a stream of positive ions, which are connected to the ionization stage, the
extraction of a stream of positive ions and the extraction of a stream of negative ions with the same amplitude ensuring electrical
neutrality of the motor. According to the invention, the extraction of a stream of positive ions and the extraction of a stream of
negative ions make it possible to ensure neutrality of the motor without having to use a neutralizer, as in the prior art. Application:
propulsion of spacecraft.
Mini-Micro Thrusters, LOX/Hydrocarbon Propulsion, and Attitude Control Systems
Propulsion – Feed System Components
FC
MSFC
Title
Auxiliary Propulsion Activities in Support of NASA's
Exploration Initiative
NTRS
Abstract
The Space Launch Initiative (SLI) procurement mechanism NRA8-30 initiated the Auxiliary Propulsion System Main Propulsion System (APS MPS)
Project in 2001 to address technology gaps and development risks for non-toxic and cryogenic propellants for auxiliary propulsion applications. These
applications include reaction control and orbital maneuvering engines, and storage, pressure control, and transfer technologies associated with onorbit maintenance of cryogens. The project has successfully evolved over several years in response to changing requirements for re-usable launch
vehicle technologies, general launch technology improvements, and, most recently, exploration technologies. Lessons learned based on actual
hardware performance have also played a part in the project evolution to focus now on those technologies deemed specifically relevant to the
Exploration Initiative. Formal relevance reviews held in the spring of 2004 resulted in authority for continuation of the Auxiliary Propulsion Project
through Fiscal Year 2005 (FY05), and provided for a direct reporting path to the Exploration Systems Mission Directorate. The tasks determined to be
relevant under the project were continuation of the development, fabrication, and delivery of three 870 lbf thrust prototype LOX ethanol reaction
control engines the fabrication, assembly, engine integration and testing of the Auxiliary Propulsion Test Bed at White Sands Test Facility and the
completion of FY04 cryogenic fluid management component and subsystem development tasks (mass gauging, pressure control, and liquid
acquisition elements). This paper presents an overview of those tasks, their scope, expectations, and results to-date as carried forward into the
Exploration Initiative.
Mini-Micro Thrusters, LOX/Hydrocarbon Propulsion, and Attitude Control Systems
Propulsion – Fundamental Propulsion Physics
U.S. Patent Applications
Patent Number
Title
US2002007233A1
Low energy method for
changing the inclinations of
orbiting satellitesUSing weak
stability boundaries and a
computer process for
implementing same
US2003197096A1
Method of controlling the
attitude and stabilization of a
satellite in low orbit
None
US2004226279A1
Wick Injection of Colloidal
Fluids for Satellite Propulsion
None
US2006070488A1
Propellantless propulsion
engine
None
US2007295164A1
Centrifugal mass drive
None
Assignee
Galaxy Development,
LLC
Abstract
When a satellite is orbiting the earth in an elliptic orbit, it has a certain inclination with respect to the earth's equator. The usual way to change
the inclination is perform a maneuver by firing the rocket engines at the periapsis of the ellipse. This then forces the satellite into the desired
inclination. There is a substantially more fuel efficient way to change the inclination. This is done by an indirect route by first doing a
maneuver to bring the satellite to the moon on a BCT (Ballistic Capture Transfer). At the moon, the satellite is in the so called fuzzy
boundary or weak stability boundary. A negligibly small maneuver can then bring it back to the earth on a reverse BCT to the desired earth
inclination. Another maneuver puts it into the new ellipse at the earth. In the case of satellites launched from Vandenberg AFB into LEO in a
circular orbit of an altitude of 700 km with an inclination of 340, approximately 6 km/s is required to change the inclination to 900. The
previous flight time associated with this method was approximately 170 days. A modification of this method also achieves a significant
savings and unexpected benefits in energy as measured by Delta-V, where the flight time is also substantially reduced to 88 or even 6 days.
For controlling the attitude of a satellite placed on a low earth orbit, components of a vector Bm of the earth's magnetic field along three
measurement axes of a frame of reference bound with the satellite (typically by means of a three-axis magnetometer)are measured. The
orientation of the earth's magnetic field in the frame of reference is computed and a derivative Bm of the vector is also computed. Magnetocouplers carried by the satellite are energized to create a torque for spinning the satellite at an angular frequency ωc about a determined spin
axis of the satellite, where ωc is greater than an orbital angular frequency 2ω0 of the satellite.
Propellant liquid is supplied to a Colloidal Thruster for Micro Satellite vehicles in Space by capillarity induced flow through a wick element
comprising a permeable porous aggregate of fibers or particles of material that is wetted by the propellant liquid. An intense electric field at
the tip of the wick element dispersed the arriving liquid into a fine spray of charged droplets. Electrodes having appropriate design, location
and potentials accelerate the charge droplets to high velocity, thereby providing reactive thrust to the vehicle. In this method of propellant
liquid introduction the flow rate and exhaust velocity, and therefore the thrust level, are determined by the applied potential difference,
thereby eliminating the need for pumps or pressurized gas and flow controllers to provide the desired flowrate for the propellant liquid.
An inertial thrust engine ( 10 ) includes a housing ( 12 ), a cover plate ( 14 ), an eccentrically disposed elliptical chamber ( 16 ), and a
weighted rotor ( 18 ). The rotor ( 18 ) has a plurality of orbital channels ( 20 a- 20 d) on one face of the rotor. The rotor ( 18 ) is keyed for
rotation on a central shaft ( 22 ). A rotary weight ( 24 ) couples with the weighted rotor ( 18 ) and is kept in an eccentrically disposed position
on the rotor ( 18 ) by the chamber ( 16 ). The rotor ( 18 ) conveys rotary energy to the weight ( 24 ) loaded on the rotor ( 18 ) to gyrate
together about a central axis ( 28 ), producing an unbalanced centrifugal force directed as a propellantless propulsion force ( 32 ).
A propellantless propulsion device comprising a rotary platform to carry and convey rotary energy to a plurality of weights in orbit about a
center of revolution. The weights are arranged in such a manner as to provide a continuous distribution of mass on one side of the rotary drive
during a cycle of revolution. The continuous distribution of mass generates a continuous output of unbalanced centrifugal force components
in one direction. A device for reducing the weights' radius of gyration for a portion of the total time in orbit about the center of revolution; the
device may include the counter-rotation of the weights for that segment of travel in the orbital trajectory. The reduction in radius of gyration
minimizes the magnitude of the centrifugal force components produced in the direction opposing the desired direction of propulsion. The
variations in the trajectory of the weights' orbit generate the unbalanced centrifugal force components that generate a propellantless
propulsion force in one direction.
Mini-Micro Thrusters, LOX/Hydrocarbon Propulsion, and Attitude Control Systems
Propulsion – Fundamental Propulsion Physics
FC
Glenn Research Center
Title
Prospects for Breakthrough Propulsion From Physics
NTRS
Abstract
"Space drives", "Warp drives", and "Wormholes" these concepts may sound like science fiction, but they are being written about in
reputable journals. To assess the implications of these emerging prospects for future spaceflight, NASA supported the Breakthrough
Propulsion Physics Project from 1996 through 2002. This Project has three grand challenges (1) Discover propulsion that eliminates the
need for propellant (2) Discover methods to achieve hyper-fast travel and (3) Discover breakthrough methods to power spacecraft.
Because these challenges are presumably far from fruition, and perhaps even impossible, a special emphasis is placed on selecting
incremental and affordable research that addresses the critical issues behind these challenges. Of 16 incremental research tasks
completed by the project and from other sponsors, about a third were found not to be viable, a quarter have clear opportunities for
sequels, and the rest remain unresolved.
Mini-Micro Thrusters, LOX/Hydrocarbon Propulsion, and Attitude Control Systems
Propulsion – Fundamental Propulsion Physics
WIPO
Patent Number
WO05035363A1
Title
WO07096539A1
Spacecraft and Method for Operating
theSpacecraft
WO07099201A1
Electric Sail for Producing Spacecraft
Propulsion
A Thruster for Propelling and Directing
a Vehicle Without Interacting with
Environment and Method for Making the
Same
Assignee
Payette, Raymond
Universite Pierre Et
Marie Curie (Paris 6),
Palais De La
Decouverte, Centre
National De La
Recherche
Scientifique
Janhunen Pekka
Abstract
A thruster for propelling and directing a vehicle without interacting with its environment without using propellant and particularly
adapted for use in space, comprising rotating means having a pair of a first and a second axes of rotation, connectable to the vehicle
such that each of the first and second axes of rotation extends opposite to each other with respect to a center of mass of the vehicle.
The thruster is also provided with actuator means for actuating a first and a second rotational movement of the rotating means
respectively around each of the first and second axis of rotation. The first rotational movement is actuated in a clockwise direction
thereby generating a first reacting torque in a counter-clockwise direction that causes a pivotal movement of the vehicle around the
first axis of rotation in the counter-clockwise direction. The second rotational movement is actuated in a counter-clockwise direction
thereby generating a second reacting torque in a clockwise direction that causes a pivotal movement of the vehicle around the second
axis of rotation in the clockwise direction. The thruster is also provided with a control mechanism to control and coordinate the
actuator means to impart propulsion and direction to the vehicle. In a further embodiment, the thruster is enclosed. In a further
embodiment, the thruster allows to spin the vehicle. A method for propelling and directing a vehicle without interacting with its
environment is also disclosed.
The present invention relates to a spacecraft (P) comprising a propulsion system allowing a force of variable intensity and variable
orientation to be exerted on the spacecraft, a control system designed to control the propulsion system in terms of intensity and in
terms of orientation so as to collect the spacecraft onto a target path around a planet, using a force which has at least one component
(f<SUP>x</SUP>, f<SUP>y</SUP>, f<SUP>z</SUP>), in the rotating frame of reference associated with the target, which is
substantially dependent linearly on the corresponding coordinate (x, y, z) of the spacecraft within this frame of reference.
A spacecraft propulsion system comprises a plurality of wires (102) or other electrically conductive elongated members deployed
from a main body (101) into respective radial directions. An electric potential generator (605) generates an electric potential on board
said main body (101). The electric coupling between the electric potential generator (605) and the elongated members is controlled
(604) so that all or some of the elongated members (102) assume a high positive potential. An auxiliary propulsion system (203)
rotates said main body around a rotational axis (502) that is perpendicular to said radial directions, thus creating a centrifugal
supporting force to the elongated members.
Mini-Micro Thrusters, LOX/Hydrocarbon Propulsion, and Attitude Control Systems
Propulsion – High Energy Propellants
Company Name
Wickman Spacecraft &
Propulsion Co.
Title
Small Launch Vehicles (SLV) Technologies
Digital Solid State
Propulsion LLC
Solid-State Electrically Controllable Rocket Motors for
Safe Attitude Control Systems
DOD SBIR Phase I
Quad Chart
Difficulties in using foreign air bases and airspace is constraining the Air Force from striking high-value, time critical targets. The Small Launch Vehicle
(SLV) and Common Aero Vehicle (CAV) that deploys weapons over the target within minutes after being launched from the United States can solve this.
The SLV can also be used to launch satellites. Wickman Spacecraft & Propulsion Company is proposing a versatile and low cost SLV that will be air
launched from Air Force cargo planes. It uses an innovative solid rocket motor with individually controlled nozzle throats providing real time controllable
thrust and steering with a fixed nozzle. With an expansion-deflection exit cone, it has an optimum expansion ratio at all altitudes. The propellant is an
environmentally friendly, low cost, phase stabilized ammonium nitrate propellant with a delivered specific impulse equal to ammonium perchlorate
propellants. The solid rocket motors are "field" reloadable. Phase I determines the feasibility of this approach while Phase II demonstrates all the enabling
technologies.
The invention of Electrically Controlled Extinguishable Solid Propellants (ECESP) over the last five years opens the door for completely new ways of
controlling solid rocket motors. When the propellant is fitted with electrodes and a current of the required voltage is applied, the propellant ignites and
continues to burn until the voltage is removed. Throttle control and multiple restarts have been demonstrated with complete extinguishment with end
burning grains. Recently we have demonstrated small "no moving parts" solid propellant core and ending motors. Many of the problems associated with
solid propulsion may be overcome as ECESP technology continues to emerge and develop. Not only are ECESPs controllable they would be extremely
safe for shipboard use. We propose a proof of concept scale up of our solid-state microthrusters to demonstrate a re-starting, 20x throttling motor with
~10 pounds-force of thrust. Heat feed back data from grains and elevated pressure (to 500 psi) electrical strand tests with determine whether scale up
beyond ACS thrusters to high burn rate, booster propulsion is feasible.
Mini-Micro Thrusters, LOX/Hydrocarbon Propulsion, and Attitude Control Systems
Propulsion – High Energy Propellants
USPTO
Patent Number
Title
US6543723
Electric orbit raising with Space Systems/Loral,
variable thrust
Inc.
Assignee
Abstract
Systems and methods for launching a satellite into orbit that optimize the mass of the satellite delivered into orbit. A satellite carrying one or more
chemical propulsion devices and one or more electric propulsion devices is launched into a transfer orbit using a launch vehicle. In one embodiment,
a selected chemical propulsion device is fired to raise the orbit of the satellite from the transfer orbit to an intermediate orbit. One or more electric
propulsion device is fired to raise the orbit of the satellite from the intermediate orbit to final geosynchronous orbit and the one or more electric
propulsion device is throttled to produce variable thrust levels so as to operate at all optimum specific impulse level to optimize the mass of the
satellite delivered into orbit. In another embodiment, the electric propulsion device may deliver the satellite into a near-geosynchronous orbit instead
of final geosynchronous orbit. In this case, a selected chemical propulsion device is fired to place the satellite in its final orbit. In yet another
embodiment, the launch vehicle may be used to deliver the satellite into the intermediate starting orbit after which the electric propulsion device is
used to achieve final orbit. In this case, the transfer orbit raising step is omitted.
Mini-Micro Thrusters, LOX/Hydrocarbon Propulsion, and Attitude Control Systems
Propulsion – MHD
Company Name
Alameda Applied Sciences Corporation
Title
MAGNETICALLY ENHANCED VACUUM ARC
THRUSTER
NASA SBIR Phase II
Field Center
Quad Chart
JPL
This thruster is itself a variant on the Vacuum Arc Thruster that has been demonstrated via a recently
concluded Phase I SBIR contracts from NASA. The MVAT approach has shown to improve thruster
performance and control contamination while keeping the system mass low (<300g).
Mini-Micro Thrusters, LOX/Hydrocarbon Propulsion, and Attitude Control Systems
Propulsion – MHD
Company Name
Alameda Applied Sciences Corp.
DOD SBIR Phase II
Title
Quad Chart
VACUUM ARC NANO THRUSTERS FOR
Advance the Vacuum Arc Thruster closer to flight qualification. Has demonstrated low system mass and performance
NANOSATELLITE SPACECRAFT CONSTELLATIONS
necessary to support precision propulsion requirements for constellation based, nano satellite missions.
Mini-Micro Thrusters, LOX/Hydrocarbon Propulsion, and Attitude Control Systems
Propulsion – MHD
FC
Marshall Space Flight Center
Title
Magnetic Nozzle and Plasma Detachment
Experiment
NTRS
Abstract
High power plasma propulsion can move large payloads for orbit transfer (such as the ISS), lunar missions, and beyond with large savings in fuel
consumption owing to the high specific impulse. At high power, lifetime of the thruster becomes an issue. Electrodeless devices with magnetically guided
plasma offer the advantage of long life since magnetic fields confine the plasma radially and keep it from impacting the material surfaces. For decades,
concerns have been raised about the plasma remaining attached to the magnetic field and returning to the vehicle along the closed magnetic field lines.
Recent analysis suggests that this may not be an issue of the magnetic field is properly shaped in the nozzle region and the plasma has sufficient energy
density to stretch the magnetic field downstream. An experiment was performed to test the theory regarding the Magneto-hydrodynamic (MHD) detachment
scenario. Data from this experiment will be presented. The Variable Specific Impulse Magnetoplasma Rocket (VASIMR) being developed by the Ad Astra
Rocket Company uses a magnetic nozzle as described above. The VASIMR is also a leading candidate for exploiting an electric propulsion test platform
being considered for the ISS.
Mini-Micro Thrusters, LOX/Hydrocarbon Propulsion, and Attitude Control Systems
Propulsion – Micro Thrusters
NASA SBIR Phase II
Quad Chart
In Phase I, Ultramet focused on the development of catalysts that could be scaled down to the dimensions required for such
small thrusters. Phase II will build on this technology, and enable the resulting catalysts to be used with other propellants,
such as hydrazine or HAN-based monopropellants.
Company Name
Ultramet
Title
HIGH-PERFORMANCE CATALYSTS FOR
SMALL IMPULSE BIT THRUSTERS, PHASE II
Field Center
GSFC
Busek Co. Inc.
MICRO RESISTOJET FOR SMALL
SATELLITES
GSFC
Micro-resistojets offer an excellent combination of simplicity, performance and wet system mass for small satellites (<100 kg,
<50 watts) requiring mN level propulsion and low to moderate delta V (<500 m/sec).
Busek Co. Inc.
RADIO FREQUENCY MICRO ION THRUSTER
FOR PRECISION PROPULSION
JPL
Busek proposes to continue development of an engineering model radio frequency discharge, gridded micro ion thruster that
produces sub-mN to mN thrust precisely adjustable over a wide dynamic thrust range.
Mide Technology Corporation
CONTROL VALVE FOR MINIATURE XENON
ION THRUSTER
JPL
Development of a Miniature Xenon Ion thruster will enable precision spacecraft positioning and formation maneuvers for
formation flying spacecraft. The current MiXI thruster prototype will provide 0.5 A-2mN thrust at 3000 sec specific impulse
and efficiencies around 50% or better. The MiXI thruster will use Xenon propellant, a noble gas, minimizing spacecraft
contamination.
Mini-Micro Thrusters, LOX/Hydrocarbon Propulsion, and Attitude Control Systems
Propulsion – Micro Thrusters
DOD SBIR Phase II
Quad Chart
In the Phase II program we propose to further develop the basic thruster array element and characterize it using unique methods for
achieving uniform propellant distribution, stable emission, and direct thrust measurement. This expected performance is unmatched by any
other electric propulsion device operating in the milliNewton thrust range, including the main competitor, liquid metal based FEEPs.
Company Name
Busek Company, Inc.
Title
SLIT COLLOID THRUSTERS USING IONIC LIQUID
PROPELLANTS
ET Materials, LLC
HIGH PERFORMANCE ELECTRICALLY CONTROLLED
SOLID SOLUTION PROPELLANT
This research is directed towards the development of a high energy electrically controlled solid propellant for maneuvering thrusters in micro
satellites. Combustion of this propellant will depend on the continuous but low level input of electrical energy. Extinguishment will take place
rapidly upon removal of this electrical energy. The propellant will have a principal oxidizer and a base binder.
Tanner Research Inc.
ADVANCED DEVELOPMENT OF MICRO THRUSTER
ACTUATORS FOR µUAV APPLICATIONS
Tanner's current micro-scale thruster devices integrate with microelectronics for digital control. Two MEMS-based fabrication processes
comprise the high risk/high reward aspects of the program: the miniaturization of the existing micro thruster concept for implementation at
very large array scale; and, the development of mass fabrication/loading techniques to demonstrate manufacturability.
Sienna Technologies, Inc. and Univ. Of
Washington
HIGH PERFORMANCE MICROTHRUSTERS FOR
MICROSATELLITES
The Phase I effort demonstrated viability of developing a new chemical propulsion system on a scale that is suitable for microsatellites. In
Phase II, Sienna Technologies, Inc. and the University of Washington characterized the proposed propellants, designed, fabricated
microthrusters and conducted laboratory level tests in preparation for implementation in microsatellites.
Mini-Micro Thrusters, LOX/Hydrocarbon Propulsion, and Attitude Control Systems
Propulsion – Micro Thrusters
U.S. Patent Applications
Patent Number
Title
Assignee
Abstract
US2002023427A1
Micro-colloid thruster system
None
A micro-colloid thruster system may be fabricated using micro electromechanical system (MEMS) fabrication techniques. A beam of
charged droplets may be extracted from an emitter tip in an emitter array by an extractor electrode and accelerated by an accelerator
electrode to produce thrust. The micro-colloid thruster system may be used as the main propulsion system for microspacecraft and for
precision maneuvers in larger spacecraft.
US2003080255A1
Attitude Control Methods and
Systems for Multiple-Payload
Spacecraft
THE BOEING
COMPANY
Attitude determination and control systems are provided that combine attitude measurements from all spacecraft payloads to determine a
master attitude estimate for a master payload and relative slave attitude estimates for the remaining slave payloads. These estimates are then
used to control the attitudes of spacecraft elements that correct the absolute and relative attitude errors. These systems significantly enhance
attitude accuracy when compared to systems that realize independent payload estimates. determine payload attitudes. These systems also
provide significant processing advantages (e.g., simpler algorithms, reduced data throughput and slower processing rate).
Mini-Micro Thrusters, LOX/Hydrocarbon Propulsion, and Attitude Control Systems
Propulsion – Micro Thrusters
USPTO
Patent
Number
US6494402
Title
Assignee
Abstract
Lateral exhaust
microthruster
The Aerospace Corporation
A microthruster having an inverted exhaust system traps burst diaphragm fragments providing a clean exhaust while an exhaust port provides
increased back pressure for efficient combustion of a propellant charge in a fuel cell. A converging diverging micronozzle provides a predictable
exhaust vector for improved microthrusting well suited for propulsion system on small spacecraft.
US6892525
Micropump-based
microthruster
Honeywell International
Inc.
A thruster for providing thrust for spacecraft positioning, which has a propellant reservoir for storing propellant, a reaction chamber for
discharging a vapor for providing thrust, a pump module comprising one or more micropumps for drawing propellant from the reservoir and for
systematically metering propellant to the reaction chamber in a controlled manner, and a controller for actuating the pump module.
Mini-Micro Thrusters, LOX/Hydrocarbon Propulsion, and Attitude Control Systems
Propulsion – Micro Thrusters
WIPO
Patent Number
WO03007311A1
Title
Description of Methods to Increase Propellant
Throughput in a Micro Pulsed Plasma Thruster
Assignee
W.E. Research Llc
Abstract
Propellant modules for Micro Pulsed Plasma Thrusters, and techniques for bundling propellant modules (30) and for using a
two-stage (36,36') discharge process to increase MicroPPT propellant (18) throughput, and decrease the output voltage
required from the power-processing unit are provided.
Mini-Micro Thrusters, LOX/Hydrocarbon Propulsion, and Attitude Control Systems
Propulsion – Micro Thrusters
FC
JPL
JPL
Title
Experimental and computational
investigation of the performance of a
micro-ion thruster
JPL micro-thrust propulsion activities
GSFC/JPL
Structural Modeling for the Terrestrial
Planet Finder Mission
JPL
Survey of the state-of-the-art in microthrust propulsion
18-Degree-of-Freedom Controller
Design for the ST7 Disturbance
Reduction System
Goddard Space Flight
Center
Jet Propulsion
Laboratory
An overview of MEMS-based
micropropulsion developments at JPL
NTRS
Abstract
A micro-ion thruster assembly with a characteristic diameter of 3-cm has been developed at JPL for testing and optimization of various system parameters.
Formation flying and microspacecraft constellation missions pose new propulsion requirements. Formationflying spacecraft, due to the tight positioning and pointing control
requirements, may need thrust control within 1- 20 uN to an accuracy of 0.1 uN for LISA and ST-7, for example. Future missions may have extended thrust ranges into the sub
- mN range. However, all do require high specific impulses ( and gt500 sec) due to long required thruster firings.
We present the most recent propulsion requirements for the Laser Interferometer Space these requirements. LISA consists of three spacecraft in heliocentric orbits, forming a
triangle with 5x lo6 km sides that are the arms of three Michelson-type interferometers. Reflective proof masses provide the reference surfaces at the end of the interferometer
arms as part of the Gravitational Reference Sensor (GRS) designed to detect gravitational waves. The microthrust propulsion system will be part of the Disturbance Reduction
System (DRS), which is responsible for maintaining each spacecraft position within approximately 10 nm around the proof masses. To provide the necessary sensitivity, the
GRS must not experience spurious accelerations and gt; 10 (exp -15) m/s(exp 2)# Hz (exp -1/2) in the 0.1 mHz to 1 Hz bandwidth, requiring precision formation flying and
drag-free operation of the LISA spacecraft. This leads to the following microthruster performance requirements: a thrust range of 2-30 microN, a thrust resolution and lt; O.1
micro N, and thrust noise and lt;0.1 micro N Hz (exp -1/2) over the LISA measurement bandwidth. The microthruster must provide this performance for 5 years continuously,
contain 10 years worth of propellant, and not disrupt the science measurements. Potential microthruster technologies include Colloid, Field Emission Electric Propulsion
(FEEP), and precision cold gas microthrusters. Each of these technologies is described in detail with focus on the NASA microthruster development of the Busek Colloid
Micro-Newton Thruster (CMNT).
No Abstract Available
This paper presents the overall design and analysis process of the spacecraft controller being developed at NASA's Goddard Space Flight Center to close the loop between
the GRS and the micro-newton colloidal thrusters. The essential dynamics of the ST7-DRS are captured in a simulation including eighteen rigid-body dynamic degrees of
freedom three translations and three rotations for the spacecraft and for each test mass. The ST7 DRS comprises three control systems the attitude control system (ACS) to
maintain a sun-pointing attitude the drag free control (DFC) to center the spacecraft about the test masses and the test mass suspension control. This paper summarizes the
control design and analysis of the ST7-DRS 18-DOF model, and is an extension of previous analyses employing a 7-DOF planar model of ST-7.
Development of MEMS (Microelectromechanical Systems) micropropulsion at the Jet Propulsion Laboratory (JPL) is reviewed. This includes a vaporizing liquid micro-thruster
for microspacecraft attitude control, a micro-ion engine for microspacecraft primary propulsion or large spacecraft fine attitude control, as well as several valve studies,
including a solenoid valve studied in collaboration with Moog Space Products Division, and a piezoelectric micro-valve. The solenoid valve features much faster actuation (as
little as 1.5 ms to open) than commercially available MEMS valves and showed no detectable leak (less than 10(sup -4) sccs GN(sub 2)) even after 1 million cycles. The
solenoid valve weighs 7 gram and is about 1 cm(sup 3). A micro-isolation valve, aimed at sealing propulsion systems at zero leak rates, was able to show burst pressures as
high as 3,000 psi even though entirely machined from silicon and Pyrex. It could be actuated with energies as little as 0.1 mJ. (copyright) 2003 Published by Elsevier Science
Ltd.
Mini-Micro Thrusters, LOX/Hydrocarbon Propulsion, and Attitude Control Systems
Propulsion – Micro Thrusters
FC
Glenn Research
Center
Title
Catalyzed Combustion In MicroPropulsion Devices Project Status
Glenn Research
Center
Catalyzed Combustion of Bipropellants
for Micro-Spacecraft Propulsion
Glenn Research
Center
Catalyzed Ignition of Bipropellants in
Microtubes
NTRS
Abstract
In recent years, there has been a tendency toward shrinking the size of spacecraft. New classes of spacecraft called micro-spacecraft have been defined by their mass,
power, and size ranges. Spacecraft in the range of 20 to 100 kg represent the class most likely to be utilized by most small sat users in the near future. There are also efforts
to develop 10 to 20 kg class spacecraft for use in satellite constellations. More ambitious efforts will be to develop spacecraft less than 10 kg, in which MEMS fabrication
technology is required. These new micro-spacecraft will require new micro-propulsion technology. Although micro-propulsion includes electric propulsion approaches, the
focus of this proposed program is micro-chemical propulsion which requires the development of microcombustors. As combustors are scaled down, the surface to volume ratio
increases. The heat release rate in the combustor scales with volume, while heat loss rate scales with surface area. Consequently, heat loss eventually dominates over heat
release when the combustor size becomes smaller, thereby leading to flame quenching. The limitations imposed on chamber length and diameter has an immediate impact on
the degree of miniaturization of a micro-combustor. Before micro-combustors can be realized, such a difficulty must be overcome. One viable combustion alternative is to take
advantage of surface catalysis. Micro-chemical propulsion for small spacecraft can be used for primary thrust, orbit insertion, trajectory-control, and attitude control. Grouping
micro-propulsion devices in arrays will allow their use for larger thrust applications. By using an array composed of hundreds or thousands of micro-thruster units, a particular
configuration can be arranged to be best suited for a specific application. Moreover, different thruster sizes would provide for a range of thrust levels (from N s to mN s) within
the same array. Several thrusters could be fired simultaneously for thrust levels higher than the basic units, or in a rapid sequence in order to provide gradual but steady low-g
acceleration. These arrays of micro-propulsion systems would offer unprecedented flexibility and redundancy for satellite propulsion and reaction control for launch vehicles. A
high-pressure bi-propellant micro-rocket engine is already being developed using MEMS technology. High pressure turbopumps and valves are to be incorporated onto the
rocket chip . High pressure combustion of methane and O2 in a micro-combustor has been demonstrated without catalysis, but ignition was established with a spark. This
combustor has rectangular dimensions of 1.5 mm by 8 mm (hydraulic diameter 3.9 mm) and a length of 4.5 mm and was operated at 1250 kPa with plans to operate it at 12.7
MPa. These high operating pressures enable the combustion process in these devices, but these pressures are not practical for pressure fed satellite propulsion systems.
Note that the use of these propellants requires an ignition system and that the use of a spark would impose a size limitation to this micro-propulsion device because the spark
unit cannot be shrunk proportionately with the thruster. Results presented in this paper consist of an experimental evaluation of the minimum catalyst temperature for initiating
supporting combustion in sub-millimeter diameter tubes. The tubes are resistively heated and reactive premixed gases are passed through the tubes. Tube temperature and
inlet pressure are monitored for an indication of exothermic reactions and composition changes in the gases.
This paper addresses the need to understand the physics and chemistry involved in propellant combustion processes in micro-scale combustors for propulsion systems on
micro-spacecraft. These spacecraft are planned to have a mass less than 50 kilograms with attitude control estimated to be in the 10 milli-Newton thrust class. These
combustors are anticipated to be manufactured using Micro Electrical Mechanical Systems (MEMS) technology and are expected to have diameters approaching the
quenching diameter of the propellants. Combustors of this size are expected to benefit significantly from surface catalysis processes. Miniature flame tube apparatus is chosen
for this study because microtubes can be easily fabricated from known catalyst materials and their simplicity in geometry can be used in fundamental simulations for validation
purposes. Experimentally, we investigated the role of catalytically active surfaces within 0.4 and 0.8 mm internal diameter microtubes, with special emphases on ignition
processes in fuel rich gaseous hydrogen and gaseous oxygen. Flame thickness and reaction zone thickness calculations predict that the diameters of our test apparatus are
below the quenching diameter of the propellants in sub-atmospheric tests. Temperature and pressure rise in resistively heated platinum and palladium microtubes was used
as an indication of exothermic reactions. Specific data on mass flow versus preheat temperature required to achieve ignition are presented. With a plug flow model, the
experimental conditions were simulated with detailed gas-phase chemistry, thermodynamic properties, and surface kinetics. Computational results generally support the
experimental findings, but suggest an experimental mapping of the exit temperature and composition is needed.
This paper addresses the need to understand the physics and chemistry involved in propellant combustion processes in micro-scale combustors for propulsion systems on
micro-spacecraft. These spacecraft are planned to have a mass less than 50 kilograms with attitude control estimated to be in the 10 milli-Newton thrust class. These
combustors are anticipated to be manufactured using Micro Electrical Mechanical Systems (MEMS) technology and are expected to have diameters approaching the
quenching diameter of the propellants. Combustors of this size are expected to benefit significantly from surface catalysis processes. Miniature flame tube apparatus is chosen
for this study because microtubes can be easily fabricated from known catalyst materials and their simplicity in geometry can be used in fundamental simulations to more
carefully characterize the measured heat transfer and pressure losses for validation purposes. Experimentally, we investigate the role of catalytically active surfaces within 0.4
and 0.8 mm internal diameter micro-tubes, with special emphases on ignition and extinction processes in fuel rich gaseous hydrogen and gaseous oxygen. Flame thickness
and reaction zone thickness calculations predict that the diameters of our test apparatus are below the quenching diameter of the propellants in sub-atmospheric tests.
Temperature and pressure rises in resistively heated platinum and palladium micro-tubes are used as an indication of exothermic reactions. Specific data on mass flow versus
preheat temperature required to achieve ignition are presented.
Mini-Micro Thrusters, LOX/Hydrocarbon Propulsion, and Attitude Control Systems
Propulsion – Micro Thrusters
NTRS
Abstract
Micropropulsion technology is essential to the success of miniaturized spacecraft and can provide ultra-precise propulsion for small spacecraft. NASA Glenn Research Center
has envisioned a micropropulsion concept that utilizes decomposing solid propellants for a valveless, leak-free propulsion system. Among the technical challenges of this
decomposing solid micropropulsion concept is optimization of miniature, rectangular nozzles. A number of flat micronozzles were tested with ambient-temperature nitrogen
and helium gas in a vacuum facility. The thrusters were etched out of silicon and had throat widths on the order of 350 microns and throat depths on the order of 250 microns.
While these were half-sections of thrusters (two would be bonded together before firing), testing provided the performance trend for nozzles of this scale and geometry. Area
ratios from 1 to 25 were tested, with thrust measured using an inverted pendulum thrust stand for nitrogen flows and a torsional thrust stand for helium. In the nitrogen testing,
peak nozzle performance was achieved around area ratio of 5. In the helium series, nozzle performance peaked for the smallest nozzle tested area ratio 1.5. For both gases,
there was a secondary performance peak above area ratio 15. At low chamber pressures ( less than 1.6 atm), nitrogen provided higher nozzle performance than helium. The
performance curve for helium was steeper, however, and it appeared that helium would provide better performance than nitrogen at higher chamber pressures.
We present the most recent propulsion requirements for the Laser Interferometer Space Antenna (LISA) Mission and describe potential microthruster technology that can
meet these requirements. LISA consists of three spacecraft in heliocentric orbits, forming a triangle with 5x l0 (exp 6) km sides that are the arms of three Michelson-type
interferometers. Reflective proof masses provide the reference surfaces at the end of the interferometer arms as part of the Gravitational Reference Sensor (GRS) designed to
detect gravitational waves. The microthrust propulsion system will be part of the Disturbance Reduction System (DRS), which is responsible for maintaining each spacecraft
position within approximate ly 10 nm around the proof masses. To provide the necessary sensitivity, the GRS must not experience spurious accelerations and gt;15 (exp -10)
m/ s(exp 2) in the 0.1 mHz to 1 Hz bandwidth, requiring precision formation flying and drag-free operation of the LISA spacecraft. This leads to the following microthruster
performance requirements: a thrust range of 2-30 Micro N, a thrust resolution and lt; 0.1 Micro N, and thrust noise and lt;0.1 Hz(exp -1/2) over the LISA measurement
bandwidth. The microthruster must provide this performance for 5 years continuously, contain 10 years worth of propellant, and not disrupt the science measurements.
Potential microthruster technologies include Colloid, Field Emission Electric Propulsion (FEEP), and precision cold gas microthrusters. Each of these technologies is described
in detail with focus on the NASA microthruster development of the Busek Colloid Micro-Newton Thruster (CMNT).
For future applications to precision formation flying missions, NASA's New Millennium Program is scheduled to test colloid micro-Newton thrusters (CMNTs) on the ST7
technology demonstration mission. These CMNTs are part of a disturbance reduction system (DRS) on the ESA SMART-2 Spacecraft or LISA Pathfinder. The goal of the ST7
DRS is to demonstrate technologies necessary to meet the nanometer precision positioning control requirements of the LISA mission. In order to achieve these goals, the
CMNTs are required to demonstrate a thrust resolution of less than 0.1 micro-N and a thrust noise of less than 0.1 micro-N square rootHz for thrust levels between 5 and 30
micro-N. Developed by Busek Co. with support from JPL in testing an design, the CMNT has been developed over the last four years into a flight-ready microthrust system.
The development, validation testing, and flight unit production of the CMNTs are described. Development tests and analysis include preliminary wear tests, propellant loading
process verification, flow testing, and performance verification. Validation and flight unit verification includes thermal and structural analysis, life testing, thermal and dynamic
load testing, and performance verification. Final delivery of the units is planned in 2007 with and planned launch and flight demonstration 2009.
FC
Glenn Research
Center
Title
Decomposing Solid Micropropulsion
Nozzle Performance Issues
Goddard Space Flight
Center; Jet
Propulsion Laboratory
Microthrust Propulsion of the LISA
Mission
Jet Propulsion
Laboratory
Microthruster Propulsion for the Space
Technology 7 (ST7) Technology
Demonstration Mission
Jet Propulsion
Laboratory
Performance characterization of the
Vaporizing Liquid Micro Thruster (VLM)
A Vaporizing Liquid Micro-Thruster (VLM) microfabricated thruster was tested on water propellant on a thrust stand and performance data obtained.
Goddard Space Flight
Center
Propulsion Options for the Global
Precipitation Measurement Core
Satellite
This study was conducted to evaluate several propulsion system options for the Global Precipitation Measurement (GPM) core satellite. Orbital simulations showed clear
benefits for the scientific data to be obtained at a constant orbital altitude rather than with a decay reboost approach. An orbital analysis estimated the drag force on the
satellite will be 1 to 12 mN during the five-year mission. Four electric propulsion systems were identified that are able to compensate for these drag forces and maintain a
circular orbit. The four systems were the UK-10 TS and the NASA 8 cm ion engines, and the ESA RMT and RITl0 EVO radio-frequency ion engines. The mass, cost, and
power requirements were examined for these four systems. The systems were also evaluated for the transfer time from the initial orbit of 400 x 650 km altitude orbit to a
circular 400 km orbit. The transfer times were excessive, and as a consequence a dual system concept (with a hydrazine monopropellant system for the orbit transfer and
electric propulsion for drag compensation) was examined. Clear mass benefits were obtained with the dual system, but cost remains an issue because of the larger power
system required for the electric propulsion system. An electrodynamic tether was also evaluated in this trade study.
Mini-Micro Thrusters, LOX/Hydrocarbon Propulsion, and Attitude Control Systems
Propulsion – Micro Thrusters
FC
Marshall Space Flight
Center
Title
Solid State Digital Propulsion "Cluster
Thrusters" For Small Satellites Using
High Performance Electrically
Controlled Extinguishable Solid
Propellants (ECESP)
Jet Propulsion
Laboratory
The Minimum Impulse Thruster
NTRS
Abstract
Electrically controlled extinguishable solid propellants (ESCSP) are capable of multiple ignitions, extinguishments and throttle control by the application of electrical power.
Both core and end burning no moving parts ECESP grains motors to three inches in diameter have now been tested. Ongoing research has led to a newer family of even
higher performance ECESP providing up to 10 higher performance, manufacturing ease, and significantly higher electrical conduction. The high conductivity was not found to
be desirable for larger motors however it is ideal for downward scaling to micro and pico- propulsion applications with a web thickness of less than 0.125 inch diameter. As a
solid solution propellant, this ECESP is molecularly uniform, having no granular structure. Because of this homogeneity and workable viscosity it can be directly cast into thin
layers or vacuum cast into complex geometries. Both coaxial and grain stacks have been demonstrated. Combining individual propellant coaxial grains and or grain stacks
together form three-dimensional arrays yield modular cluster thrusters. Adoption of fabless manufacturing methods and standards from the electronics industry will provide
custom, highly reproducible micro-propulsion arrays and clusters at low costs. These stack and cluster thruster designs provide a small footprint saving spacecraft surface
area for solar panels and or experiments. The simplicity of these thrusters will enable their broad use on micro-pico satellites for primary propulsion, ACS and formation flying
applications. Larger spacecraft may find uses for ECESP thrusters on extended booms, on-orbit refueling, pneumatic actuators, and gas generators.
The Minimum Impulse Thruster (MIT) was developed to improve the state-of-the-art minimum impulse capability of hydrazine monopropellant thrusters. Specifically, a new fast
response solenoid valve was developed, capable of responding to a much shorter electrical pulse width, thereby reducing the propellant flow time and the minimum impulse
bit. The new valve was combined with the Aerojet MR-103, 0.2 lbf (0.9 N) thruster and put through an extensive Delta-qualification test program, resulting in a factor of 5
reduction in the minimum impulse bit, from roughly 1.1 milli-lbf-seconds (5 milliNewton seconds) to - 0.22 milli-lbf-seconds (1 mN-s). To maintain it's extensive heritage, the
thruster itself was left unchanged. The Minimum Impulse Thruster provides mission and spacecraft designers new design options for precision pointing and precision
translation of spacecraft.
Mini-Micro Thrusters, LOX/Hydrocarbon Propulsion, and Attitude Control Systems
Propulsion – MonoPropellants
Company Name
Orion Propulsion, Inc.
Title
Oxygen-Methane Thruster
Field Center
MSFC
NASA SBIR Phase I
Quad Chart
Two main innovations will be developed in the Phase II effort that are fundamentally associated with our gaseous oxygen/gaseous
methane RCS thruster. The first innovation is that of an integrated torch igniter/injector which, provides simple and reliable ignition for the
thruster. The second innovation is the thruster's capability to operate with both gaseous and liquid propellants. This innovation enables
greater system flexibility that is inherent in the capacity to function over a much wider range of operating conditions. Orion has been
approached by several prospective customers concerning affordable, reusable thrusters. In the current commercial market, this does not
exist and the commercial sector cannot afford traditional propulsion solutions. For this reason, Orion started developing a thruster that
offers the following features: ? Provides an affordable solution as compared to what is currently available ? Uses a different manned
approach since it is reusable ? Uses a modular design, which allows for the robust inspection and replacement of parts ? Uses green
propellants and capitalizes on in situ propellant production ? Provides greater system flexibility since it operates with both liquid and
gaseous propellants Not only does the gas/gas operation of the thruster attract companies such as t/Space and Bigleow Aerospace, but
its liquid/liquid operation broadens this customer base and makes this design applicable to NASA programs such as the Crew
Exploration Vehicle (CEV).
Mini-Micro Thrusters, LOX/Hydrocarbon Propulsion, and Attitude Control Systems
Propulsion – MonoPropellants
Company Name
Cornerstone Research
Group, Inc.
Title
Vectored-Thrust MAV for BIA Operations in Urban Environments
DOD SBIR Phase I
Quad Chart
Cornerstone Research Group, Inc. (CRG), proposes to design and demonstrate a vertical takeoff and landing (VTOL) micro air
vehicle (MAV) with vectored thrust for high maneuverability and platform stability. CRG will design and demonstrate a MAV
providing the capabilities necessary for operating in cluttered urban environments while transferring real time reconnaissance
information for bomb impact assessment (BIA). The proposed multi-degree-of-freedom (DOF) MAV will use a vectored thrust
propulsion system designed to allow the MAV body to remain level during all stages of flight while providing high maneuverability
for collision avoidance in low-altitude operations. Lightweight structural components are essential to the design of a VTOL MAV
where thrust-to-weight ratios greater than 1 are required. To accomplish this, CRG will employ its lightweight high-strength
syntactic materials in the design of the MAV structure. In Phase I CRG will develop and validate performance of the vectored thrust
propulsion system, design a lightweight integrated structure, and select and evaluate microelectronics for control architecture
(stability, navigation, communications). In addition, CRG will fabricate and demonstrate a radio-controlled (R/C) prototype MAV.
Mini-Micro Thrusters, LOX/Hydrocarbon Propulsion, and Attitude Control Systems
Propulsion – MonoPropellants
FC
GRC
Title
HAN-Based Monopropellant Technology
Development
GRC
High-Performance Monopropellants and Catalysts
Evaluated
NTRS
Abstract
NASA Glenn Research Center is sponsoring efforts to develop technology for high-performance, high-density, low-freezing point, low-hazards
monopropellant systems. The program is focused on a family of monopropellant formulations composed of an aqueous solution of hydroxylammonium
nitrate (HAN) and a fuel component. HAN-based monopropellants offer significant mass and volume savings to small (less than 100 kg) satellite for
orbit raising and on-orbit propulsion applications. The low-hazards characteristics of HAN-based monopropellants make them attractive for
applications where ground processing costs are a significant concern. A 1-lbf thruster has been demonstrated to a 20-kg satellite orbit insertion duty
cycle, using a formulation compatible with currently available catalysts. To achieve specific impulse levels above those of hydrazine, catalyst materials
that can withstand the high-temperature, corrosive combustion environment of HAN-based monopropellants have to be developed. There also needs
to be work done to characterize propellant properties, burning behavior, and material compatibility. NASA is coordinating their monopropellant efforts
with those of the United States Air Force.
The NASA Glenn Research Center is sponsoring efforts to develop advanced monopropellant technology. The focus has been on monopropellant
formulations composed of an aqueous solution of hydroxylammonium nitrate (HAN) and a fuel component. HAN-based monopropellants do not have
a toxic vapor and do not need the extraordinary procedures for storage, handling, and disposal required of hydrazine (N2H4). Generically, HAN-based
monopropellants are denser and have lower freezing points than N2H4. The performance of HAN-based monopropellants depends on the selection of
fuel, the HAN-to-fuel ratio, and the amount of water in the formulation. HAN-based monopropellants are not seen as a replacement for N2H4 per se,
but rather as a propulsion option in their own right. For example, HAN-based monopropellants would prove beneficial to the orbit insertion of small,
power-limited satellites because of this propellant's high performance (reduced system mass), high density (reduced system volume), and low
freezing point (elimination of tank and line heaters). Under a Glenn-contracted effort, Aerojet Redmond Rocket Center conducted testing to provide
the foundation for the development of monopropellant thrusters with an I(sub sp) goal of 250 sec. A modular, workhorse reactor (representative of a 1lbf thruster) was used to evaluate HAN formulations with catalyst materials. Stoichiometric, oxygen-rich, and fuelrich formulations of HAN-methanol
and HAN-tris(aminoethyl)amine trinitrate were tested to investigate the effects of stoichiometry on combustion behavior. Aerojet found that fuelrich
formulations degrade the catalyst and reactor faster than oxygen-rich and stoichiometric formulations do. A HAN-methanol formulation with a
theoretical Isp of 269 sec (designated HAN269MEO) was selected as the baseline. With a combustion efficiency of at least 93 percent demonstrated
for HAN-based monopropellants, HAN269MEO will meet the I(sub sp) 250 sec goal.
Mini-Micro Thrusters, LOX/Hydrocarbon Propulsion, and Attitude Control Systems
Propulsion – Propellant Storage
FC
Glenn Research Center
Title
Clay Nanocomposite Aerogel
Sandwich Structures for
Cryotanks
Glenn Research Center
Epoxy and Layered Silicate
Nanocomposite Tanks Produced
and Tested for Cryogen Storage
Applications
Marshall Space Flight
Center
Hydrogen Permeability of a
Polymer Based Composite Tank
Material Under Tetra-Axial Strain
NTRS
Abstract
GRC research has led to the development of epoxy-clay nanocomposites with 60-70 lower gas permeability than the base epoxy resin. Filament wound carbon fiber reinforced
tanks made with this nanocomposite had a five-fold lower helium leak rate than the corresponding tanks made without clay. More recent work has produced new composites
with more than a 100-fold reduction in helium permeability. Use of these advanced, high barrier composites would eliminate the need for a liner in composite cryotanks, thereby
simplifying construction and reducing propellant leakage. Aerogels are attractive materials for use as cryotank insulation because of their low density and low thermal
conductivity. However, aerogels are fragile and have poor environmental stability, which have limited their use to certain applications in specialized environments (e.g., in certain
types of nuclear reactors as Cerenkov radiation detectors, and as thermal insulators aboard space rovers on Mars). New GRC developed polymer crosslinked aerogels (XAerogels) retain the low density of conventional aerogels, but they demonstrate a 300-fold increase in their mechanical strength. Currently, our strongest materials combine a
density of approx. 0.45 g cc, a thermal conductivity of approx. 0.04 W mK and a compressive strength of 185 MPa. Use of these novel aerogels as insulation materials structural
components in combination with the low permeability of epoxy-clay nanocomposites could significantly reduce cryotank weight and improve durability.
It is envisioned that next-generation space exploration vehicles will have integral liquid hydrogen and liquid oxygen cryogenic fuel tanks that not only contain fuel, but function as
load-carrying structures during launch and flight operations. Traditionally, metallic tanks have been used for housing cryogenic fluids. The advantages of such tanks include high
strength, high stiffness, and low permeability. Presently, it appears that the replacement of traditional metallic cryogenic fuel tanks with polymer matrix composite tanks may
decrease weight significantly, and hence, increase load-carrying capabilities (ref. 1). However, the tanks must be able to withstand flight loads and temperatures ranging from
250 to 120 C, without loss of cryogenic fuel due to microcracking or delamination. Research of the NASA Glenn Research Center has led to the development of epoxy-clay
nanocomposites with up to 70-percent lower hydrogen permeability than that of the base epoxy resin. Filament-wound carbon-fiber-reinforced tanks made with this
nanocomposite had a fivefold lower helium leak rate than the corresponding tanks made without clay. Use of these advanced composites would eliminate the need for a liner in
composite cryotanks, thereby simplifying construction and reducing propellant leakage.
In order to increase the performance of future expendable and reusable launch vehicles and reduce per-pound payload launch costs, weight reductions have been sought in
vehicle components. Historically, the cryogenic propellant tanks for launch vehicles have been constructed from metal. These are some of the largest structural components in
the vehicle and contribute significantly to the vehicles total dry weight. A successful replacement material will be conformable, have a high strength to weight ratio, and have a
low gas-permeability to the cryogens being stored, i.e., oxygen and hydrogen. Polymer-based composites are likely candidates to fill this role. Polymer and polymer-based
composites in general are known to have acceptable gas permeation properties in their as-cured state.1 The use of polymer-based composites for this application has been
proposed for some time.2 Some successes have been reported with oxygen3, but other than the DC-XA experience, those with hydrogen have been limited. The primary reason
for this has been the small molecular diameter of hydrogen, the lower temperatures of the liquid, and that the composite materials examined to date have all been susceptible to
microcrack formation in response to the thermal-mechanical cycles experienced in the use-environment. There have been numerous accounts of composite materials with
reported acceptable resistance to the formation of microcracks when exposed to various mechanical and or thermal cycles. However, virtually all of these studies have
employed uniaxial loads and there has been no discussion or empirical evidence pertaining to how these loads relate to the biaxial state of stress in the material in its use
environment. Furthermore, many of these studies have suffered from a lack of instrument sensitivity in detecting hydrogen permeability, no standards, insufficient documentation
of test conditions, testing of cycled materials in their unload state, and or false assumptions about the nature of the microcracks in the material. This paper documents the results
of hydrogen permeability testing on a Bismaleimide (BMI) based graphite fiber composite material under a variety of tetra-axial strain states.
Mini-Micro Thrusters, LOX/Hydrocarbon Propulsion, and Attitude Control Systems
Propulsion – Propellant Storage
FC
MSFC
Title
Ion Storage Tests with the High
Performance Antimatter Trap
(HiPAT)
Glenn Research Center
Polymer-Layered Silicate
Nanocomposites for Cryotank
Applications
Jet Propulsion
Laboratory; Johnson
Space Center; White
Sands Test Facility
Stress Rupture Testing and
Analysis of the NASA WSTF-JPL
Carbon Overwrapped Pressure
Vessels
Marshall Space Flight
Center
Toroidal Tank Development for
Upper-stages
Jet Propulsion Laboratory
Ultralight propellant tank for Mars
exploration rover cruise stage
NTRS
Abstract
The NASA Marshall Space Flight Centers (NASA MSFC) Propulsion Research Center (PRC) is evaluating an antiproton storage system, referred to as the High Performance
Antiproton Trap (HiPAT). This interest stems from the sheer energy represented by matter antimatter annihilation process with has an energy density approximately 10 order of
magnitude above that of chemical propellants. In other terms, one gram of antiprotons contains the equivalent energy of approximately 23 space shuttle external tanks or ET's
(each ET contains roughly 740,000 kgs of fuel and oxidizer). This incredible source of stored energy, if harnessed, would be an enabling technology for deep space mission
where both spacecraft weight and propulsion performance are key to satisfying aggressive mission requirements. The HiPAT hardware consists of a 4 Tesla superconductor
system, an ultra high vacuum test section (vacuum approaching 10(exp -12) torr), and a high voltage confinement electrode system (up to 20 kvolts operation). The current
laboratory layout is illustrated. The HiPAT designed objectives included storage of up to 1 trillion antiprotons with corresponding lifetimes approaching 18 days. To date, testing
has centered on the storage of positive hydrogen ions produced in situ by a stream of high-energy electrons that passes through the trapping region. However, due to space
charge issues and electron beam compression as it passes through the HiPAT central field, current ion production is limited to less then 50,000 ions. Ion lifetime was determined
by counting particle populations at the end of various storage time intervals. Particle detection was accomplished by destructively expelling the ions against a micro-channel
plate located just outside the traps magnetic field. The effect of radio frequency (RF) stabilization on the lifetime of trapped particles was also examined. This technique, referred
to as a rotating wall, made use of a segmented electrode located near the center of the trap on which various phases of a particular frequency were applied. Various
experiments were performed illustrating the ability of an RF drive to both prolong and reduced the lifetimes of various ion species depending on the selected frequency. HiPAT is
now being reconfigured for testing with an ion source that will provide both positive and negative hydrogen ions from an external source. This ion system shall provide higher fill
capacity (order of million of ions per shot), stacking of multiple shots, and injection schemes typical of a realistic antiproton delivery system.
Previous composite cryotank designs have relied on the use of conventional composite materials to reduce microcracking and permeability. However, revolutionary advances in
nanotechnology derived materials may enable the production of ultra-lightweight cryotanks with significantly enhanced durability and damage tolerance, as well as reduced
propellant permeability. Layered silicate nanocomposites are especially attractive in cryogenic storage tanks based on results that have been reported for epoxy nanocomposite
systems. These materials often exhibit an order of magnitude reduction in gas permeability when compared to the base resin. In addition, polymer-silicate nanocomposites have
been shown to yield improved dimensional stability, strength, and toughness. The enhancement in material performance of these systems occurs without property trade-offs
which are often observed in conventionally filled polymer composites. Research efforts at NASA Glenn Research Center have led to the development of epoxy-clay
nanocomposites with 70 lower hydrogen permeability than the base epoxy resin. Filament wound carbon fiber reinforced tanks made with this nanocomposite had a five-fold
lower helium leak rate than the corresponding tanks made without clay. The pronounced reduction observed with the tank may be due to flow induced alignment of the clay
layers during processing. Additionally, the nanocomposites showed CTE reductions of up to 30, as well as a 100 increase in toughness.
Carbon composite overwrapped pressure vessels (COPVs) are widely used in applications from spacecraft to life support. COPV technology provides a pressurized media
storage advantage over amorphous technology with weight savings on the order of 30 percent. The National Aeronautics and Space Administration (NASA) has been supporting
the development of this technology since the early 1970's with an interest in safe application of these components to reduce mass to orbit. NASA White Sands Test Facility
(WSTF) has been testing components in support of this objective since the 1980s and has been involved in test development and analysis to address affects of impact,
propellant and cryogenic fluids exposure on Kevlar and carbon epoxy. The focus of this paper is to present results of a recent joint WSTF-Jet Propulsion Laboratories (JPL)
effort to assess safe life of these components. The WSTF-JPL test articles consisted of an aluminum liner and a carbon fiber overwrap in an industry standard epoxy resin
system. The vessels were specifically designed with one plus-minus helical wrap and one hoop wrap over the helical and they measured 4.23 x 11.4 in. long. 120 test articles
were manufactured in August of 1998 of one lot fiber and resin and the 110 test articles were delivered to WSTF for test. Ten of the 120 test articles were burst tested at the
manufacturer to establish the delivered fiber stress. Figure 1 shows a test article in a pre burst condition and with a hoop fiber failure (no leak of pressurized media) and post
burst (failure of liner and loss of pressurized media).
The advantages, development, and fabrication of toroidal propellant tanks are profiled in this viewgraph presentation. Several images are included of independent research and
development (IR and ampD) of toroidal propellant tanks at Marshall Space Flight Center (MSFC). Other images in the presentation give a brief overview of Thiokol conformal
tank technology development. The presentation describes Thiokol's approach to continuous composite toroidal tank fabrication in detail. Images are shown of continuous and
segmented toroidal tanks fabricated by Thiokol.
The design of the ultralight propellant tanks and the PMD device, flight qualification testing of the tank and the PMD system, and damage control methods is discussed.
Mini-Micro Thrusters, LOX/Hydrocarbon Propulsion, and Attitude Control Systems
Propulsion—Solar
Company Name
New Era Technology, Inc. And
Pennsylvania State University
Title
DEVELOPMENT OF A MODULAR SOLAR THERMAL
PROPULSION ENGINE
NASA SBIR Phase II
Field Center
Quad Chart
MSFC
A new modular solar powered rocket engine is proposed to provide a few milligrams to a few kilograms of thrust
at 1100-1600 seconds of specific impulse by heating hydrogen propellant to average temperatures. This
revolutionary improvement over existing solar-electric powered space propulsion systems could lead to the
development of an ultra-light, compact, and an order of magnitude lower cost propulsion system.
Mini-Micro Thrusters, LOX/Hydrocarbon Propulsion, and Attitude Control Systems
Propulsion—Solar
FC
Marshall Space Flight
Center
Title
Application of Solar-Electric
Propulsion to Robotic Missions in
Near-Earth Space
Marshall Space Flight
Center
Magnetic inflation produced by the
Mini-Magnetospheric Plasma
Propulsion (M2P2) prototype
Marshall Space Flight
Center
Results of Evaluation of Solar
Thermal Propulsion
Glenn Research Center
Solar Electric Propulsion
Technologies Being Designed for
Orbit Transfer Vehicle Applications
NTRS
Abstract
Interest in applications of solar electric propulsion (SEP) is increasing. Application of SEP technology is favored when (1) the mission is compatible with low-thrust propulsion,
(2) the mission needs high total delta V such that chemical propulsion is disadvantaged and (3) performance enhancement is needed. If all such opportunities for future
missions are considered, many uses of SEP are likely. Representative missions are surveyed and several SEP applications selected for analysis, including orbit raising, lunar
science and robotic exploration, and planetary science. These missions span SEP power range from 10 kWe to about 100 kWe. A SEP design compatible with small
inexpensive launch vehicles, and capable of lunar science missions, is presented. Modes of use and benefits are described, and potential SEP evolution is discussed.
Mini-Magnetospheric Plasma Propulsion (M2P2) seeks the creation of a magnetic wall or bubble (i.e. a magnetosphere) attached to a spacecraft that will intercept the solar
wind and thereby provide high-speed propulsion with little expenditure of propellant. The prototype uses a helicon source embedded asymmetrically in a dipole-like magnetic
field. Breakdown of the plasma can be produced at high densities 10(sup 12)-10(sup 13) cm(sup -3) with a temperature of several eV. The plasma pressure is sufficient to
cause the outward expansion or inflation of the mini-magnetosphere. This expansion has now been measured directly by magnetic field probes. Computer simulations of the
laboratory geometry show the presence of magnetic field perturbations that have similar magnitude and temporal variations as seen in the experiments. The field line mapping
from the model has similar features to the optical images taken during laboratory prototype. The agreement between the laboratory experiments and the computer simulations
provide quantitative evidence that inflation of a dipole field can be achieved in the laboratory, essentially all the way out to the chamber walls which in the large chamber
experiments corresponds to several tens of magnet radii. The operation characteristics of the prototype are consistent with the initial simulations that indicated that if such a
device were operated in space then it could produce a mini-magnetosphere of the order of about 15-20 km. Such a mini-magnetosphere would experience 1-3 N of thrust from
the solar wind, while requiring on 1-2 kWe of power to sustain the mini-magnetosphere. copyright 2002 American Institute of Physics.
The solar thermal propulsion evaluation reported here relied on prior research for all information on solar thermal propulsion technology and performance. Sources included
personal contacts with experts in the field in addition to published reports and papers. Mission performance models were created based on this information in order to estimate
performance and mass characteristics of solar thermal propulsion systems. Mission analysis was performed for a set of reference missions to assess the capabilities and
benefits of solar thermal propulsion in comparison with alternative in-space propulsion systems such as chemical and electric propulsion. Mission analysis included estimation
of delta V requirements as well as payload capabilities for a range of missions. Launch requirements and costs, and integration into launch vehicles, were also considered.
The mission set included representative robotic scientific missions, and potential future NASA human missions beyond low Earth orbit. Commercial communications satellite
delivery missions were also included, because if STP technology were selected for that application, frequent use is implied and this would help amortize costs for technology
advancement and systems development. A C3 Topper mission was defined, calling for a relatively small STP. The application is to augment the launch energy (C3) available
from launch vehicles with their built-in upper stages. Payload masses were obtained from references where available. The communications satellite masses represent the
range of payload capabilities for the Delta IV Medium and or Atlas launch vehicle family. Results indicated that STP could improve payload capability over current systems, but
that this advantage cannot be realized except in a few cases because of payload fairing volume limitations on current launch vehicles. It was also found that acquiring a more
capable (existing) launch vehicle, rather than adding an STP stage, is the most economical in most cases.
There is increasing interest in employing Solar Electric Propulsion (SEP) for new missions requiring transfer from low Earth orbit to the Earth-Moon Lagrange point, L1.
Mission architecture plans place the Gateway Habitat at L1 in the 2011 to 2016 timeframe. The Gateway Habitat is envisioned to be used for Lunar exploration, space
telescopes, and planetary mission staging. In these scenarios, an SEP stage, or "tug," is used to transport payloads to L1--such as the habitat module, lunar excursion and
return vehicles, and chemical propellant for return crew trips. SEP tugs are attractive because they are able to efficiently transport large (less than 10,000 kg) payloads while
minimizing propellant requirements. To meet the needs of these missions, a preliminary conceptual design for a general-purpose SEP tug was developed that incorporates
several of the advanced space power and in-space propulsion technologies (such as high-power gridded ion and Hall thrusters, high-performance thin-film photovoltaics,
lithium-ion batteries, and advanced high-voltage power processing) being developed at the NASA Glenn Research Center. A spreadsheet-based vehicle system model was
developed for component sizing and is currently being used for mission planning. This model incorporates a low-thrust orbit transfer algorithm to make preliminary
determinations of transfer times and propellant requirements. Results from this combined tug mass estimation and orbit transfer model will be used in a higher fidelity
trajectory model to refine the analysis.
Mini-Micro Thrusters, LOX/Hydrocarbon Propulsion, and Attitude Control Systems
Propulsion—Solar
FC
Glenn Research Center
Title
Solar Electric Propulsion Vehicle
Design Study for Cargo Transfer to
Earth-Moon L1
NTRS
Abstract
A design study for a cargo transfer vehicle using solar electric propulsion was performed for NASA's Revolutionary Aerospace Systems Concepts program. Targeted for 2016,
the solar electric propulsion (SEP) transfer vehicle is required to deliver a propellant supply module with a mass of approximately 36 metric tons from Low Earth Orbit to the
first Earth-Moon libration point (LL1) within 270 days. Following an examination of propulsion and power technology options, a SEP transfer vehicle design was selected that
incorporated large-area (approx. 2700 sq m) thin film solar arrays and a clustered engine configuration of eight 50 kW gridded ion thrusters mounted on an articulated boom.
Refinement of the SEP vehicle design was performed iteratively to properly estimate the required xenon propellant load for the out-bound orbit transfer. The SEP vehicle
performance, including the xenon propellant estimation, was verified via the SNAP trajectory code. Further efforts are underway to extend this system model to other orbit
transfer missions.
Mini-Micro Thrusters, LOX/Hydrocarbon Propulsion, and Attitude Control Systems
Propulsion—Tethers
FC
Marshall Space Flight Center
Title
A Model for Dynamic Simulation and Analysis of Tether
Momentum Exchange
NASA (non Center Specific)
The muTORQUE momentum-exchange tether experiment
NTRS
Abstract
Momentum-exchange electrodynamic reboost (MXER) tether systems may enable high-energy missions to the Moon, Mars, and beyond
by serving as an 'upper stage in space'. Existing rockets that use an MXER tether station could double their capability to launch
communications satellites and help improve US competitiveness. A MXER tether station would boost spacecraft from low Earth orbit to
a high-energy orbit quickly, like a high-thrust rocket. Then, using the same principles that make an electric motor work, it would slowly
rebuild its orbital momentum by pushing against the Earth's magnetic field-without using any propellant. One of the significant
challenges in developing a momentum-exchange electrodynamic reboost tether systems is in the analysis and design of the capture
mechanism and its effects on the overall dynamics of the system. This paper will present a model for a momentum-exchange tether
system that can simulate and evaluate the performance and requirements of such a system.
Long, high-strength tethers can provide a mechanism for transferring orbital momentum and energy from one space object to another
without the consumption of propellant. By providing a highly-reusable transportation architecture, systems built upon such "momentumexchange" tethers may be able to achieve significant cost reductions for a number of in-space propulsion missions. Before such systems
could be placed into operation, however, a number of technical challenges must be met, including flight demonstration of high-strength,
highly survivable tethers, demonstration of the ability to control the dynamics of a rotating tether system, and the ability for a tether
system to rendezvous with, capture, and then toss a payload. In this paper, we discuss a concept design for a small momentum
exchange tether experiment that is intended to serve as the first step in demonstrating these key technologies. The "Microsatellite
Tethered Orbit Raising Qualification Experiment" (muTORQUE) will be designed to fly as a secondary payload on an upper stage of a
rocket used to deliver a satellite to GEO. The muTORQUE experiment will remain on the upper stage left in a GTO trajectory. After the
primary satellite has been deployed into GEO, the muTORQUE experiment will deploy a microsatellite at the end of a 20 km long tether.
Utilizing tether reeling and or electrodynamic propulsion, the muTORQUE system will set the tether in rotation around the upper stage,
accelerating the rotation until the tip velocity is approximately 400 m s. The experiment will then release the microsatellite when the
system is at its perigee, tossing the payload into a near-minimum-energy transfer to the Moon. The microsatellite can then utilize a
Belbruno weak-boundary trajectory to transfer into a lunar orbit using only a few m s of delta-V. Preliminary analyses indicate that the
tether system could be mass-competitive with a chemical propellant system for the same mission. copyright 2002 American Institute of
Physics.
Mini-Micro Thrusters, LOX/Hydrocarbon Propulsion, and Attitude Control Systems
Structures--Launch and Flight Vehicle
Company Name
Honeybee Robotics
Title
Structural Attachments for Rapid
Assembly of Satellites
DOD SBIR Phase I
Quad Chart
A goal of the Operationally Responsive Space (ORS) thrust is to enable turn around of a tactical satellite within six days, from mission call-up to on-orbit
operation. The ability to produce such a spacecraft would be a vast departure from the norm of large, complex, costly custom spacecraft that require a period of
years to deploy. Considerable attention has been paid to modular architectures for standardizing spacecraft. Honeybee Robotics Spacecraft Mechanisms
Corporation, in partnership with leading small spacecraft provider AeroAstro, proposes to develop a rapid assembly system for the quick completion of
structural, electrical, and thermal connections between modular satellite panels. The system will build upon Honeybee’s patented "Quick Insertion Nut"
fastening technology, which has the benefit of being similar to the standard bolted joint typically used in the aerospace industry. The system will enable reduced
satellite build time, and will facilitate the rapid assembly
and disassembly of panels to keep pace with the possibility of changing mission requirements. The proposed Phase 1 effort will include requirements
derivation, concept development, detailed design and analysis, and iterative breadboarding and demonstration testing. The demonstration will consist of
assembling and subsequently disassembling 3 panels with the attachment mechanism to form a representative "corner" of a spacecraft and verify electrical and
thermal continuity across the joint. This breadboard will demonstrate the feasibility of the attachment concept itself for fit, speed and ease of assembly and
disassembly, stiffness, and electrical and thermal connection.
Mini-Micro Thrusters, LOX/Hydrocarbon Propulsion, and Attitude Control Systems
Structures--Launch and Flight Vehicle
WIPO
Patent Number
WO05073085A1
Title
In Orbit Space Transportation &
Recovery System
Assignee
Iostar
Corporation,
Jehan, Robert
Abstract
An In Orbit Transportation & Recovery System (10) is disclosed. One preferred embodiment of the present invention comprises a space tug
powered by a nuclear reactor (19). The system includes a collapsible boom (110 connected at one end to a propellant tank (13) which stores fuel for
an electric propulsion system (12). This end of the boom (11) is equipped with docking hardware (14) that is able to grasp and hold a satellite (15)
and as a means to refill the tank (13). Radiator panels (16) mounted on a boom (11) dissipate heat from the reactor (19). A radiation shield (20) is
situated next to the reactor (19) to protect the satellite payload (15) at the far end of the boom 9110. The system (10) will be capable of
accomplishing rendezvous and docking maneuvers which will enable it to move spacecraft between a low Earth parking orbit and positions in
higher orbits or to other locations in our Solar System.
Mini-Micro Thrusters, LOX/Hydrocarbon Propulsion, and Attitude Control Systems
Structures—Tankage
Company Name
Microcosm, Inc.
Title
Lightweight, Composite Cryogenic Tank
Structures
Field Center
LARC
NASA SBIR Phase I
Quad Chart
Microcosm has developed and qualified strong, all-composite LOX tanks for launch vehicles. Our new 42-inch diameter tank design weighs
486 lbs and burst without leaking at 2,125 psi, within 3.5% of the predicted burst pressure. This SBIR will analyze, design, build, and test much
lighter weight all composite cryogenic tanks and examine, develop, and test alternative insulation techniques to minimize boil-off. This SBIR
will also examine the reuse of propellant tanks as crew and storage habitats. During Phase I, we will design and fabricate 12 10-inch diameter
and 2 25-inch diameter cryogenic tanks with a design burst pressure of approximately 850 psi. Eight of the 10-inch tanks and one 25-inch tank
will be thermally cycled and burst tested using liquid nitrogen to obtain statistical data. The remaining 4 10-inch tanks will first be thermally
cycled, then flushed out and re-pressurized with gaseous helium to simulate reuse as a crew habitat. The remaining 25-inch tank will be
delivered to NASA for further testing. Phase II will fabricate, build, and test larger tanks and tanks specifically intended to meet the needs of
future NASA programs, and alternative insulation approaches will be evaluated to minimize boil-off.
Mini-Micro Thrusters, LOX/Hydrocarbon Propulsion, and Attitude Control Systems
Verification and Validation--Simulation Modeling Environment
Company Name
James R. Gloudemans
Title
An advanced open-source aircraft design
platform for personal air vehicle geometry,
aerodynamics, and structures
Field Center
LARC
NASA SBIR Phase I
Quad Chart
Innovators working to revolutionize air travel through personal aviation pioneers need innovative aircraft design tools. Vehicle Sketch Pad (VSP)
is an aircraft geometry tool for rapid evaluation of advanced design concepts. VSP will be extended to include support for the modeling of aircraft
structural layout and a modular system for integrating engineering analyses. These modifications will allow VSP to unify geometry,
aerodynamics, and structures in the early design of advanced concepts. VSP will be released as open-source; a community for its development
will be initiated and fostered. Open distribution will ensure that VSP is available to all, thereby supporting personal aviation innovators,
universities, NASA, and the whole aerospace industry. This will enable a new level of fidelity and accuracy in personal aircraft design needed to
meet the aggressive goals required for the success of personal air vehicles. An improved and open VSP will catalyze personal aviation by
supporting breakthroughs in noise and cost reduction and ease of operations.
Mini-Micro Thrusters, LOX/Hydrocarbon Propulsion, and Attitude Control Systems
Verification and Validation--Simulation Modeling Environment
Company Name
Metacomp Technologies, Inc.
Title
Combustion Stability Innovations for Liquid
Rocket
DOD SBIR Phase I
Quad Chart
The proposed effort can provide the framework for a unified modeling & simulation program to substantially improve the capability for understanding,
analyzing, and predicting the underlying mechanisms dictating the combustion stability characteristics of liquid-propellant rocket engines. Emphasis will be
placed on the effects of atomization, mixing, and distributed combustion of liquid oxygen (LOX) and kerosene propellants under conditions representative of
oxygen-rich preburner staged-combustion cycle (ORPSC) engines. New models and solution methodologies will be developed and validated, where
necessary, to further enhance the solution accuracy and efficiency. The Phase I effort will focus on the flow evolution and flame dynamics of single element
coaxial swirl injectors typical of the preburner and main chamber applications with liquid/liquid and gas/liquid mixtures, respectively.
Mini-Micro Thrusters, LOX/Hydrocarbon Propulsion, and Attitude Control Systems
Verification and Validation--Simulation Modeling Environment
DOD SBIR Phase II
Quad Chart
This technology addresses the limitations of existing computational fluid dynamic models with advances in numeric's and parallel code
architecture, as well as by inclusion of pertinent physics and thermochemistry. This “smart” parallel adaptive domain decomposition for
structured and unstructured grids and related numerical advances will reduce run time by an order of magnitude.
Company Name
Combustion Research & Flow
Technology, Inc.
Title
TRANSIENT JET-INTERACTION COMBUSTION
MODELING
SRS Technologies (A subsidiary of
ManTech
International Corporation)
VISUAL MODELING TOOL FOR KILL VEHICLE
PROPULSION SYSTEM EVALUATION
SRS is now prepared to incrementally develop the Phase I DACS Modeling Tool into a fully capable, user-friendly, KV propulsion modeling
tool, to ultimately enable lighter, faster, more reliable and less expensive missile defense kill vehicles and interceptors. The modeling
options will be expanded to include non-traditional solid DACS systems, liquid DACS systems, gel DACS systems, and cold-gas thrusters.
Technical Solutions Inc.
VIRTUAL SYSTEM INTEGRATION LAB (VSIL) – A
FLEXIBLE SYTEM INTEGRATION TOOL FOR
VIRTUAL PROTOTYPING & SIMULATION
VSIL provides a means of simulating, testing and evaluating vehicle system components and their interfaces and integration into a
functioning system. The VSIL will include an Internet-accessible Repository of models, simulations and interface specifications to facilitate
development of subsystem-specific simulations by second tier contractors, thereby increasing the assurance of interoperability in emerging
systems.
Mini-Micro Thrusters, LOX/Hydrocarbon Propulsion, and Attitude Control Systems
Verification and Validation--Simulation Modeling Environment
FC
Goddard Space Flight Center
Title
A Survey Of Earth-Moon Libration Orbits
Stationkeeping Strategies And Intra-Orbit
Transfers
Jet Propulsion Laboratory
Low-thrust orbit transfer optimization with refined
Q-law and multi-objective genetic algorithm
NTRS
Abstract
Cislunar space is a readily accessible region that may well develop into a prime staging area in the effort to colonize space near Earth or to colonize the
Moon. While there have been statements made by various NASA programs regarding placement of resources in orbit about the Earth-Moon Lagrangian
locations, there is no survey of the total cost associated with attaining and maintaining these unique orbits in an operational fashion. Transfer trajectories
between these orbits required for assembly, servicing, and positioning of these resources have not been extensively investigated. These orbits are
dynamically similar to those used for the Sun-Earth missions, but differences in governing gravitational ratios and perturbation sources result in unique
characteristics. We implement numerical computations using high fidelity models and linear and nonlinear targeting techniques to compute the various
maneuver (Delta)V and temporal costs associated with orbits about each of the Earth-Moon Lagrangian locations (L1, L2, L3, L4, and L5). From a dynamical
system standpoint, we speak to the nature of these orbits and their stability. We address the cost of transfers between each pair of Lagrangian locations.
An optimization method for low-thrust orbit transfers around a central body is developed using the Q-law and a multi-objective genetic algorithm. in the
hybrid method, the Q-law generates candidate orbit transfers, and the multi-objective genetic algorithm optimizes the Q-law control parameters in order to
simultaneously minimize both the consumed propellant mass and flight time of the orbit tranfer. This paper addresses the problem of finding optimal orbit
transfers for low-thrust spacecraft.
Mini-Micro Thrusters, LOX/Hydrocarbon Propulsion, and Attitude Control Systems
Verification and Validation--Testing Facilities
FC
JSC
Title
A Facility for Testing High-Power Electric Propulsion
Systems in Space A Design Study
GRC/JSC/MSFC
A High-power Electric Propulsion Test Platform in Space
NTRS
Abstract
This paper will describe the results of the preliminary phase of a NASA design study for a facility to test high-power electric propulsion systems in
space. The results of this design study are intended to provide a firm foundation for a subsequent detailed design and development activities leading
to the deployment of a valuable space facility supporting the new vision of space exploration. The objectives for human and robotic exploration of
space can be accomplished affordably, safely and effectively with high-power electric propulsion systems. But, as thruster power levels rise to the
hundreds of kilowatts and up to megawatts, their testing will pose stringent and expensive demands on existing Earth-based vacuum facilities. These
considerations and the access to near-Earth space provided by the International Space Station (ISS) have led to a renewed interest in space testing.
The ISS could provide an excellent platform for a space-based test facility with the continuous vacuum conditions of the natural space environment
and no chamber walls to modify the open boundary conditions of the propulsion system exhaust. The platform would be designed to accommodate the
side-by-side testing of multiple types of electric thrusters currently under development and thus provide a strong basis for comparing their relative
performance. The utility of testing on the station is further enhanced by the human presence, enabling close interaction with and modification of the
test hardware in a true laboratory environment. These conditions facilitate rapid development and flight certification at potentially lower cost than with
conventional Earth-bound facilities. As an added benefit, the propulsive effect of these tests could provide some drag compensation for the station,
reducing the re-boost cost for the orbital facility. While it is expected that the ISS will not be capable of generating continuous levels of high power, the
utilization of state-of-the-art energy storage media would be sufficient to achieve very high power levels over intervals short enough to be feasible and
long enough to provide ample demonstration of steady-state operation. This paper will outline the results of the preliminary phase of the design study
with emphasis on the requirements that will dictate the system design.
This paper will describe the results of the preliminary phase of a NASA design study for a facility to test high-power electric propulsion systems in
space. The results of this design study are intended to provide a firm foundation for subsequent detailed design and development activities leading to
the deployment of a valuable space facility. The NASA Exploration Systems Mission Directorate is sponsoring this design project. A team from the
NASA Johnson Space Center, Glenn Research Center, the Marshall Space Flight Center and the International Space Station Program Office is
conducting the project. The test facility is intended for a broad range of users including government, industry and universities. International
participation is encouraged. The objectives for human and robotic exploration of space can be accomplished affordably, safely and effectively with
high-power electric propulsion systems. But, as thruster power levels rise to the hundreds of kilowatts and up to megawatts, their testing will pose
stringent and expensive demands on existing Earth-based vacuum facilities. These considerations and the human access to near-Earth space
provided by the International Space Station (ISS) have led to a renewed interest in space testing. The ISS could provide an excellent platform for a
space-based test facility with the continuous vacuum conditions of the natural space environment and no chamber walls to modify the open boundary
conditions of the propulsion system exhaust. The test platform could take advantage of the continuous vacuum conditions of the natural space
environment. Space testing would provide open boundary conditions without walls, micro-gravity and a realistic thermal environment. Testing on the
ISS would allow for direct observation of the test unit, exhaust plume and space-plasma interactions. When necessary, intervention by on-board
personnel and post-test inspection would be possible. The ISS can provide electrical power, a location for diagnostic instruments, data handling and
thermal control. The platform will be designed to accommodate the side-by-side testing of multiple types of electric thrusters. It is intended to be a
permanent facility in which different thrusters can be tested over time. ISS crews can provide maintenance for the platform and change out thruster
test units as needed. The primary objective of this platform is to provide a test facility for electric propulsion devices of interest for future exploration
missions. These thrusters are expected to operate in the range of hundreds of kilowatts and above. However, a platform with this capability could also
accommodate testing of thrusters that require much lower power levels. Testing at the higher power levels would be accomplished by using power
fiom storage devices on the platform, which would be gradually recharged by the ISS power generation system. This paper will summarize the results
of the preliminary phase of the study with an explanation of the user requirements and the initial conceptual design. The concept for test operations will
also be described. The NASA project team is defining the requirements but they will also reflect the inputs of the broader electric propulsion
community including those at universities, commercial enterprises and other government laboratories. As a facility on the International Space Station,
the design requirements are also intended to encompass the needs of international users. Testing of electric propulsion systems on the space station
will help advance the development of systems needed for exploration and could also serve the needs of other customers. Propulsion systems being
developed for commercial and military applications could be tested and certification testing of mature thrusters could be accomplished in the space
environment.
Low-Cost Assembly, Integration, and Testing
Avionics and Astrionics--Guidance, Navigation and Control
FC
Jet Propulsion Laboratory
Title
Deep Impact Sequence Planning Using MultiMission Adaptable Planning Tools With
Integrated Spacecraft Models
NTRS
Abstract
The Deep Impact mission was ambitious and challenging. JPL's well proven, easily adaptable multi-mission sequence planning tools combined with
integrated spacecraft subsystem models enabled a small operations team to develop, validate, and execute extremely complex sequence-based
activities within very short development times. This paper focuses on the core planning tool used in the mission, APGEN. It shows how the multimission design and adaptability of APGEN made it possible to model spacecraft subsystems as well as ground assets throughout the lifecycle of the
Deep Impact project, starting with models of initial, high-level mission objectives, and culminating in detailed predictions of spacecraft behavior during
mission-critical activities.
Low-Cost Assembly, Integration, and Testing
Communication--Autonomous Control and Monitoring
Company Name
Invocon, Inc.
Title
High Impact G-Loading Data Acquisition
System (HI-G-DAS)
DOD SBIR Phase I
Quad Chart
A system is proposed for high-performance wireless data acquisition in high-impact environments, specifically the launch environment of Naval payloads.
The High Impact G-Loading Data Acquisition System (HI-G-DAS) is a miniature, low power instrumentation system designed to maximize flexibility by
combining precision data
gathering capability, user programmability, and long battery life. HI-G-DAS will use state of the art techniques for component selection, PCB design, and
modeling of critical elements to withstand acceleration on the order of 1000 g's. The robust system will be capable of interfacing with multiple types of
sensors through the use of proven modular techniques and channel-specific programmability. Using wireless communication and battery charging
techniques, the system can be fully protected from water ingress in harsh sea conditions. Additionally, its automated underwater beacon will help to locate
the device in the event that recovery is necessary. The use of standard software & hardware interfaces as well as common wireless protocols will help to
ensure that operation of the system is simple and intuitive.
Low-Cost Assembly, Integration, and Testing
Information--Autonomous Reasoning/Artificial Intelligence
FC
Langley Research Center
Title
Integrated System-Level Optimization for Concurrent
Engineering With Parametric Subsystem Modeling
NTRS
Abstract
The introduction of concurrent design practices to the aerospace industry has greatly increased the productivity of engineers and teams
during design sessions as demonstrated by JPL's Team X. Simultaneously, advances in computing power have given rise to a host of potent
numerical optimization methods capable of solving complex multidisciplinary optimization problems containing hundreds of variables,
constraints, and governing equations. Unfortunately, such methods are tedious to set up and require significant amounts of time and
processor power to execute, thus making them unsuitable for rapid concurrent engineering use. This paper proposes a framework for
Integration of System-Level Optimization with Concurrent Engineering (ISLOCE). It uses parametric neural-network approximations of the
subsystem models. These approximations are then linked to a system-level optimizer that is capable of reaching a solution quickly due to the
reduced complexity of the approximations. The integration structure is described in detail and applied to the multiobjective design of a
simplified Space Shuttle external fuel tank model. Further, a comparison is made between the new framework and traditional concurrent
engineering (without system optimization) through an experimental trial with two groups of engineers. Each method is evaluated in terms of
optimizer accuracy, time to solution, and ease of use. The results suggest that system-level optimization, running as a background process
during integrated concurrent engineering sessions, is potentially advantageous as long as it is judiciously implemented.
Low-Cost Assembly, Integration, and Testing
Information--Computer Systems Architectures
Company Name
AeroAstro, Inc.
Title
FLEXIBLE AND EXTENSIBLE BUS FOR SMALL
SATELLITES (FEBSS)
DOD SBIR Phase II
Quad Chart
AeroAstro proposes to further develop modular stackable spacecraft system; the Plug'n'Sense software architecture standard which
allows spacecraft subsystems to interact without a great deal of custom hardware or software; the Universal Small Payload Interface to
make launch vehicle integration more transparent; and finally, the Flexible Extensible Bus for Small Satellites.
Low-Cost Assembly, Integration, and Testing
Information--Data Acquisition and End-to-End Management
Company Name
Vektrex Electronic Systems Inc.
Title
AUTOMATIC TEST MARKUP LANGUAGE
DOD SBIR Phase II
Quad Chart
Vektrex has a long-standing commitment to related ATS standards and technology, particularly instrumentation and instrument I/O. Current work
being done in the ATML standards group extends existing schema’s, defining additional schema, and demonstrating an implementation validating
the feasibility of XML application to ATS in a real-world application on the AF ESTS platform.
Low-Cost Assembly, Integration, and Testing
Information--Software Tools for Distributed Analysis and Simulation
Company Name
Global Strategic Solutions LLC
Title
Eliminating Legacy Performance Barriers Imposed on New Systems
DOD SBIR Phase I
Quad Chart
Cost savings and significant gains in system performance, system utilization, and throughput could be realized, when upgrading,
modernizing or replacing subsystems, if the integration of new technologies into the legacy systems were not hindered by the
constraints and restrictions imposed by the existing system design structure. Focusing on the issue of Automatic Test System (ATS)
modernization, this project investigates the feasibility of developing and implementing a new, non-traditional system integration
approach (paradigm) to minimize, or eliminate the constraints and restrictions which surface when new test technologies are
introduced into existing ATS (legacy) system designs, concepts or frameworks. The effort considers the application of a model-driven
system integration strategy, eliminating the dependency of the legacy ATS applications to the specific test resources used in the
system, and formulating a new ATS design framework to facilitate the insertion of new technologies into legacy systems throughout
their life cycle. Application of the approach to other systems and subsystem domains is also considered.
Low-Cost Assembly, Integration, and Testing
Propulsion--Fundamental Propulsion Physics
Company Name
Spectral Sciences, Inc.
Title
Applying Novel Computational Chemistry Methods to Plume Signature
Modeling of Low Thrust Propulsion Systems
DOD SBIR Phase I
Quad Chart
Understanding and modeling low thrust systems requires a fundamental understanding of the underlying physics and
chemistry, which is not well known. Spectral Sciences, Inc. (SSI) proposes to develop and apply novel computational
chemistry methods to quantify the key chemical processes giving rise to the plume signatures of low thrust systems. These
processes will be integrated into the current suite of state-of-the-art propulsion system/plume modeling codes, including the
Air Force Research Laboratory radiation and flow code, SOCRATES-P co-developed at SSI, enabling accurate end-to-end
(combustion chamber to far field plume) simulation of low thrust plume signatures. In Phase I, we will develop and
demonstrate two innovative techniques that will form the core of a new computational chemistry toolkit called DIVERT (Directdynamics Information for Very Energetic Reaction Trajectories) targeting low thrust chemistry. We will demonstrate the proofof-concept through application of the method to key chemical mechanisms, integrate the results into SOCRATES-P, and
validate them against field and laboratory data. In Phase II, experiments will be performed characterizing engine exhaust, the
new modeling methods will be applied to all critical chemical reactions, and a wide-range of validation studies will be
performed.
Low-Cost Assembly, Integration, and Testing
Structures--Launch and Flight Vehicle
Company Name
Garvey Spacecraft Corp.
Title
Demonstration and Analysis of Reusable Launch
Vehicle Operations
DOD SBIR Phase I
Quad Chart
The increased use of reusable systems continues to be one of the most promising options for creating advancements in the daily maintenance of
rocket systems, lowering hours for preparation and diminishing expenses for preparation. However, since the end of the DC-X/XA Delta Clipper
program, flight testing of candidate reusable launch vehicle (RLV) designs, technologies and operations has come to a halt. This project provides
an opportunity to quickly jump-start domestic RLV flight-based test and evaluation through the innovative use of an proven test vehicle design
that has already been developed and flown by a joint industry-academic team. By leveraging previous design and development efforts, flight
testing will actually commence in Phase I. This research will focus on conducting a rapid (i.e. - same day) turn-around between flights to identify
and assess key operational factors. A subsequent Phase II vehicle would incorporate lessons learned from these initial demonstrations, along
with more advanced technologies to expand the RLV flight research envelope. It could also serve as a prototype for an operational RLV that
might eventually replace the expendable first stage(s) of small launch vehicles (SLVs) now in development.
Low-Cost Assembly, Integration, and Testing
Verification and Validation--Operations Concepts and Requirements
NTRS
Abstract
The Responsive Space initiative has several implications for flight software that need to be addressed not only within the run-time element, but the development infrastructure and
software life-cycle process elements as well. The runtime element must at a minimum support Plug and amp Play, while the development and process elements need to incorporate
methods to quickly generate the needed documentation, code, tests, and all of the artifacts required of flight quality software. Very rapid response times go even further, and imply little
or no new software development, requiring instead, using only predeveloped and certified software modules that can be integrated and tested through automated methods. These
elements have typically been addressed individually with significant benefits, but it is when they are combined that they can have the greatest impact to Responsive Space. The Flight
Software Branch at NASA's Goddard Space Flight Center has been developing the runtime, infrastructure and process elements needed for rapid integration with the Core Flight
software System (CFS) architecture. The CFS architecture consists of three main components the core Flight Executive (cFE), the component catalog, and the Integrated Development
Environment (DE). This paper will discuss the design of the components, how they facilitate rapid integration, and lessons learned as the architecture is utilized for an upcoming
spacecraft.
Modular, Reconfigurable, and Rapid-response (MR(sup 2)) space systems represent a paradigm shift in the way space assets of all sizes are designed, manufactured, integrated,
tested, and flown. This paper will describe the MR(sup 2) paradigm in detail, and will include guidelines for its implementation. The Remote Sensing Advanced Technology microsatellite
(RSAT) is a proposed flight system test-bed used for developing and implementing principles and best practices for MR(sup 2) spacecraft, and their supporting infrastructure. The initial
goal of this test-bed application is to produce a lightweight (approx. 100 kg), production-minded, cost-effective, and scalable remote sensing micro-satellite capable of high performance
and broad applicability. Such applications range from future distributed space systems, to sensor-webs, and rapid-response satellite systems. Architectures will be explored that strike a
balance between modularity and integration while preserving the MR(sup 2) paradigm. Modularity versus integration has always been a point of contention when approaching a design
whereas one-of-a-kind missions may require close integration resulting in performance optimization, multiple and flexible application spacecraft benefit and ampom modularity, resulting
in maximum flexibility. The process of building spacecraft rapidly ( and lt 7 days), requires a concerted and methodical look at system integration and test processes and pitfalls.
Although the concept of modularity is not new and was first developed in the 1970s by NASA's Goddard Space Flight Center (Multi-Mission Modular Spacecraft), it was never
modernized and was eventually abandoned. Such concepts as the Rapid Spacecraft Development Office (RSDO) became the preferred method for acquiring satellites. Notwithstanding,
over the past 30 years technology has advanced considerably, and the time is ripe to reconsider modularity in its own right, as enabler of R(sup 2), and as a key element of
transformational systems. The MR2 architecture provides a competitive advantage over the old modular approach in its rapid response to market needs that are difficult to predict both
from the perspectives of evolving technology, as well as mission and application requirements.
The proposed Prometheus 1 spacecraft would utilize nuclear electric propulsion to propel the spacecraft to its ultimate destination where it would perform its primary mission. As part of
the Prometheus 1 Phase A studies, system models were developed for each of the spacecraft subsystems that were integrated into one overarching system model. The Electric
Propulsion System (EPS) model was developed using data from the Prometheus 1 electric propulsion technology development efforts. This EPS model was then used to provide both
performance and mass information to the Prometheus 1 system model for total system trades. Development of the EPS model is described, detailing both the performance calculations
as well as its evolution over the course of Phase A through three technical baselines. Model outputs are also presented, detailing the performance of the model and its direct relationship
to the Prometheus 1 technology development efforts. These EP system model outputs are also analyzed chronologically showing the response of the model development to the four
technical baselines during Prometheus 1 Phase A.
FC
GSFC
Title
Implications of Responsive Space on
the Flight Software Architecture
GSFC
Modular, Reconfigurable, and Rapid
Response Space Systems The Remote
Sensing Advanced Technology
Microsatellite
GRC/JPL
Electric Propulsion System Modeling for
the Proposed Prometheus 1 Mission
JPL
Model-based engineering design for
space missions
The basic elements of model-based design for space missions have existed for almost a decade, awaiting an opportunity to implement them in the same place at the same time. In early
design phases, combinations of models, concurrent engineering methods, and scenario-driven design have been used for several years with results that have exceeded even optimistic
expectations but the goal of extending these methods to later phases of design has been more elusive. JPL's Model-Based Engineering Design (MBED) initiative will provide opportunity
to reach that goal. It enables advanced systems engineering practice through a series of integrated, increasingly detailed models that provide continuity from architectural concept
through detailed design. It extends current capability for rapid conceptual design, allowing thorough exploration of design tradespaces and selection of an optimal design point with
associated cost and rationale and it provides seamless connection to subsystem models and detailed design tool suites. In this paper we will review the goals and status of MBED and
show the expected interconnectivity between conceptual and detailed design.
GSFC
The Role of Integrated Modeling in the
Design and Verification of the James
Webb Space Telescope
This viewgraph presentation gives an overview of the architecture of the James Webb Space Telescope, and explains how integrated modeling is useful for analyzing wavefront, thermal
distortion, subsystems, and image motion jitter for the telescope design.
Low-Cost Assembly, Integration, and Testing
Verification and Validation--Testing Requirements and Architectures
Company Name
GMA Industries, Inc.
Title
AUTOMATED TEST PROGRAM SET DEVELOPMENT
USING INTEGRATED CIRCUIT ELECTROMAGNETIC
EMISSIONS
DOD SBIR Phase II
Quad Chart
This proposal describes the automated development of a test program set utilizing electromagnetic emissions from
integrated circuits to determine UUT operational status.
Low-Cost Assembly, Integration, and Testing
Verification and Validation--Testing Requirements and Architrectures
FC
GSFC
Title
Integration and Test of Shuttle Small Payloads
GSFC
Testing of Environmental Satellite Bus-Instrument Interfaces
Using Engineering Models
NTRS
Abstract
Recommended approaches for space shuttle small payload integration and test (I and T) are presented. The paper is intended for
consideration by developers of shuttle small payloads, including I and T managers, project managers, and system engineers. Examples and
lessons learned are presented based on the extensive history of NASA's Hitchhiker project. All aspects of I and T are presented, including:
(1) I and T team responsibilities, coordination, and communication; (2) Flight hardware handling practices; (3) Documentation and
configuration management; (4) I and T considerations for payload development; (5) I and T at the development facility; (6) Prelaunch
operations, transfer, orbiter integration and interface testing; (7) Postflight operations. This paper is of special interest to those payload
projects that have small budgets and few resources: that is, the truly faster, cheaper, better projects. All shuttle small payload developers
are strongly encouraged to apply these guidelines during I and T planning and ground operations to take full advantage of today's limited
resources and to help ensure mission success.
This paper discusses the formulation and execution of a laboratory test of the electrical interfaces between multiple atmospheric science
instruments and the spacecraft bus that carries them. The testing, performed in 2002, used engineering models of the instruments that will
be flown on the Aura s p a c m and of the Aura spacecraft bus electronics. Aura is one of NASA's Earth Observing System OS) Program
missions managed by the Goddard Space Flight Center. The test was designed to evaluate the complex interfaces in the spacecraft and
instrument command and data handling (C and ampDH) subsystems prior to integration of the complete flight instruments on the spacecraft.
A problem discovered during (and not before) the flight hardware integration phase can cause significant cost and schedule impacts. The
testing successfully surfaced problems and led to their resolution before the full-up integration phase, saving significant cost and schedule
time. This approach could be used on future environmental satellite programs involving multiple, complex scientific instruments being
integrated onto a bus.
Autonomous Multi-Mission Virtual Ground and Spacecraft Operations
Avionics and Astrionics--Gudiance, Navigation, and Control
Company Name
Toyon Research Corp.
Title
A Miniature Ultra-Tightly Coupled Software-Defined
Navigation (SDN) System with Direction-Finding,
Attitude-Determination and Anti-Jam GPS
DOD SBIR Phase I
Quad Chart
Toyon Research Corporation proposes to develop a miniature ultra-tightly coupled software-defined navigation (SDN) system that can provide anti-jam
(AJ) GPS capability and attitude measurements independent of an inertial measurement unit (IMU). Using a single-aperture 4-cm antenna, the system
will provide angle-of-arrival (AoA) measurements for GPS jammers as well as signals outside the GPS band. These AoA measurements can assist
emitter localization algorithms that use time-difference-of-arrival (TDOA) and frequency-difference-of-arrival (FDOA) methods, and can be used for
threat identification and elimination. The SDN system will also accept a variety of IMU measurements for optimum AJ GPS and ultra-tightly coupled
GPS/inertial navigation system (INS) performance. These may include low-cost micro-electro-mechanical systems (MEMS), magnetic compass,
magnetometer, altimeter, time-of-arrival (TOA) inputs, and electro-optical (EO)/infrared (IR) sensors. The drift-free GPS-based attitude estimates will
augment measurements obtained by low-cost MEMS gyros with large rate biases, thereby improving the overall system performance. Moreover,
Toyon has teamed with Analytical Graphics, Inc. (AGI), the developer of the industry standard Navigation Tool Kit (NTK) and Satellite Tool Kit (STK)
software to analyze and verify the accuracy of the SDN system and sensor mixture specifically for TDOA and FDOA precision geolocation
applications. Furthermore, the SDN module will be developed as open-source software.
Autonomous Multi-Mission Virtual Ground and Spacecraft Operations
Avionics and Astrionics--Guidance, Navigation, and Control
NTRS
Abstract
This paper describes the architecture of the JPL multi-mission sequencing system and the sequence automation process, discusses the cost
savings associated with both of these changes.
The Deep Impact mission was ambitious and challenging. JPL's well proven, easily adaptable multi-mission sequence planning tools combined with
integrated spacecraft subsystem models enabled a small operations team to develop, validate, and execute extreme
FC
Jet Propulsion Laboratory
Title
Cost reduction from multi-mission sequencing software
Jet Propulsion Laboratory
Deep Impact Sequence Planning Using Multi-Mission
Adaptable Planning Tools With Integrated Spacecraft
Models
Jet Propulsion Laboratory
Jet Propulsion Laboratory
Multi-mission sequencing software
Proposed fully automated multi-mission uplink sequence
generation tool suite
Reduce costs with multimission sequencing and a
multimission operations system
Reduce costs with multi-mission sequencing and a multimission operations system
No Abstract Available
No Abstract Available
NASA (non Center Specific)
Scheduling Software for Complex Scenarios
Preparing a vehicle and its payload for a single launch is a complex process that involves thousands of operations. Because the equipment and
facilities required to carry out these operations are extremely expensive and limited in number, optimal assignment and efficient use are critically
important. Overlapping missions that compete for the same resources, ground rules, safety requirements, and the unique needs of processing
vehicles and payloads destined for space impose numerous constraints that, when combined, require advanced scheduling. Traditional scheduling
systems use simple algorithms and criteria when selecting activities and assigning resources and times to each activity. Schedules generated by
these simple decision rules are, however, frequently far from optimal. To resolve mission-critical scheduling issues and predict possible problem
areas, NASA historically relied upon expert human schedulers who used their judgment and experience to determine where things should happen,
whether they will happen on time, and whether the requested resources are truly necessary.
Jet Propulsion Laboratory
Science Opportunity Analyzer (SOA) a multi-mission approach No Abstract Available
to science planning
Jet Propulsion Laboratory
Jet Propulsion Laboratory
The paper will then propose extending this multi-mission philosophy to skeleton timeline development, science sequencing, and spacecraft
sequencing. Finally, the paper will investigate a multi-mission approach to MOS development.
Mission sequencing involves merging science and engineering inputs into an integrated, constraint-checked sequence and producing review and
spacecraft command products. This task employs processes, procedures, and tools which have a high degree of commonality across all missions.
The JPL Multi-Mission Office (MMO) Mission Planning and Sequencing Team (MPST) has successfully baselined these processes, procedures, and
tools so that they are readily adaptable to missions of varying complexity. As a result, the MPST can quickly assemble a team that provides mission
sequencing to very different missions at a fraction of previous costs. This paper will discuss the MMO MPST approach of adapting core processes,
procedures, and tools to multiple missions. The paper will then propose extending this multi-mission philosophy to skeleton timeline development,
science sequencing, and spacecraft sequencing. Finally, the paper will investigate a multi-mission approach to Mission Operations System (MOS)
development.
Autonomous Multi-Mission Virtual Ground and Spacecraft Operations
Avionics and Astrionics--Guidance, Navigation, and Control
FC
Goddard Space Flight Center;
Wallops Flight Facility
Title
A Space Based Internet Protocol System for Launch Vehicle
Tracking and Control
Jet Propulsion Laboratory
A Ten-Meter Ground-Station Telescope for Deep-Space
Optical Communications A Preliminary Design
Jet Propulsion Laboratory
Background and architecture for an autonomous ground
station controller
NTRS
Abstract
Personnel from the Goddard Space Flight Center Wallops Flight Facility (GSFC WFF) in Virginia are responsible for the overall management of the
NASA Sounding Rocket and Scientific Balloon Programs. Payloads are generally in support of NASA's Space Science Enterprise's missions and
return a variety of scientific data as well as providing a reasonably economical means of conducting engineering tests for instruments and devices
used on satellites and other spacecraft. Sounding rockets used by NASA can carry payloads of various weights to altitudes from 50 km to more than
1,300 km. Scientific balloons can carry a payload weighing as much as 3,630 Kg to an altitude of 42 km. Launch activities for both are conducted
not only from established ranges, but also from remote locations worldwide requiring mobile tracking and command equipment to be transported
and set up at considerable expense. The advent of low earth orbit (LEO) commercial communications satellites provides an opportunity to
dramatically reduce tracking and control costs of these launch vehicles and Unpiloted Aerial Vehicles (UAVs) by reducing or eliminating this ground
infrastructure. Additionally, since data transmission is by packetized Internet Protocol (IP), data can be received and commands initiated from
practically any location. A low cost Commercial Off The Shelf (COTS) system is currently under development for sounding rockets that also has
application to UAVs and scientific balloons. Due to relatively low data rate (9600 baud) currently available, the system will first be used to provide
GPS data for tracking and vehicle recovery. Range safety requirements for launch vehicles usually stipulate at least two independent tracking
sources. Most sounding rockets flown by NASA now carry GP receivers that output position data via the payload telemetry system to the ground
station. The Flight Modem can be configured as a completely separate link thereby eliminating the requirement for tracking radar. The system
architecture that integrates antennas, GPS receiver, commercial satellite packet data modem, and a single board computer with custom software is
described along with the technical challenges and the plan for their resolution. These include antenna development, high Doppler rates, reliability,
environmental ruggedness, hand over between satellites, and data security. An aggressive test plan is included which, in addition to environmental
testing, measures bit error rate, latency and antenna patterns. Actual launches on a sounding rocket and various aircraft flights have taken place.
Flight tests are planned for the near future on aircraft, long duration balloons and sounding rockets. These results, as well as the current status of
the project, are reported.
This article describes a telescope design for a 10-m optical ground station for deep-space communications. The design for a direct-detection optical
communications telescope differs dramatically from a telescope for imaging applications. In general, the requirements for optical manufacturing and
tracking performance are much less stringent for direct detection of optical signals. The technical challenge is providing a design that will operate in
the daytime nighttime conditions required for a Deep Space Network tracking application. The design presented addresses these requirements. The
design will provide higher performance at lower cost than existing designs.
The Deep Space Station Controller (DSSC) is state of the art ground station control architecture being developed at JPL. During the past few years
the technology development program at JPL demonstrated a series of increasingly competent automated ground station prototypes of which the
DSSC is the latest. It has been designed for robust closed loop control of ground stations utilized for forward and return link communications with
NASA's deep space exploration missions.
Autonomous Multi-Mission Virtual Ground and Spacecraft Operations
Avionics and Astrionics--Guidance, Navigation, and Control
FC
Marshall Space Flight Center
Title
Case for Deploying Complex Systems Utilizing Commodity
Components
Glenn Research Center
Inflatable Membrane Reflector and Shape-Memory Polymer
Antenna Developed for Space and Ground Communications
Application
Dryden Flight Research
Center
Internet Protocol Over Telemetry Testing for Earth Science
Capability Demo Summary
NTRS
Abstract
When the International Space Station (ISS) finally reached an operational state, many of the Payload Operations and Integration Facility (POIF)
hardware components were reaching end of life, COTS product costs were soaring, and the ISS budget was becoming severely constrained.
However, most requirement development was complete. In addition, the ISS program is a fully functioning program with at least fifteen years of
operational life remaining. Therefore it is critical that any upgrades, refurbishments, or enhancements be accomplished in realtime with minimal
disruptions to service. For these and other reasons, it was necessary to ensure the viability of the POIF. Due to the to the breadth of capability of the
POIF (a NASA ground station), it is believed that the lessons to be learned by other complex systems are applicable and any solutions garnered by
the POIF are applicable to other complex systems as well. With that in mind, a number of new approaches have been investigated to increase the
portability of the POIF and reduce the cost of refurbishment, operations, and maintenance. These new approaches were directed at the Total Cost
of Ownership (TCO) not only the refurbishment but also current operational difficulties, licensing, and anticipation of the next refurbishment. Our
basic premise is that technology had evolved dramatically since the concept of the POIF ground system and we should leverage our experience on
this new technological landscape. Fortunately, Moore's law and market forces have changed the landscape considerably. These changes are
manifest in five (5) ways that are particularly relevant to POIF 1. Complex Instruction Set Computing (CISC) processors have advanced to
unprecedented levels of compute capacity with a dramatic cost break, 2. Linux has become a major operating system supported by most vendors
on a broad range of platforms, 3. Windows(TradeMark) based desktops are pervasive in the office environment, 4. Stable and affordable
WindowsTM development environments and tools are available and offer a rich set of capabilities, 5. The WindowsTM 2000 provides a stable client
platform, Therefore, five studies were proposed, developed, and are in the current process of deployment which dramatically reduces the cost of
operations, maintenance, refurbishment, and deployment of a ground system. Restating and refining the basic premise stated earlier, it is possible
to enhance operations through the replacement of hardware and software components with commodity based items wherever applicable. This will
dramatically reduce the overall lifecycle cost of the project. The first study leveraged the POIF S secure, three-tier, web architecture to replace the
client workstations with lower cost PC platforms. A second study initiated a review of COTS products to examine the level of added value of each
product. This study included replacement of some COTS products with custom code, deletions, substitutions, and consolidation of COTS products.
Studies three and four reviewed the server architectures of the data distribution systems and Enhanced HOSC System (EHS) command and
telemetry system to propose migration to new platforms, both software and hardware. The final study reviewed current IP communication
technologies, developed an operational model for flight operations, and demonstrated that voice over IP was practical and could be integrated into
operations.
Communications requirements derived from the human and robotic exploration vision dictate a sophisticated, layered architecture that is
dramatically different than what was required for the low-rate, point-to-point communications of the Apollo era. Emerging, but relatively near-term,
exploration initiative needs include teleoperation and autonomous robotic missions, lunar reconnaissance and orbiting relay satellites, cooperative
spacecraft, lunar surface wireless local area networks, extremely wide bandwidth links to support hyperspectral imaging, synthetic aperture radar,
and other novel applications, such as high-definition television and telemedicine. For example, data rates on the order of 1 gigabit per second may
be required from Mars areostationary relay satellites. Such enormous data rates and extreme link distances will require 10-meter-class microwave
antennas, likely operating at Ka-band frequencies. The NASA Glenn Research Center has been involved in several efforts to meet these
requirements. This year, Glenn and SRS Technologies developed a 4- by 6-m offset-parabolic inflatable membrane reflector. Glenn, SRS, and the
Georgia Institute of Technology developed test apparatus and software to demonstrate that a novel ground station composed of an array of
relatively small apertures can economically replace a single, expensive tracking ground station. In addition, Johns Hopkins University delivered a
prototype 2-m version of an antenna based on a novel shape-memory composite structure.
The development and flight tests described here focused on utilizing existing pulse code modulation (PCM) telemetry equipment to enable onvehicle networks of instruments and computers to be a simple extension of the ground station network. This capability is envisioned as a necessary
component of a global range that supports test and development of manned and unmanned airborne vehicles.
Autonomous Multi-Mission Virtual Ground and Spacecraft Operations
Avionics and Astrionics--Guidance, Navigation, and Control
FC
Goddard Space Flight Center;
Jet Propulsion Laboratory
Title
James Webb Space Telescope Ka-Band Trade
Goddard Space Flight Center
LEO Download Capacity Analysis for a Network of Adaptive
Array Ground Stations
Jet Propulsion Laboratory;
Stennis Space Center
Vexcel Spells Excellence for Earth and Space
NTRS
Abstract
In August 2003 James Webb Space Telescope (JWST) had its Initial Review Confirmation Assessment Briefing with NASA HQ management. This
is a major milestone as the project was approved to proceed from Phase A to B, and NASA will commit funds for the project towards meeting its
science goals from the Earth-Sun s Lagrange 2 (L2) environment. At this briefing, the Project was asked, 'to take another look' into using, the JPL s
Deep Space Network (DSN) as the provider of ground stations and evaluate other ground station options. The current operations concept assumes
S-band and X-band communications with a daily and hour contact using the DSN with the goal of transmitting over 250 Gigabit (Gb) of data to the
ground. The Project has initiated a trade study to look at this activity, and we would like to share the result of the trade in the conference. Early
concept trades tends to focus on the 'normal' operation mode of supporting telemetry (science and engineering), command and radio metrics.
Entering the design phase, we find that we have the unique ranging requirement for our L2 orbit using alternating ground stations located in different
hemispheres. The trade must also address emergency operations (which are covered when using the DSN). This paper describes the issues
confronting this Project and how the DSN and the JWST Project are working together to find an optimized approach for meeting these issues. We
believe this trade is of major interest for future Code S and other L2 missions in that JWST will set the standard.
To lower costs and reduce latency, a network of adaptive array ground stations, distributed across the United States, is considered for the downlink
of a polar-orbiting low earth orbiting (LEO) satellite. Assuming the X-band 105 Mbps transmitter of NASA s Earth Observing 1 (EO-1) satellite with a
simple line-of-sight propagation model, the average daily download capacity in bits for a network of adaptive array ground stations is compared to
that of a single 11 m dish in Poker Flats, Alaska. Each adaptive array ground station is assumed to have multiple steerable antennas, either
mechanically steered dishes or phased arrays that are mechanically steered in azimuth and electronically steered in elevation. Phased array
technologies that are being developed for this application are the space-fed lens (SFL) and the reflectarray. Optimization of the different boresight
directions of the phased arrays within a ground station is shown to significantly increase capacity for example, this optimization quadruples the
capacity for a ground station with eight SFLs. Several networks comprising only two to three ground stations are shown to meet or exceed the
capacity of the big dish, Cutting the data rate by half, which saves modem costs and increases the coverage area of each ground station, is shown
to increase the average daily capacity of the network for some configurations.
With assistance from Stennis Space Center, Vexcel was able to strengthen the properties of its Apex Ground Station(TM), an affordable, end-to-end
system that comes complete with a tracking antenna that permits coverage within an approximate 2,000-kilometer radius of its location, a high
speed direct-to-disk data acquisition system that can download information from virtually any satellite, and data processing software for virtually all
synthetic aperture radar and optical satellite sensors. Vexcel is using an Apex system linked to the Terra satellite to help scientists and NASA
personnel measure land and ocean surface temperatures, detect fires, monitor ocean color and currents, produce global vegetation maps and data,
and assess cloud characteristics and aerosol concentrations. In addition, Vexcel is providing NASA with close-range photogrammetry software for
the International Space Station. The technology, commercially available as FotoG(TM), was developed with SBIR funding and support from NASA's
Jet Propulsion Laboratory. Commercially, FotoG is used for demanding projects taken on by engineering firms, nuclear power plants, oil refineries,
and process facilities. A version of Vexcel's close-range photo measurement system was also used to create virtual 3-D backdrops for a high-tech
science fiction film.
Autonomous Multi-Mission Virtual Ground and Spacecraft Operations
Communications--Architectures and Networks
Company Name
James Cutler
Title
Software-Defined Ground StationsEnhancing Multi-Mission Support
Field Center
ARC
NASA SBIR Phase I
Quad Chart
This SBIR Phase 1 proposal to NASA requests $99,055.69 to enhance multiple mission support in ground stations through the use of software
defined radios and virtual machines to create a software-defined ground station (SDGS). This proposal responds to NASA SBIR topic S4.09
"Autonomous Multi-Mission Virtual Ground and Spacecraft Operations". The significance of our SDGS work is that the monolithic, stovepipe, and
hardware centric nature of ground stations will be reduced. Major system components will be moved to software, thereby promoting remote,
network-based maintenance, upgrades, and new technology development. Off the shelf software modules will be available, but also mechanism
for low-level ground station customization for mission specifics; all done remotely over the Internet. Costly hardware upgrades will be reduced or
eliminated. Our innovation is in the intelligent combination of software-defined radio techniques and virtual machines. This enables a near
complete software solution to primary ground station functions. It simplifies ground station hardware and enables flexible application support. In
Phase 1 we propose to architect an SDGS system for support of expected small satellite missions. We will prototype basic elements with an on
orbit or engineering model satellite system from our partners. Commercial applications
include communication support for satellites and high altitude balloon systems. Our customers will include NASA, NSF, DoD, and private satellite
builders such as universities and venture space. The PI, Dr. James Cutler, has extensive small satellite and ground station experience. He has
prototype a global ground station network to operate satellites as if they were nodes on the Internet. Our facilities located in Northern California
have tools for computer and radio development, and access to small satellite systems and ground station resources.
Autonomous Multi-Mission Virtual Ground and Spacecraft Operations
Communications--Architectures and Networks
NASA SBIR Phase II
Quad Chart
From a functional perspective, the BioNet middleware provides a standards-based command and control capability, while the Mobile Data
Acquistion and Communications System (MDACS) provides a generic space-rated hardware platform for general-purpose computations
and communications.
Company Name
Invocon, Inc. and University of
Colorado
Title
INTEGRATED DATA ASSIMILATION
ARCHITECTURE
Field Center
GRC
Summation Research, Inc.
PROGRAMMABLE HIGH-RATE MULTIMISSION RECEIVER FOR SPACE
COMMUNICATIONS
GSFC
A modular and flexible High Rate Receiver Backbone (HRRB) would allow customization of some processing firmware and should
accommodate advances in deployed link formats more easily than units “factory loaded” for particular signal types.
Toyon Research Corporation
MULTI-MISSION SDR
GRC
Specifically, we propose to develop a modular, but highly integrated, digital and analog signal processing platform along with a standardscompliant waveform. In Phase II Toyon will pursue development of a waveform that is standards-based in order to further promote reuse
and interoperability.
Autonomous Multi-Mission Virtual Ground and Spacecraft Operations
Communications--Architectures and Networks
Company Name
Innoflight, Inc. and Southwest Research
Institute
Title
TARGETS FOR RADAR CALIBRATION AND TEST OF
ADVANCED DISCRIMINATION TECHNOLOGIES AND
CONCEPTS
DOD SBIR Phase II
Quad Chart
This Phase-II project will build and demonstrate a sophisticated Microsat Encryption Unit (MEU) system that will meet the needs of
DoD missions. It will be designed to mechanical and electrical specifications that meet the small satellite market needs. In addition,
it will enable advanced TCP/IP communications in compliance with the DoD's Transformational Communications concept of
operations.
Autonomous Multi-Mission Virtual Ground and Spacecraft Operations
Communications--Architectures and Networks
U.S. Patent Applications
Patent Number
Title
Assignee
Abstract
US2003046336A1
Persistent link for broadband mobile
platform communications
systemsUSing proxy servers
None
A communications system according to the invention provides a communications link between a distributed communications system and a
mobile platform via a satellite. The communications system includes a ground station and a parent proxy server connected to the ground
station. A distributed communications system is connected to the parent proxy server. A satellite communicates with the ground station. A
transceiver, a router and a child proxy server are located on the mobile platform. A user communication device (UCD) is connected to the child
proxy server. The child and parent proxy servers establish a persistent transmission control protocol (TCP) link between the mobile platform
and the ground station. The UCD connects to the child proxy server using a first group of TCP settings. The child and parent proxy servers
communicate using a second group of TCP settings to optimize the TCP link for long delay satellite links.
US2003220106A1
Method and system for object
tracking and communication
None
A system and method for communicating between ground stations and mobile satellites effectively eliminates the need for transferring
information from one ground station to another. Communication between a ground station and a mobile satellite is accomplished via an
intermediate network of geosynchronous control satellites. The mobile satellite is positioned within a three-dimensional cellular grid formed by
the network of geosynchronous satellites. This system is particularly applicable to a terrestrial network communicatively coupled to the
geosynchronous network via one or more ground stations. The terrestrial network of ground stations no longer needs to perform expensive,
time-consuming and manual synchronization and hand-off procedures. Instead, the position of a ground station antenna can remain fixed on
one specific satellite in the geosynchronous network, while the mobile satellite is tracked by the geosynchronous network utilizing the threedimensional cellular grid formed by intermediate control satellites.
US2006094351A1
System and method of providing Ntiered enterprise/web-based
management, procedure
coordination, and control of a
geosynchronous satellite fleet
SES Americom,
Inc.
An N-Tiered enterprise and/or N-Tiered Web-based enterprise system and methods for managing, monitoring, coordinating the procedures of
and optionally controlling a fleet of geosynchronous satellites through a distributed network such as the Internet. The system and associated
object oriented software architecture seamlessly supports monitoring and analyzing real-time, near real-time, historical/playback, simulated and
archived satellite telemetry as well as Tracking, Telemetry and Control (TT&C) ground system generated data products such as real-time alerts,
archived alerts and ground system server statuses from a group of legacy distributed TT&C wide area network (WAN) ground station systems
located around the world.
Autonomous Multi-Mission Virtual Ground and Spacecraft Operations
Communications--Architectures and Networks
USPTO
Patent Number
Title
Assignee
Abstract
US6411806
Virtual network
configuration and
management system for
satellite communications
system
Mobile Satellite
Ventures LP
US6466569
Uplink transmission and
reception techniques for
a processing
satelliteation satellite
TRW Inc.
US6553208
Methods and apparatus
Motorola, Inc.
forUSing ground based
processing resources to
support orbiting satellites
US6625129
Demand assigned spatial
multiplexing in satellite
communication systems
In a mobile satellite system, a system for providing satellite communication between multiple users in a virtual network arrangement includes first and second
mobile earth terminals (METs) responsively connected to and registering with the mobile satellite system. The first MET selects a virtual network identifier
(VN ID) representing a virtual network group including the first and second METs to establish voice communication therewith and transmits the VN ID to a
central controller. The central controller receives the VN ID from the first MET, validates the first MET for communication, validates the VN ID, allocates a
frequency for the virtual network group, and broadcasts the message to the virtual network group including the second MET informing the virtual network
group of the allocated frequency and the voice communication associated therewith. The second MET tunes to the frequency in response to the message
broadcast by the central controller.
Uplink transmission and reception techniques for a processing satellite including one or more earth terminals 400 connected to receive ATM data cells. One or
more encoders 418 are connected to coordinate four data cells with an error correction code to generate data bursts and to coordinate the data bursts with
synchronizing bursts to generate data frames. One or more modulators 420 are connected to modulate the data frames by frequency division multiple access
modulation to enable placement of the modulated data frames into a plurality of channels. One or more antennas 406 transmit the modulated data frames to a
satellite 100 over 48 beams with various forms of polarization. In satellite 100, a receiving multibeam antenna and feed 106 responds to one or more beams of
radiocarrier signals having one or more forms of polarization. One or more demodulators 138 demodulate the radio carrier signals into data frames from various
channels including a plurality of channel types. One or more decoders 146 decode data bursts and synchronizing bursts from the data frames and decode four
data cells from the bursts using error correction code.
A plurality of orbiting satellites (102-112), each comprised of an onboard computer (142). Each of the orbiting satellites and a plurality of ground-based
computers (114-124) are utilized to perform delay-insensitive functions associated with the orbiting satellites. An ubiquitous data link (128-138) is maintained
between each satellite and one or more of the ground-based computers. The integrated system is configured to allow a satellite to generate a remote procedure
call (RPC), transmit the RPC to a ground-based computer, whereupon the ground-based computer processes the RPC and transmits a response back to the
satellite. Delay-sensitive tasks and functions are performed by the computer onboard the satellite; delay-insensitive tasks are advantageously performed by the
ground-based computer assets.
The methods (300, 900, and 1200), and apparatus (140 and 800) of the present invention employ demand assigned spatial multiplexing in a satellite
communication system (100) to efficiently accommodate varying demand for high data rate service while minimizing satellite antenna and radio system
complexity in a ubiquitous coverage satellite system optimized for transporting packet data. In a preferred embodiment, the methods (300, 900, 1000 and 1200)
and apparatus (140 and 800) of the present invention use a first set of signaling beams (210) to establish signaling uplinks and then direct traffic spot beams
(220) at particular UE devices to support traffic uplinks in accordance with demand assigned multiple access (DAMA) parameters when and only when the UE
devices have traffic data to transmit to a satellite. In a preferred embodiment, UE device (140) also executes a method (1200) that includes steps for determining
appropriate steering angles for a servicing satellite (110) to use to direct a traffic spot beam (220) to the UE device's location when the UE has traffic data to
transmit.
Motorola, Inc.
Autonomous Multi-Mission Virtual Ground and Spacecraft Operations
Communications--Architectures and Networks
USPTO
Patent Number
Title
Assignee
Abstract
US6678520
Method and apparatus
for providing wideband
servicesUSing medium
and low earth orbit
satellites
Hughes
Electronics
Corporation
A method and apparatus for mitigating communications interference between satellite communications systems in different orbits is disclosed. The method
comprises the steps of evaluating a geometrical relationship between a second ground station and the satellites in the second satellite constellation, and directing
communications between the second ground station and the second satellite according to the evaluated geometrical relationship. In one embodiment
communications are handed over from a first satellite to another satellite when the first satellite is no longer at the highest elevation angle of visible satellites. In
another embodiment, handover occurs when the first satellite drops below a minimum elevation angle.
US6690934
Multiple access satellite
communications network
US6823170
Satellite communications
systemUSing multiple
earth stations
US6931247
Aircraft control method
Universal Space A shared single system of ground stations (120) for servicing any number of satellite owners, permits the owners to transfer command information to their
Network, Inc.
satellite (110), and collect data streams that arm sent back from the satellite all via a standardized global communication system maintained and operated by a
commercial satellite communications company. A remotely controlled ground station (120) is operated and controlled from a central controller. Each user
creates and stores a ground station configuration file at the central controller, which file contains the data necessary to configure the remotely controlled ground
station (120) to communicate with the user's satellite (110). When the user desires to communicate with the user's satellite, the user schedules a communication
session with the central controller, which downloads the configuration file to the appropriate ground station. The appropriate ground station is determined based
on current orbital characteristics of the satellite in question. A server at the ground station then uses the data in the configuration file to configure equipment at
the ground station to communicate with the desired satellite.
Ericsson Inc.
A satellite communications system uses multiple ground stations and one or more satellites for communicating between mobile subscribers and a land-based
communications network, such as the PSTN or the Internet. Multiple ground stations geographically dispersed minimize toll charges incurred routing calls from
a mobile subscriber through the land network by reducing the need for long-distance calling. Further, because each ground station communicates with a given
satellite using the same frequency spectrum, the subscriber capacity of the system increases and/or bandwidth requirements for the communications link
between ground stations and satellites may be reduced. The present system uses ground-based beamforming techniques enabling each satellite to transmit
signals in multiple transmission beams, each beam supporting one or more mobile subscribers. Each beam may reuse the same frequency spectrum, thereby
increasing the number of subscribers supported by each satellite. Multiple ground stations cooperatively relay signals through a given satellite in a manner
complementary with ground-based beamforming.
Aerovironment, This disclosure provides a solar rechargeable aircraft that is inexpensive to produce, is steerable, and can remain airborne almost indefinitely. The preferred
Inc.
aircraft is a span-loaded flying wing, having no fuselage or rudder. Traveling at relatively slow speeds, and having a two-hundred foot wingspan that mounts
photovoltaic cells on most all of the wing's top surface, the aircraft uses only differential thrust of its eight propellers to turn. Each of five segments of the wing
has one or more motors and photovoltaic arrays, and produces its own lift independent of the other segments, to avoid loading them. Five two-sided
photovoltaic arrays, in all, are mounted on the wing, and receive photovoltaic energy both incident on top of the wing, and which is incident also from below,
through a bottom, transparent surface. The aircraft includes hinges and actuators capable of providing an adjustable dihedral for the wing. The actuators can be
motors or control surfaces. Alternately, the actuators can be movable masses within the wing, which may be capable of deforming the wing to alter the
aerodynamics of the wing, and thereby actuate the hinges. Because of wing dihedral, the aircraft includes motors both above and below the center of drag, and
the aircraft uses differential thrust to control aircraft pitch. The aircraft has a wide variety of applications, which include serving as a long term high altitude
platform that serves to link a ground station using radio wave signals and a satellite using optical signals.
Autonomous Multi-Mission Virtual Ground and Spacecraft Operations
Communications--Architectures and Networks
USPTO
Patent Number
Title
Assignee
Abstract
US6934609
Space-based integrated
multi-mission broadband
architecture
Lockheed
Martin
Corporation
US6950625
Communications
apparatus and method
ICO Services
Limited
Spacecraft network and communication method thereof. A spacecraft network includes a first server spacecraft disposed in a first server orbit, a first client
spacecraft disposed in a first client orbit, and a wireless local area network formed between at least the first server spacecraft and the first client spacecraft. The
wireless local area network includes at least one communication channel to transmit and receive spatial information, at least one receiver to receive a first
communication signal including at least routing information, at least one routing system to determine a desired route, and one transmitter to transmit the first
communication signal. The first client spacecraft is free from the at least one routing system, and the first server spacecraft includes one of the at least one
routing system.
A method of reusing frequency bands between base stations of a terrestrial mobile communications network and a satellite communications network,
comprising allocating the frequency bands using integrated resource management and other mitigation techniques in a such a way as to minimize interference
between both the systems, thus making optimum usage of valuable frequency spectrum.
US6952580
Multiple link internet
protocol mobile
communications system
and method therefor
Mobile communications
networkUSing point-topoint protocol over
ethernet
The DIRECTV
Group, Inc.
A communication system has a plurality of high altitude devices that are coupled to user terminals through a plurality of dynamic links. The terminal monitors
and changes the multiple dynamic links as the position of the user terminal relative to the high altitude devices changes. The terminal transmits to and receives
from the user terminals through the high altitude devices using a plurality of datagrams.
The Boeing
Company
US7068615
Adaptable forward link
data rates in
communications systems
for mobile platforms
The Boeing
Company
US7103280
Architecture for an
optical satellite
communication network
The DirecTV
Group, Inc.
A communications system that provides broadband access to passengers of mobile platforms includes a router located on the mobile platform. A network is
connected to the router. User communication devices (UCDs) connected to the network, wherein the UCDs establish point-to-point over Ethernet (PPPoE)
sessions with the router. A transmitter and a receiver are connected to the router. A satellite and a ground station are in communication with the transmitter and
the receiver. A distributed communications system includes virtual private networks (VPN) and is connected to the ground station. A first address manager
leases the use of public IP addresses by the mobile platform. A second address manager assigns the public IP addresses to UCDs when the UCDs request access
to the VPNs and private IP addresses for other network service. The UCDs employ IPSec protocol when accessing the VPNs.
A communication system for mobile platforms includes mobile platforms with transceivers identified by Internet Protocol (IP) addresses. A satellite relays a
forward link from a ground station to the mobile platforms. The forward link contains IP packet data that is modulated by variable length orthogonal (VLO)
spreading codes and that has different information data rates. The VLO spreading code for each IP packet is selected to optimize a desired link margin of the IP
packet that is received by the addressed transceiver. The IP packets can also be modulated using a pseudonoise (PN) spreading code. Forward error correction
(FEC) may also be applied. The transceivers include a feedback circuit that generates an Eb/No estimate.
A satellite constellation has a plurality of satellites. Each of the satellites has an RF ground link for communicating with a ground station and an optical link for
communication with at least one of the plurality of satellites. Each of the satellites has a reconfigurable optical transmitter for sending and receiving data
streams. Each reconfigurable optical transmitter has a first optical carrier associated therewith and a reconfigurable optical receiver. The plurality of satellites is
arranged to have a first subset of satellites. The first subset of satellites is configured to communicate. The plurality of satellites is reconfigured to have a second
subset of satellites having at least one different satellites than that of said first subset. The second subset supercedes the first subset. The second subset of
satellites is configured to communicate. Various subset around the globe may form local area networks. The local area networks are preferably optically coupled
to form a wide area network.
US7054322
Autonomous Multi-Mission Virtual Ground and Spacecraft Operations
Communications--Architectures and Networks
USPTO
Patent Number
Title
Assignee
Abstract
US7151929
Satellite payload data
communications and
processing techniques
Northrop
Grumman
Corporation
A satellite communication system ( 10 ) includes a satellite ( 20 ) which receives first and second data from sources. Each satellite receives control data from a
satellite control center ( 100 ). An earth processing center (PC) is arranged to process the data received from the satellites, and a wide band network ( 30 ) is
arranged to transmit the data to the processing center. A receptor terminal (A) is arranged to receive the first data and to place the first data on the network ( 30 )
for transmission to the processing center (PC). A second receptor terminal (B) is arranged to receive the second data and to place the second data on the network
( 30 ) for transmission to the processing center (PC).
Autonomous Multi-Mission Virtual Ground and Spacecraft Operations
Communications--Architectures and Networks
FC
Jet Propulsion Laboratory
Title
A Multi-mission Event-Driven
Component-Based System for
Support of Flight Software
Development, ATLO, and
Operations first used by the Mars
Science Laboratory (MSL) Project
GRC
Advanced Communications
Architecture Demonstration Made
Significant Progress
Glenn Research Center
Command and Control of Space
Assets Through Internet-Based
Technologies Demonstrated
Jet Propulsion Laboratory
Converging voice and data over
mission-critical networks
Jet Propulsion Laboratory
Design and implementation of an
inter-agency, multi-mission space
flight operations network interface
NTRS
Abstract
This paper details an architectural description of the Mission Data Processing and Control System (MPCS), an event-driven, multi-mission ground data processing
components providing uplink, downlink, and data management capabilities which will support the Mars Science Laboratory (MSL) project as its first target mission. MPCS is
developed based on a set of small reusable components, implemented in Java, each designed with a specific function and well-defined interfaces. An industry standard
messaging bus is used to transfer information among system components. Components generate standard messages which are used to capture system information, as well
as triggers to support the event-driven architecture of the system. Event-driven systems are highly desirable for processing high-rate telemetry (science and engineering)
data, and for supporting automation for many mission operations processes.
Simulation for a ground station located at 44.5 deg latitude. The Advanced Communications Architecture Demonstration (ACAD) is a concept architecture to provide highrate Ka-band (27-GHz) direct-to-ground delivery of payload data from the International Space Station. This new concept in delivering data from the space station targets
scientific experiments that buffer data onboard. The concept design provides a method to augment the current downlink capability through the Tracking Data Relay Satellite
System (TDRSS) Ku-band (15-GHz) communications system. The ACAD concept pushes the limits of technology in high-rate data communications for space-qualified
systems. Research activities are ongoing in examining the various aspects of high-rate communications systems including (1) link budget parametric analyses, (2) antenna
configuration trade studies, (3) orbital simulations (see the preceding figure), (4) optimization of ground station contact time (see the following graph), (5) processor and
storage architecture definition, and (6) protocol evaluations and dependencies.
The NASA Glenn Research Center successfully demonstrated a transmission-control-protocol Internet-protocol- (TCP IP) based approach to the command and control of
onorbit assets over a secure network. This is a significant accomplishment because future NASA missions will benefit by using Internet-standards-based protocols. Benefits
of this Internet-based space command and control system architecture include reduced mission costs and increased mission efficiency. The demonstration proved that this
communications architecture is viable for future NASA missions. This demonstration was a significant feat involving multiple NASA organizations and industry. Phillip
Paulsen, from Glenn's Project Development and Integration Office, served as the overall project lead, and David Foltz, from Glenn's Satellite Networks and Architectures
Branch, provided the hybrid networking support for the required Internet connections. The goal was to build a network that would emulate a connection between a space
experiment on the International Space Station and a researcher accessing the experiment from anywhere on the Internet, as shown. The experiment was interfaced to a
wireless 802.11 network inside the demonstration area. The wireless link provided connectivity to the Tracking and Data Relay Satellite System (TDRSS) Internet Link
Terminal (TILT) satellite uplink terminal located 300 ft away in a parking lot on top of a panel van. TILT provided a crucial link in this demonstration. Leslie Ambrose, NASA
Goddard Space Flight Center, provided the TILT TDRSS support. The TILT unit transmitted the signal to TDRS 6 and was received at the White Sands Second TDRSS
Ground Station. This station provided the gateway to the Internet. Coordination also took place at the White Sands station to install a Veridian Firewall and automated
security incident measurement (ASIM) system to the Second TDRSS Ground Station Internet gateway. The firewall provides a trusted network for the simulated space
experiment. A second Internet connection at the demonstration area was implemented to provide Internet connectivity to a group of workstations to serve as platforms for
controlling the simulated space experiment. Installation of this Internet connection was coordinated with an Internet service provider (ISP) and local NASA Johnson Space
Center personnel. Not only did this TCP IP-based architecture prove that a principal investigator on the Internet can securely command and control on-orbit assets, it also
demonstrated that valuable virtual testing of planned on-orbit activities can be conducted over the Internet prior to actual deployment in space.
The U.S. National Aeronautics and Space Administration (NASA) Deep Space Network DSN - is an international network of antennas that supports interplanetary spacecraft
missions and radio and radar astronomy observations for the exploration of the solar system and the universe. The network also supports selected Earth-orbiting missions.
This paper describes the ground communications network of the DSN and ways network infrastructure costs are being reduced by the introduction of new technology.
Fundamentally, the DSN ground network architecture is a star network, and the hub is at JPL in Pasadena, California. Communications to customer sites are designed to
minimize NASA costs and may be either shared IP backbone networks or dedicated circuits. One of the primary features of the network is its ability to support real-time data,
voice, and video communications among antenna stations, an automated multi-mission operations systems facility at JPL (AMMOS), and mission operations centers
(MOCs) at NASA and non-NASA facilities. Funding for advanced engineering to reduce ground network infrastructure costs was provided by the NASA Office of Space
Science.
An advanced network interface was designed and implemented by a team from the Jet Propulsion Lab with support from the European Space Operations Center. This
poster shows the requirements for the interface, the design, the topology, the testing and lessons learned from the whole implementation.
Autonomous Multi-Mission Virtual Ground and Spacecraft Operations
Communications--Architectures and Networks
FC
Jet Propulsion Laboratory
Goddard Space Flight Center
Title
Enabling Space Science and
Exploration
Ethernet for Space Flight
Applications
Goddard Space Flight Center
Experimenting with an Evolving
Ground Space-based Software
Architecture to Enable Sensor
Webs
Goddard Space Flight Center
Internet Access to Spacecraft
Johnson Space Center
NASA's Core Trajectory SubSystem Project Using JBoss
Enterprise Middleware for
Building Software Systems Used
to Support Spacecraft Trajectory
Operations
NTRS
Abstract
This viewgraph presentation on enabling space science and exploration covers the following topics 1) Today s Deep Space Network 2) Next Generation Deep Space
Network 3) Needed technologies 4) Mission IT and networking and 5) Multi-mission operations.
NASA's Goddard Space Flight Center (GSFC) is adapting current data networking technologies to fly on future spaceflight missions. The benefits of using commercially
based networking standards and protocols have been widely discussed and are expected to include reduction in overall mission cost, shortened integration and test (I and
ampT) schedules, increased operations flexibility, and hardware and software upgradeability scalability with developments ongoing in the commercial world. The networking
effort is a comprehensive one encompassing missions ranging from small University Explorer (UNEX) class spacecraft to large observatories such as the Next Generation
Space Telescope (NGST). Mission aspects such as flight hardware and software, ground station hardware and software, operations, RF communications, and security
(physical and electronic) are all being addressed to ensure a complete end-to-end system solution. One of the current networking development efforts at GSFC is the
SpaceLAN (Spacecraft Local Area Network) project, development of a space-qualifiable Ethernet network. To this end we have purchased an IEEE 802.3-compatible 10
100 1000 Media Access Control (MAC) layer Intellectual Property (IP) core and are designing a network node interface (NNI) and associated network components such as a
switch. These systems will ultimately allow the replacement of the typical MIL-STD-1553 1773 and custom interfaces that inhabit most spacecraft. In this paper we will
describe our current Ethernet NNI development along with a novel new space qualified physical layer that will be used in place of the standard interfaces. We will outline our
plans for development of space qualified network components that will allow future spacecraft to operate in significant radiation environments while using a single onboard
network for reliable commanding and data transfer. There will be a brief discussion of some issues surrounding system implications of a flight Ethernet. Finally, we will show
an onboard network architecture for a proposed new mission using Ethernet for science data transport.
A series of ongoing experiments are being conducted at the NASA Goddard Space Flight Center to explore integrated ground and space-based software architectures
enabling sensor webs. A sensor web, as defined by Steve Talabac at NASA Goddard Space Flight Center(GSFC), is a coherent set of distributed nodes interconnected by a
communications fabric, that collectively behave as a single, dynamically adaptive, observing system. The nodes can be comprised of satellites, ground instruments,
computing nodes etc. Sensor web capability requires autonomous management of constellation resources. This becomes progressively more important as more and more
satellites share resource, such as communication channels and ground station,s while automatically coordinating their activities. There have been five ongoing activities
which include an effort to standardize a set of middleware. This paper will describe one set of activities using the Earth Observing 1 satellite, which used a variety of ground
and flight software along with other satellites and ground sensors to prototype a sensor web. This activity allowed us to explore where the difficulties that occur in the
assembly of sensor webs given today s technology. We will present an overview of the software system architecture, some key experiments and lessons learned to facilitate
better sensor webs in the future.
The Operating Missions as Nodes on the Internet (OMNI) project at NASA's Goddard Space flight Center (GSFC), is demonstrating the use of standard Internet protocols for
spacecraft communication systems. This year, demonstrations of Internet access to a flying spacecraft have been performed with the UoSAT-12 spacecraft owned and
operated by Surrey Satellite Technology Ltd. (SSTL). Previously, demonstrations were performed using a ground satellite simulator and NASA's Tracking and Data Relay
Satellite System (TDRSS). These activities are part of NASA's Space Operations Management Office (SOMO) Technology Program, The work is focused on defining the
communication architecture for future NASA missions to support both NASA's 'faster, better, cheaper' concept and to enable new types of collaborative science. The use of
standard Internet communication technology for spacecraft simplifies design, supports initial integration and test across an IP based network, and enables direct
communication between scientists and instruments as well as between different spacecraft, The most recent demonstrations consisted of uploading an Internet Protocol (IP)
software stack to the UoSAT- 12 spacecraft, simple modifications to the SSTL ground station, and a series of tests to measure performance of various Internet applications.
The spacecraft was reconfigured on orbit at very low cost. The total period between concept and the first tests was only 3 months. The tests included basic network
connectivity (PING), automated clock synchronization (NTP), and reliable file transfers (FTP). Future tests are planned to include additional protocols such as Mobile IP, email, and virtual private networks (VPN) to enable automated, operational spacecraft communication networks. The work performed and results of the initial phase of tests
are summarized in this paper. This work is funded and directed by NASA GSFC with technical leadership by CSC in arrangement with SSTL, and Vytek Wireless.
NASA's Johnson Space Center (JSC) Flight Design and Dynamics Division (DM) has prototyped the use of Open Source middleware technology for building its next
generation spacecraft mission support system. This is part of a larger initiative to use open standards and open source software as building blocks for future mission and
safety critical systems. JSC is hoping to leverage standardized enterprise architectures, such as Java EE, so that its internal software development efforts can be focused
on the core aspects of their problem domain. This presentation will outline the design and implementation of the Trajectory system and the lessons learned during the
exercise.
Autonomous Multi-Mission Virtual Ground and Spacecraft Operations
Communications--Architectures and Networks
NTRS
Abstract
In this work, we present sensing performance using an architecture for a reconfigurable protocol chip for spacebased applications. Toward utilizing the IP packet
architecture, utilizing data link layer framing structures for multiplexed data on a channel are the targeted application considered for demonstration purposes. Specifically, we
examine three common framing standards and present the sensing performance of these standards and their relative de-correlation metrics. Some analysis is performed to
investigate the impact of lossy links and on the number of packets required to perform a decision with some probability. Finally, we present results on a demonstration
platform that integrated reconfigurable sensing technology into the Ground Station Interface Device (GRID) for End-to-End IP demonstrations in space.
FC
Jet Propulsion Laboratory
Title
Protocol Sensing Across Multiple
Space Missions
Jet Propulsion Laboratory
Reconfigurable Protocol Sensing
in an End-to-End Demonstration
In this work, we present sensing performance using an architecture for a reconfigurable protocol chip for spacebased applications. Toward utilizing the IP packet
architecture, utilizing data link layer framing structures for multiplexed data on a channel are the targeted application considered for demonstration purposes. Specifically, we
examine three common framing standards and present the sensing performance of these standards and their relative de-correlation metrics. Some analysis is performed to
investigate the impact of lossy links. Finally, we present results on a demonstration platform that integrated reconfigurable sensing technology into the Ground Station
Interface Device (GRID) for End-to-End IP demonstrations in space.
Glenn Research Center
Secure, Network-Centric
Operations of a Space-Based
Asset Cisco Router in Low Earth
Orbit (CLEO) and Virtual Mission
Operations Center (VMOC)
This report documents the design of network infrastructure to support operations demonstrating the concept of network-centric operations and command and control of
space-based assets. These demonstrations showcase major elements of the Transformal Communication Architecture (TCA), using Internet Protocol (IP) technology. These
demonstrations also rely on IP technology to perform the functions outlined in the Consultative Committee for Space Data Systems (CCSDS) Space Link Extension (SLE)
document. A key element of these demonstrations was the ability to securely use networks and infrastructure owned and or controlled by various parties. This is a sanitized
technical report for public release. There is a companion report available to a limited audience. The companion report contains detailed networking addresses and other
sensitive material and is available directly from William Ivancic at Glenn Research Center.
Glenn Research Center
Shuttle Payload Ground
Command and Control An
Experiment Implementation
Combustion Module-2 Software
Development, STS-107
This presentation covers the design of a command and control architecture developed by the author for the Combustion Module-2 microgravity experiment, which flew
aboard the STS-107 Shuttle mission, The design was implemented to satisfy a hybrid network that utilized TCP IP for both the onboard segment and ground segment, with
an intermediary unreliable transport for the space to ground segment. With the infusion of Internet networking technologies into Space Shuttle, Space Station, and
spacecraft avionics systems, comes the need for robust methodologies for ground command and control. Considerations of high bit error links, and unreliable transport over
intermittent links must be considered in such systems. Internet protocols applied to these systems, coupled with the appropriate application layer protections, can provide
adequate communication architectures for command and control. However, there are inherent limitations and additional complexities added by the use of Internet protocols
that must be considered during the design. This presentation will discuss the rationale for the framework and protocol algorithms developed by the author. A summary of
design considerations, implantation issues, and learned lessons will be will be presented. A summary of mission results using this communications architecture will be
presented. Additionally, areas of further needed investigation will be identified.
Johnson Space Center
Sustainable, Reliable MissionSystems Architecture
A mission-systems architecture, based on a highly modular infrastructure utilizing open-standards hardware and software interfaces as the enabling technology is essential
for affordable md sustainable space exploration programs. This mission-systems architecture requires (8) robust communication between heterogeneous systems, (b) high
reliability, (c) minimal mission-to-mission reconfiguration, (d) affordable development, system integration, end verification of systems, and (e) minimal sustaining engineering.
This paper proposes such an architecture. Lessons learned from the Space Shuttle program and Earthbound complex engineered systems are applied to define the model.
Technology projections reaching out 5 years are made to refine model details.
Autonomous Multi-Mission Virtual Ground and Spacecraft Operations
Communications--Architectures and Networks
NTRS
Abstract
Project Constellation implements NASA's vision for space exploration to expand human presence in our solar system. The engineering focus of this project is developing a
system of systems architecture. This architecture allows for the incremental development of the overall program. Systems can be built and connected in a "Lego style"
manner to generate configurations supporting various mission objectives. The development of the avionics or control systems of such a massive project will result in
concurrent engineering. Also, each system will have software and the need to communicate with other (possibly heterogeneous) systems. Fortunately, this design problem
has already been solved during the creation and evolution of systems such as the Internet and the Department of Defense's successful effort to standardize distributed
simulation (now IEEE 1516). The solution relies on the use of a standard layered software framework and a communication protocol. A standard framework and
communication protocol is suggested for the development and maintenance of Project Constellation systems. The ARINC 653 standard is a great start for such a common
software framework. This paper proposes a common system software framework that uses the Real Time Publish Subscribe protocol for framework-to-framework
communication to extend ARINC 653. It is highly recommended that such a framework be established before development. This is important for the success of concurrent
engineering. The framework provides an infrastructure for general system services and is designed for flexibility to support a spiral development effort.
FC
Johnson Space Center
Title
System Software Framework for
System of Systems Avionics
Goddard Space Flight Center
The Center for Space
Telemetering and
Telecommunications Systems
This report comprises the final technical report for the research grant 'Center for Space Telemetering and Telecommunications Systems' sponsored by the National
Aeronautics and Space Administration's Goddard Space Flight Center. The grant activities are broken down into the following technology areas (1) Space Protocol Testing
(2) Autonomous Reconfiguration of Ground Station Receivers (3) Satellite Cluster Communications and (4) Bandwidth Efficient Modulation. The grant activity produced a
number of technical reports and papers that were communicated to NASA as they were generated. This final report contains the final summary papers or final technical
report conclusions for each of the project areas. Additionally, the grant supported students who made progress towards their degrees while working on the research.
Goddard Space Flight Center
The Earth Based Ground Stations
Element of the Lunar Program
The Lunar Architecture Team (LAT) is responsible for developing a concept for building and supporting a lunar outpost with several exploration capabilities such as rovers,
colonization, and observatories. The lunar outpost is planned to be located at the Moon's South Pole. The LAT Communications and Navigation Team (C and amp;N) is
responsible for defining the network infrastructure to support the lunar outpost. The following elements are needed to support lunar outpost activities: A Lunar surface
network based on industry standard wireless 802.xx protocols, relay satellites positioned 180 degrees apart to provide South Pole coverage for the half of the lunar 28-day
orbit that is obscured from Earth view, earth-based ground stations deployed at geographical locations 120 degrees apart. This paper will focus on the Earth ground stations
of the lunar architecture. Two types of ground station networks are discussed. One provides Direct to Earth (DTE) support to lunar users using Kaband 23/26Giga-Hertz
(GHz) communication frequencies. The second supports the Lunar Relay Satellite (LRS) that will be using Ka-band 40/37GHz (Q-band). This paper will discuss strategies to
provide a robust operational network in support of various lunar missions and trades of building new antennas at non-NASA facilities, to improve coverage and provide site
diversification for handling rain attenuation.
Goddard Space Flight Center
Autonomic Computing for
Spacecraft Ground Systems
Autonomic computing for spacecraft ground systems increases the system reliability and reduces the cost of spacecraft operations and software maintenance. In this paper,
we present an autonomic computing solution for spacecraft ground systems at NASA Goddard Space Flight Center (GSFC), which consists of an open standard for a
message oriented architecture referred to as the GMSEC architecture (Goddard Mission Services Evolution Center), and an autonomic computing tool, the Criteria Action
Table (CAT). This solution has been used in many upgraded ground systems for NASA 's missions, and provides a framework for developing solutions with higher
autonomic maturity.
Goddard Space Flight Center
Integrating Automation into a
Multi-Mission Operations Center
NASA Goddard Space Flight Center's Space Science Mission Operations (SSMO) Project is currently tackling the challenge of minimizing ground operations costs for
multiple satellites that have surpassed their prime mission phase and are well into extended mission. These missions are being reengineered into a multi-mission operations
center built around modern information technologies and a common ground system infrastructure. The effort began with the integration of four SMEX missions into a similar
architecture that provides command and control capabilities and demonstrates fleet automation and control concepts as a pathfinder for additional mission integrations. The
reengineered ground system, called the Multi-Mission Operations Center (MMOC), is now undergoing a transformation to support other SSMO missions, which include
SOHO, Wind, and ACE. This paper presents the automation principles and lessons learned to date for integrating automation into an existing operations environment for
multiple satellites.
Autonomous Multi-Mission Virtual Ground and Spacecraft Operations
Communications--Autonomous Control and Monitoring
U.S. Patent Applications
Patent Number
Title
Asignee
Abstract
US2006100752A1
Automatic operation system and method
for automating satellite control operation
and satellite ground control systemUSing
the same
None
Provided are an automatic operation system for automating satellite control operation, a method thereof, and an automatic satellite
ground control system using the same. The object of the present research is to provide an automatic operation system that can reduce
operation manpower and cost by automatically operating an entire satellite ground control system, which requires a plurality of operators
to be properly operated, all the times, a method thereof, and an automatic satellite ground control system using the same. The satellite
ground control system of the present research includes: a mission timeline receiving unit, an operation procedure editing unit, an
operation procedure analyzing and code transforming unit, an operation procedure executing unit, a subsystem process state monitoring
unit, and an access managing unit for providing interface with each of the subsystems.
Autonomous Multi-Mission Virtual Ground and Spacecraft Operations
Communications--Autonomous Control and Monitoring
FC
Goddard Space Flight Center
Title
Agent Based Software for the Autonomous Control of
Formation Flying Spacecraft
Goddard Space Flight Center
Experiences with a Requirements-Based Programming
Approach to the Development of a NASA Autonomous Ground
Control System
NTRS
Abstract
Distributed satellite systems is an enabling technology for many future NASA DoD earth and space science missions, such as MMS,
MAXIM, Leonardo, and LISA 1, 2, 3. While formation flying offers significant science benefits, to reduce the operating costs for these
missions it will be essential that these multiple vehicles effectively act as a single spacecraft by performing coordinated observations.
Autonomous guidance, navigation, and control as part of a coordinated fleet-autonomy is a key technology that will help accomplish this
complex goal. This is no small task, as most current space missions require significant input from the ground for even relatively simple
decisions such as thruster burns. Work for the NMP DS1 mission focused on the development of the New Millennium Remote Agent
(NMRA) architecture for autonomous spacecraft control systems. NMRA integrates traditional real-time monitoring and control with
components for constraint-based planning, robust multi-threaded execution, and model-based diagnosis and reconfiguration. The
complexity of using an autonomous approach for space flight software was evident when most of its capabilities were stripped off prior
to launch (although more capability was uplinked subsequently, and the resulting demonstration was very successful).
Requirements-to-Design-to-Code (R2D2C) is an approach to the engineering of computer-based systems that embodies the idea of
requirements-based programming in system development. It goes further however, in that the approach offers not only an underlying
formalism, but full formal development from requirements capture through to the automatic generation of provably-correct code. As
such, the approach has direct application to the development of systems requiring autonomic properties. We describe a prototype tool
to support the method, and illustrate its applicability to the development of LOGOS, a NASA autonomous ground control system, which
exhibits autonomic behavior. Finally, we briefly discuss other areas where the approach and prototype tool are being considered for
application.
Autonomous Multi-Mission Virtual Ground and Spacecraft Operations
Communications—Laser
Company Name
LSA
Title
PASSIVE COMMUNICATION OPTIONS FOR MINIATURE
SATELLITES
DOD SBIR Phase II
Quad Chart
There are many uses for miniature satellites, from Intelligence, Surveillance and Reconnaissance (ISR) to communications relays
and platforms for scientific experiments. The small size and inexpensive nature of these platforms permits a larger number of
satellites to be launched from one vehicle, and also increases the total functionality of our space assets by encouraging earlier
launches of leading edge technologies and devices.
Autonomous Multi-Mission Virtual Ground and Spacecraft Operations
Communications—RF
Company Name
Time Domain Corp.
Title
Low-Probability-of-Intercept/Low-Probability-of-Detection (LPI/LPD)
Data Link
DOD SBIR Phase I
Quad Chart
Time Domain Corporation (TDC) proposes using an Ultra-wideband (UWB) communication system to provide a reliable 30 km RF link
between an unmanned aerial vehicle and a ground station. Pseudo random flipped and time hopped codes provide a whitened pulse
train with very low power spectral density (PSD). The PSD looks like Gaussian distributed noise to most narrowband low noise
detection systems and would be very difficult to detect with wideband systems. The intended UWB receiver uses coherent integration
to pull the signals out of the noise and to reconstruct and decode the information symbols. TDC has demonstrated several long range
line of sight ground-to-ground links (approximately 15 km) using a high gain directional receive antenna, and shorter range LOS
ground-to-UAV links (approximately 1 km) using very small omni-directional antennas and radiated effective isotropic power (E.I.R.P.)
of 9.5 dBm = 9 mW. In Phase I, TDC will define requirements, build a demonstration platform, perform link budget analysis, conduct
LPI/LPD analysis, and plan the prototype design. The program will determine feasibility of meeting size, weight, and power
consumption requirements with UWB radios that provide sufficient data rate over a 30 km link.
Autonomous Multi-Mission Virtual Ground and Spacecraft Operations
Electronics--Radiation-Hard/Resistant Electronics
Company Name
Williams-Pyro, Inc.
Title
Reliable, Lightweight, Low Cost, and Volume Efficient Electrical
Circuitry Development that Facilitates Integration and Checkout for the
Space Tracking
DOD SBIR Phase I
Quad Chart
To facilitate faster assembly and improve the reliability of STSS satellites, Williams-Pyro, Inc. (WPI) proposes to develop a
Multifunctional Flex cable, "MultiFlex." MultiFlex reconfigures cabling to improve the way cables are used in aerospace systems.
MultiFlex consists of integrated micro-interconnects shielded by simple, low-cost, multi-layer EMI shielding for high data rate (e.g.,
USB 2.0). MultiFlex will be integrated with VLSI electronics to allow low-bandwidth communications on existing power distribution
networks, which reduces satellite cabling and minimizes satellite volume and mass. More importantly, MultiFlex will have a built-in
test feature with health indicators (LEDs), which significantly reduces touch labor and time for cable check out during satellite
integration. MultiFlex offers the following distinct advantages over existing cabling configurations: (1) Highly integrated, flexible,
low volume and mass; (2) Low cost, reliable, easy to fabricate and install; (3) Built-in test capability
with health indicators; and (4) Integrated power line communication (PLC) technology.
Autonomous Multi-Mission Virtual Ground and Spacecraft Operations
Information--Computer System Architectures
Company Name
Emergent Space
Technologies, Inc.
Title
Ground Enterprise Management System
Field Center
ARC
NASA SBIR Phase I
Quad Chart
Spacecraft ground systems are on the cusp of achieving "plug-and-play" capability, i.e., they are approaching the state in which the various
components can be quickly integrated with near-automatic interoperability. When properly architected, such systems offer advantages over their
traditional counterparts, such as improved upgradeability and extensibility. They can also be more easily automated and can provide better fault
tolerance. These characteristics are important for all NASA spacecraft, from Exploration to Science. However, plug-and-play systems also pose
some interesting challenges that can undermine their effectiveness. For example, they can be more complex in terms of information management
and system administration. This is especially true when automation is used to reduce the workload of the operators. In fact, as the number of
components increases, as much "situational awareness" of the ground system is needed the spacecraft. A software framework that supports plugand-play integration, while also providing information management and system coordination, is required. Emergent Space Technologies Inc.
therefore proposes to research and develop an Enterprise Management System for spacecraft ground systems. Called GEMS, it will provide
spacecraft controllers and crew members with "situational awareness" of the current state of the ground system as well as understanding of how
events and automated actions are affecting the system in real-time. The innovation lies in the development of algorithms and software adapters that
gather data from the various components of the ground system
to construct a data model that captures the system state and displays it to the controllers.
Autonomous Multi-Mission Virtual Ground and Spacecraft Operations
Information--Computer System Architectures
Company Name
RAM Laboratories, Inc.
Title
REDUCING LATENCIES FOR DISTRIBUTED
ENVIRONMENTS
DOD SBIR Phase II
Quad Chart
This Communications Infrastructure can be used to reduce data latencies across distributed environments while providing a path to
grow as the GMD system expands to consider mobile platforms and new users, including the transition to Link-16 and IPv6
protocols.
Autonomous Multi-Mission Virtual Ground and Spacecraft Operations
Information--Data Acquistion and End-to-End Management
Company Name
TenXsys Inc.
Title
DISTRIBUTED OPERATIONS SOFTWARE
FOR WIRELESS HANDHELD COMPUTERS
Field Center
JSC
NASA SBIR Phase II
Quad Chart
TenXsys will extend the Adaptive Data Retrieval and Optimized, Intelligent Transfer (ADROIT) system that works in conjunction
with one or more applications on a personal digital assistant (PDA) to deliver mission-essential information to distributed operators
over a wireless communications infrastructure to integrate voice, video, and more complex 3-D image displays and manipulation
into the design.
Autonomous Multi-Mission Virtual Ground and Spacecraft Operations
Information--Data Acquistion and End-toEnd Management
WIPO
Patent Number
WO03021866A2
Title
Point-To-Point Protocol
Over Ethernet for Mobile
Platforms
Assignee
The Boeing Company
Abstract
A communications system that provides broadband access to passengers of mobile platforms includes a router located on the mobile platform. A
network is connected to the router. User communication decives (UCDs)connected to the network, wherein in the UCDs establish point-to-point
over Ethernet (PPPoE) sessions with the router. A transmitter and a receiver are connected to the router. A satelite and aground station are in
communication with the transmitter and the receiver. A distributed communications system includes virtual private networks (VPN) and is
connected to the ground station. A first address manager assigns the public Ip addresses to UCDs when the UCDs request access to the VPNs and
private IP addresses for other network service. The UCDs employ IPSec protocol when accessing the VPNs.
Autonomous Multi-Mission Virtual Ground and Spacecraft Operations
Information--Data Acquistion and End-to-End Management
FC
Jet Propulsion Laboratory
Jet Propulsion Laboratory
Title
Frame synchronization in Jet Propulsion Laboratory's
Advanced Multi-Mission Operations System (AMMOS)
Lessons learned from five years of multi-mission operations
NTRS
Abstract
The Jet Propulsion Laboratory's Advanced Multi-Mission Operations System system processes data received from deep-space
spacecraft, where error rates can be high, bit rates are low, and data is unique precious.
No Abstract Available
Autonomous Multi-Mission Virtual Ground and Spacecraft Operations
Information--Human-Computer Interfaces
FC
GSFC
Title
A Virtual Mission Operations Center
- Collaborative Environment
Glenn
Research
Center
Development of Network-based
Communications Architectures for
Future NASA Missions
Jet Propulsion
Laboratory
Development Roadmap of an
Evolvable and Extensible MultiMission Telecom Planning and
Analysis Framework
Goddard Space
Flight Center
REACH Real-Time Data Awareness
in Multi-Spacecraft Missions
NTRS
Abstract
Development of technologies that enable significant reductions in the cost of space mission operations is critical if constellations, formations, federations and sensor webs, are to be
economically feasible. One approach to cost reduction is to infuse automation technologies into mission operations centers so that fewer personnel are needed for mission support. But
missions are more culturally and politically adverse to the risks of automation. Reducing the mission risk associated with increased use of automation within a MOC is therefore of great
importance. The belief that mission risk increases as more automation is used stems from the fact that there is inherently less direct human oversight to investigate and resolve anomalies
in an unattended MOC. The Virtual Missions Operations Center - Collaborative Environment (VMOC-CE) project was launched to address this concern. The goal of the VMOC-CE project
is to identify, develop, and infuse technology to enable mission operations between onsite operators and on-call personnel in geographically dispersed locations. VMOC-CE enables
missions to more readily adopt automation because off-site operators and engineers can more easily identify, investigate, and resolve anomalies without having to be present in the MOC.
The VMOC-CE intent is to have a single access point for all resources used in a collaborative mission operations environment. Team members will be able to interact during spacecraft
operations, specifically for resolving anomalies, utilizing a desktop computer and the Internet. Mission operations management can use the VMOC-CE as a tool to participate in and
monitor status of anomaly resolution or other mission operations issues. In this paper we present the VMOC-CE project, system capabilities and technologies, operations concept, and
results of its pilot in support of the Earth Science Mission Operations System (ESMOS).
Since the Vision for Space Exploration (VSE) announcement, NASA has been developing a communications infrastructure that combines existing terrestrial techniques with newer
concepts and capabilities. The overall goal is to develop a flexible, modular, and extensible architecture that leverages and enhances terrestrial networking technologies that can either be
directly applied or modified for the space regime. In addition, where existing technologies leaves gaps, new technologies must be developed. An example includes dynamic routing that
accounts for constrained power and bandwidth environments. Using these enhanced technologies, NASA can develop nodes that provide characteristics, such as routing, store and
forward, and access-on-demand capabilities. But with the development of the new infrastructure, challenges and obstacles will arise. The current communications infrastructure has been
developed on a mission-by-mission basis rather than an end-to-end approach; this has led to a greater ground infrastructure, but has not encouraged communications between spacebased assets. This alone provides one of the key challenges that NASA must encounter. With the development of the new Crew Exploration Vehicle (CEV), NASA has the opportunity to
provide an integration path for the new vehicles and provide standards for their development. Some of the newer capabilities these vehicles could include are routing, security, and
Software Defined Radios (SDRs). To meet these needs, the NASA/Glenn Research Center s (GRC) Network Emulation Laboratory (NEL) has been using both simulation and emulation to
study and evaluate these architectures. These techniques provide options to NASA that directly impact architecture development. This paper identifies components of the infrastructure
that play a pivotal role in the new NASA architecture, develops a scheme using simulation and emulation for testing these architectures and demonstrates how NASA can strengthen the
new infrastructure by implementing these concepts.
In this paper, we describe the development roadmap and discuss the various challenges of an evolvable and extensible multi-mission telecom planning and analysis framework. Our longterm goal is to develop a set of powerful flexible telecommunications analysis tools that can be easily adapted to different missions while maintain the common Deep Space
Communication requirements. The ability of re-using the DSN ground models and the common software utilities in our adaptations has contributed significantly to our development efforts
measured in terms of consistency, accuracy, and minimal effort redundancy, which can translate into shorter development time and major cost savings for the individual missions. In our
roadmap, we will address the design principles, technical achievements and the associated challenges for following telecom analysis tools (i) Telecom Forecaster Predictor - TFP (ii)
Unified Telecom Predictor - UTP (iii) Generalized Telecom Predictor - GTP (iv) Generic TFP (v) Web-based TFP (vi) Application Program
NASA's Advanced Architectures and Automation Branch at the Goddard Space Flight Center (Code 588) saw the potential to reduce the cost of constellation missions by creating new
user interfaces to the ground system health-and-safety data. The goal is to enable a small Flight Operations Team (FOT) to remain aware and responsive to the increased amount of
ground system information in a multi-spacecraft environment. Rather than abandon the tried and true, these interfaces were developed to run alongside existing ground system software to
provide additional support to the FOT. These new user interfaces have been combined in a tool called REACH. REACH-the Real-time Evaluation and Analysis of Consolidated Health-is a
software product that uses advanced visualization techniques to make spacecraft anomalies easy to spot, no matter how many spacecraft are in the constellation. REACH reads
numerous real-time streams of data from the ground system(s) and displays synthesized information to the FOT such that anomalies are easy to pick out and investigate.
Autonomous Multi-Mission Virtual Ground and Spacecraft Operations
Information--Human-Computer Interfaces
FC
Goddard Space
Flight Center
Title
REACH Real-Time Data Awareness
in Multi-Spacecraft Missions
Dryden Flight
Research
Center;
Marshall Space
Flight Center
The Real Time Mission Monitor A
Situational Awareness Tool For
Managing Experiment Assets
NTRS
Abstract
Missions have been proposed that will use multiple spacecraft to perform scientific or commercial tasks. Indeed, in the commercial world, some spacecraft constellations already exist.
Aside from the technical challenges of constructing and flying these missions, there is also the financial challenge presented by the tradition model of the flight operations team (FOT)
when it is applied to a constellation mission. Proposed constellation missions range in size from three spacecraft to more than 50. If the current ratio of three-to-five FOT personnel per
spacecraft is maintained, the size of the FOT becomes cost prohibitive. The Advanced Architectures and Automation Branch at the Goddard Space Flight Center (GSFC Code 588) saw
the potential to reduce the cost of these missions by creating new user interfaces to the ground system health-and-safety data. The goal is to enable a smaller FOT to remain aware and
responsive to the increased amount of ground system information in a multi-spacecraft environment. Rather than abandon the tried and true, these interfaces were developed to run
alongside existing ground system software to provide additional support to the FOT. These new user interfaces have been combined in a tool called REACH. REACH-the Real-time
Evaluation and Analysis of Consolidated Health-is a software product that uses advanced visualization techniques to make spacecraft anomalies easy to spot, no matter how many
spacecraft are in the constellation. REACH reads a real-time stream of data from the ground system and displays it to the FOT such that anomalies are easy to pick out and investigate.
Data visualization has been used in ground system operations for many years. To provide a unique visualization tool, we developed a unique source of data to visualize the REACH
Health Model Engine. The Health Model Engine is rule-based software that receives real-time telemetry information and outputs 'health' information related to the subsystems and
spacecraft that the telemetry belong to. The Health Engine can run out-of-the-box or can be tailored with a scripting language. Out of the box, it uses limit violations to determine the health
of subsystems and spacecraft when tailored, it determines health using equations combining the values and limits of any telemetry in the spacecraft. The REACH visualizations then 'roll
up' the information from the Health Engine into high level, summary displays. These summary visualizations can be 'zoomed' into for increasing levels of detail. Currently REACH is
installed in the Small Explorer (SMEX) lab at GSFC, and is monitoring three of their five spacecraft. We are scheduled to install REACH in the Mid-sized Explorer (MIDEX) lab, which will
allow us to monitor up to six more spacecraft. The process of installing and using our 'research' software in an operational environment has provided many insights into which parts of
REACH are a step forward and which of our ideas are missteps. Our paper explores both the new concepts in spacecraft health-and-safety visualization, the difficulties of such systems in
the operational environment, and the cost and safety issues of multi-spacecraft missions.
The NASA Real Time Mission Monitor (RTMM) is a situational awareness tool that integrates satellite, airborne and surface data sets; weather information; model and forecast outputs;
and vehicle state data (e.g., aircraft navigation, satellite tracks and instrument field-of-views) for field experiment management RTMM optimizes science and logistic decision-making
during field experiments by presenting timely data and graphics to the users to improve real time situational awareness of the experiment's assets. The RTMM is proven in the field as it
supported program managers, scientists, and aircraft personnel during the NASA African Monsoon Multidisciplinary Analyses experiment during summer 2006 in Cape Verde, Africa. The
integration and delivery of this information is made possible through data acquisition systems, network communication links and network server resources built and managed by
collaborators at NASA Dryden Flight Research Center (DFRC) and Marshall Space Flight Center (MSFC). RTMM is evolving towards a more flexible and dynamic combination of sensor
ingest, network computing, and decision-making activities through the use of a service oriented architecture based on community standards and protocols.
Autonomous Multi-Mission Virtual Ground and Spacecraft Operations
Information--Software Development Environments
FC
Ames Research Center; Jet
Propulsion Laboratory
Title
Leveraging Existing Mission Tools in a Re-Usable, ComponentBased Software Environment
NTRS
Abstract
Emerging methods in component-based software development offer significant advantages but may seem incompatible with existing
mission operations applications. In this paper we relate our positive experiences integrating existing mission applications into
component-based tools we are delivering to three missions. In most operations environments, a number of software applications have
been integrated together to form the mission operations software. In contrast, with component-based software development chunks of
related functionality and data structures, referred to as components, can be individually delivered, integrated and re-used. With the
advent of powerful tools for managing component-based development, complex software systems can potentially see significant
benefits in ease of integration, testability and reusability from these techniques. These benefits motivate us to ask how componentbased development techniques can be relevant in a mission operations environment, where there is significant investment in software
tools that are not component-based and may not be written in languages for which component-based tools even exist. Trusted and
complex software tools for sequencing, validation, navigation, and other vital functions cannot simply be re-written or abandoned in
order to gain the advantages offered by emerging component-based software techniques. Thus some middle ground must be found.
We have faced exactly this issue, and have found several solutions. Ensemble is an open platform for development, integration, and
deployment of mission operations software that we are developing. Ensemble itself is an extension of an open source, componentbased software development platform called Eclipse. Due to the advantages of component-based development, we have been able to
vary rapidly develop mission operations tools for three surface missions by mixing and matching from a common set of mission
operation components. We have also had to determine how to integrate existing mission applications for sequence development,
sequence validation, and high level activity planning, and other functions into a component-based environment. For each of these, we
used a somewhat different technique based upon the structure and usage of the existing application.
Autonomous Multi-Mission Virtual Ground and Spacecraft Operations
Information--Software Development Environments
Company Name
Command and Control Technologies
Corporation
Title
DATA DESCRIPTION EXCHANGE
SERVICES FOR HETEROGENEOUS
VEHICLE AND SPACEPORT CONTROL AND
MONITORING SYSEMS
Field Center
KSC
NASA SBIR Phase II
Quad Chart
This new capability will reduce development, operations, and support costs for legacy and future systems that are part of
ground and space based distributed command and control systems. It will also establish a space systems information
exchange model that can support future highly interoperable and mobile software systems.
Autonomous Multi-Mission Virtual Ground and Spacecraft Operations
Information--Software Tools for Distributed Analysis and Simulation
Company Name
RAM Laboratories, Inc.
Title
MANAGEMENT OF DISTRIBUTED REAL-TIME
DATABASES
DOD SBIR Phase II
Quad Chart
Command and Control, Battle Management, and Communications (C2BMC) rely on a publish/subscribe environment to address battle
management challenges associated with utilizing available data sources to detect, discriminate, and destroy enemy missiles in a
manner of minutes, if not seconds. Latencies for these systems range from tens of seconds down to sub-seconds for real-time and
near-real-time elements of the system.
Autonomous Multi-Mission Virtual Ground and Spacecraft Operations
Propulsion--Aircraft Engines
Company Name
Technical Directions, Inc.
Title
Aeropropulsion and Power Technology
DOD SBIR Phase I
Quad Chart
Small low-cost propulsion systems are required for a variety of new unmanned air vehicle applications. Turbojet engine-generator systems are
considered desirable for these small aircraft applications where compact size and low-cost are critical parameters. Technical Directions Inc.
has developed a family of very low-cost turbojet engines for these aircraft applications using automotive turbocharger rotating components that
satisfy many of the requirements in these applications. Future aircraft are being considered with increased mission complexity and additional
range requirements. With frame sizes already established in many of the small aircraft applications, finding space for additional fuel to satisfy
the mission requirements becomes almost an impossible assignment. This program will create the technology for a performance upgrade to
existing low-cost engines to permit a 10-20% range extension. New design methods will be explored for the aerodynamic engine components
that would provide operating efficiency gains of 5-15%. In these small turbojet engines, good component efficiencies are more difficult to
achieve due to the effects of additional leakage and manufacturing tolerances making the performance improvements even more challenging.
Accomplishing the program objectives herein will provide the desired additional range and mission capability without the need for additional
fuel.
Autonomous Multi-Mission Virtual Ground and Spacecraft Operations
Sensors and Sources--Microwave/Submillimeter
FC
Glenn Research Center
Title
Steerable Space Fed Lens Array for Low-Cost Adaptive
Ground Station Applications
NTRS
Abstract
The Space Fed Lens Array (SFLA) is an alternative to a phased array antenna that replaces large numbers of expensive solid-state phase
shifters with a single spatial feed network. SFLA can be used for multi-beam application where multiple independent beams can be
generated simultaneously with a single antenna aperture. Unlike phased array antennas where feed loss increases with array size, feed
loss in a lens array with more than 50 elements is nearly independent of the number of elements, a desirable feature for large apertures. In
addition, SFLA has lower cost as compared to a phased array at the expense of total volume and complete beam continuity. For ground
station applications, both of these tradeoff parameters are not important and can thus be exploited in order to lower the cost of the ground
station. In this paper, we report the development and demonstration of a 952-element beam-steerable SFLA intended for use as a low cost
ground station for communicating and tracking of a low Earth orbiting satellite. The dynamic beam steering is achieved through switching to
different feed-positions of the SFLA via a beam controller.
Autonomous Multi-Mission Virtual Ground and Spacecraft Operations
Verification and Validation--Operations Concepts and Requirements
FC
Jet Propulsion Laboratory
Title
Multi-mission ground data systems - breakthroughs and
challenges
NTRS
Abstract
Given the cost-constrained nature of JPL Flight Projects, especially Discovery Class missions, there is more and more pressure to reduce
the costs associated with mission operations, particularly the costs for the Ground Data System. This paper explores the successes (and
failures) of using the Mission Management Office (MMO) GDS team to provide a common set of services, tools, procedures, and products
to JPL Flight Projects.
Autonomous Multi-Mission Virtual Ground and Spacecraft Operations
Verification and Validation--Simulation Modeling Environment
FC
Goddard Space Flight
Center
Title
MPGT - THE MISSION
PLANNING
GRAPHICAL TOOL
NTRS
Abstract
The Mission Planning Graphical Tool (MPGT) provides mission analysts with a mouse driven graphical representation of the spacecraft and environment data used in spaceflight planning.
Developed by the Flight Dynamics Division at NASA's Goddard Space Flight Center, MPGT is designed to be a generic tool that can be configured to analyze any specified earth orbiting
spacecraft mission. The data is presented as a series of overlays on top of a 2-dimensional or 3-dimensional projection of the earth. Up to six spacecraft orbit tracks can be drawn at one
time. Position data can be obtained by either an analytical process or by use of ephemeris files. If the user chooses to propagate the spacecraft orbit using an ephemeris file, then Goddard
Trajectory Determination System (GTDS) formatted ephemeris files must be supplied. The MPGT User's Guide provides a complete description of the GTDS ephemeris file format so that
users can create their own. Other overlays included are ground station antenna masks, solar and lunar ephemeris, Tracking Data and Relay Satellite System (TDRSS) coverage, a field-ofview swath, and orbit number. From these graphical representations an analyst can determine such spacecraft-related constraints as communication coverage, interference zone
infringement, sunlight availability, and instrument target visibility. The presentation of time and geometric data as graphical overlays on a world map makes possible quick analyses of trends
and time-oriented parameters. For instance, MPGT can display the propagation of the position of the Sun and Moon over time, shadowing of sunrise sunset terminators to indicate
spacecraft and Earth day night, and color coding of the spacecraft orbit tracks to indicate spacecraft day night. With the 3-dimensional display, the user specifies a vector that represents the
position in the universe from which the user wishes to view the earth. From these "viewpoint" parameters the user can zoom in on or rotate around the earth. The zoom feature is also
available with the 2-dimensional map image. The program contains data files of world map continent coordinates, contour information, antenna mask coordinates, and a sample star catalog.
Since the overlays are designed to be mission independent, no software modifications are required to satisfy the different requirements of various spacecraft. All overlays are generic with
communication zone contours and spacecraft terminators generated analytically based on spacecraft altitude data. Interference zone contours are user-specified through text-edited data
files. Spacecraft orbit tracks are specified via Keplerian, Cartesian, or DODS (Definitive Orbit Determination System) orbit vectors. Finally, all time-related overlays are based on a usersupplied epoch. A user interface subsystem allows the user to alter any system mission or graphics parameter through a series of pull-down menus and pop-up data entry panels. The user
can specify, load, and save mission and graphic data files, control graphical presentation formats, enter a DOS shell, and terminate the system. The interface automatically performs error
checking and data validation on all data input from either a file or the keyboard. A help facility is provided. MPGT also includes a software utility called ShowMPGT which displays screen
images that were generated and saved with the MPGT system. Specific sequences of images can be recalled without having to reset graphics and mission related parameters. The MPGT
system does not provide hardcopy capabilities however this capability will be present in the next release. To obtain hardcopy graphical output, the PC must be configured with a printer that
captures the video signal and copies it onto a hardcopy medium. MPGT is written in FORTRAN, C, and Macro Assembler for use on IBM PC compatibles running MS-DOS v3.3 or higher
which are configured with the following hardware an 80X87 math coprocessor, an EGA or VGA board, 1.3Mb of disk space and 620K of RAM. Due to this memory requirement, it is
recommended that a memory manager or memory optimizer be run prior to executing MPGT. A mouse is supported, but is optional. The provided MPGT system executables were created
using the following compilers Microsoft FORTRAN v5.1, Microsoft C compiler v6.0 and Microsoft Macro Assembler v6.0. These MPGT system executables also incorporate object code from
two proprietary programs HALO Professional Kernel Graphics System v2.0 (copyright Media Cybernetics, Inc., 1981-1992), which is distributed under license agreement with Media
Cybernetics, Incorporated and The Screen Generator v5.2, which is distributed with permission of The West Chester Group. To build the system executables from the provided source code,
the three compilers and two commercial programs would all be required. Please note that this version of MPGT is not compatible with Halo '88. The standard distribution medium for MPGT
is a set of two 3.5 inch 720K MS-DOS format diskettes. The contents of the diskettes are compressed using the PKWARE archiving tools. The utility to unarchive the files, PKUNZIP.EXE
v2.04g, is included. MPGT was developed in 1989 and version 3.0 was released in 1992. HALO is a Registered trademark of Media Cybernetics, Inc. Microsoft and MS-DOS are Registered
trademarks of Microsoft Corporation. PKWARE and PKUNZIP are Registered trademarks of PKWARE, Inc. All trademarks mentioned in this abstract appear for identification purposes only
and are the property of their respective companies.
Autonomous Multi-Mission Virtual Ground and Spacecraft Operations
Verification and Validation--Simulation Modeling Environment
FC
Glenn Research Center
Title
Space Communications
Emulation Facility
Jet Propulsion
Laboratory
An Application of the
"Virtual Spacecraft"
Concept in Evaluation
of the Mars Pathfinder
Lander Low Gain
Antenna
NTRS
Abstract
Establishing space communication between ground facilities and other satellites is a painstaking task that requires many precise calculations dealing with relay time, atmospheric conditions,
and satellite positions, to name a few. The Space Communications Emulation Facility (SCEF) team here at NASA is developing a facility that will approximately emulate the conditions in
space that impact space communication. The emulation facility is comprised of a 32 node distributed cluster of computers each node representing a satellite or ground station. The objective
of the satellites is to observe the topography of the Earth (water, vegetation, land, and ice) and relay this information back to the ground stations. Software originally designed by the
University of Kansas, labeled the Emulation Manager, controls the interaction of the satellites and ground stations, as well as handling the recording of data. The Emulation Manager is
installed on a Linux Operating System, employing both Java and C programming codes. The emulation scenarios are written in extensible Markup Language, XML. XML documents are
designed to store, carry, and exchange data. With XML documents data can be exchanged between incompatible systems, which makes it ideal for this project because Linux, MAC and
Windows Operating Systems are all used. Unfortunately, XML documents cannot display data like HTML documents. Therefore, the SCEF team uses XML Schema Definition (XSD) or just
schema to describe the structure of an XML document. Schemas are very important because they have the capability to validate the correctness of data, define restrictions on data, define
data formats, and convert data between different data types, among other things. At this time, in order for the Emulation Manager to open and run an XML emulation scenario file, the user
must first establish a link between the schema file and the directory under which the XML scenario files are saved. This procedure takes place on the command line on the Linux Operating
System. Once this link has been established the Emulation manager validates all the XML files in that directory against the schema file, before the actual scenario is run. Using some very
sophisticated commercial software called the Satellite Tool Kit (STK) installed on the Linux box, the Emulation Manager is able to display the data and graphics generated by the execution
of a XML emulation scenario file. The Emulation Manager software is written in JAVA programming code. Since the SCEF project is in the developmental stage, the source code for this type
of software is being modified to better fit the requirements of the SCEF project. Some parameters for the emulation are hard coded, set at fixed values. Members of the SCEF team are
altering the code to allow the user to choose the values of these hard coded parameters by inserting a toolbar onto the preexisting GUI.
The virtual spacecraft concept is embodied in a set of subsystems, either in the form of hardware or computational models, which together represent all, or a portion of, a spacecraft. For
example, the telecommunications transponder may be a hardware prototype while the propulsion system may exist only as a simulation. As the various subsystems are realized in
hardware, the spacecraft becomes progressively less virtual. This concept is enabled by JPL's Mission System Testbed which is a set of networked workstations running a message passing
operating system called "TRAMEL" which stands for Task Remote Asynchronous Message Exchange Layer. Each simulation on the workstations, which may in fact be hardware controlled
by the workstation, "publishes" its operating parameters on TRAMEL and other simulations requiring those parameters as input may "subscribe" to them. In this manner, the whole
simulation operates as a single virtual system. This paper describes a simulation designed to evaluate a communications link between the earth and the Mars Pathfinder Lander module as it
descends under a parachute through the Martian atmosphere toward the planet's surface. This link includes a transmitter and a low gain antenna on the spacecraft and a receiving antenna
and receiver on the earth as well as a simulation of the dynamics of the spacecraft. The transmitter, the ground station antenna, the receiver and the dynamics are all simulated
computationally while the spacecraft antenna is implemented in hardware on a very simple spacecraft mockup. The dynamics simulation is a record of one output of the ensemble of outputs
of a Monte Carlo simulation of the descent. Additionally, the antenna spacecraft mock-up system was simulated using APATCH, a shooting and bouncing ray code developed by Demaco,
Inc. The antenna simulation, the antenna hardware, and the link simulation are all physically located in different facilities at JPL separated by several hundred meters and are linked via the
local area network (LAN).
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