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Search indexed NASA NTRS and DOE OSTI research on propulsion, heat transfer, battery materials and energy systems. Follow report and document links to the original sources.
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Megawatt ion-propulsion power conditioning system design for manned interplanetary missions, noting payload-mass ratio problems, advantages, etc
This CD contains files that support the talk (see CASI ID 20100021404). There are 24 models that relate to the ADAPT system and 1 Excel worksheet. In the paper an investigation into the use of Bayesian networks to construct large-scale diagnostic systems is described. The high-level specifications, Bayesian networks, clique trees, and arithmetic circuits representing 24 different electrical power systems are described in the talk. The data in the CD are the models of the 24 different power systems.
Various space electrical power system masses are analyzed with particular emphasis on the power management and distribution (PMAD) portion. The electrical power system (EPS) is divided into functional blocks: source, interconnection, storage, transmission, distribution, system control and load. The PMAD subsystem is defined as all the blocks between the source, storage and load, plus the power conditioning equipment required for the source, storage and load. The EPS mass of a wide range of spacecraft is then classified as source, storage or PMAD and tabulated in a database. The intent of the database is to serve as a reference source for PMAD masses of existing and in-design spacecraft. The PMAD masses in the database range from 40 kg/kW to 183 kg/kW across the spacecraft systems studied. Factors influencing the power system mass are identified. These include the total spacecraft power requirements, total amount of load capacity and physical size of the spacecraft. It is found that a new utility class of power systems, represented by Space Station Freedom, is evolving.
Various space electrical power system masses are analyzed with particular emphasis on the power management and distribution (PMAD) portion. The electrical power system (EPS) is divided into functional blocks: source, interconnection, storage, transmission, distribution, system control and load. The PMAD subsystem is defined as all the blocks between the source, storage and load, plus the power conditioning equipment required for the source, storage and load. The EPS mass of a wide range of spacecraft is then classified as source, storage or PMAD and tabulated in a database. The intent of the database is to serve as a reference source for PMAD masses of existing and in-design spacecraft. The PMAD masses in the database range from 40 kg/kW to 183 kg/kW across the spacecraft systems studied. Factors influencing the power system mass are identified. These include the total spacecraft power requirements, total amount of load capacity and physical size of the spacecraft. It is found that a new utility class of power systems, represented by Space Station Freedom, is evolving.
Primary electrical power for loads in the Apollo Command and Service Modules is furnished by a system of three fuel cell powerplants. The powerplants convert cryogenic hydrogen and oxygen into direct current. Water formed in the fuel cell reactions is supplied to the spacecraft for crew use. Waste heat is rejected through a coolant loop to a spacecraft radiator. The ability of the fuel cell to meet its voltage-power requirements and to control its own operating temperature under the environmental extremes of launch, ascent, earth orbit, translunar flight, and lunar orbit was demonstrated on successively more difficult missions leading up to the first lunar landing.
During orbit day, solar energy collected by the solar cell arrays and transformed into electrical energy is used to power the spacecraft subsystems, including the control system. In conventional spacecraft designs, a portion of the energy collected during the light portion of the orbit is stored in a set of batteries for use during orbit night. In the Integrated Power/Attitude Control System (IPACS) approach, that energy is stored in the rotating flywheel in the form of kinetic energy. Umbra electrical power demands are satisfied by attaching a generator to the wheel shaft and despinning the rotor. Through this approach, the battery system is no longer required and thus is eliminated.
A brief description is given of the schemes suggested for providing electrical power aboard spacecraft. An attempt was made to list the major advantages and disadvantages of each scheme and to identify the particular areas in which it appears likely further research work may prove fruitful. Special attention was given to those research areas that appear suitable for colleges and university work. Chemical, solar, and nuclear power sources are considered together with the appropriate thermoelectric and turboelectric conversion systems.
As power levels of spacecraft rise to the 50 to 100 kW range, it becomes apparent that low voltage (28 V) dc power distribution and management systems will not operate efficiently at these higher power levels. The concept of transforming a solar array voltage at 150 V dc into a 1000 V ac distribution system operating at 20 kHz is examined. The transformation is accomplished with series-resonant inverter by using a rotary transformer to isolate the solar array from the spacecraft. The power can then be distributed in any desired method such as three phase delta to delta. The distribution voltage can be easily transformed to any desired load voltage and operating frequency. The reasons for the voltage limitations on the solar array due to plasma interactions and the many advantages of a high voltage, high frequency at distribution system are discussed.
Power system developments for rockets and space flights - fuel cells, improved zinc-silver oxide battery, brushless dc motors, switching transformer regulator, and Pegasus electricity
The results of electrical power system (EPS) analysis of the planned Shuttle Transportation System Flight 1 mission are presented. The capability of the orbiter EPS to support the planned flight and to provide program tape information and supplementary data specifically requested by the flight operations directorate was assessed. The analysis was accomplished using the orbiter version of the spacecraft electrical power simulator program, operating from a modified version of orbiter electrical equipment utilization baseline revision four. The results indicate that the nominal flight, as analyzed, is within the capabilities of the orbiter power generation system, but that a brief, and minimal, current overload may exist between main distributor 1 and mid power controlled 1, and that inverter 9 may the overloaded for extended periods of time. A comparison of results with launch commit criteria also indicated that some of the presently existing launch redlines may be violated during the terminal countdown.
The current understanding of the power transfer mechanisms by which power is supplied to a planet's magnetosphere by the kinetic energy of planetary spin and the energy flux of the impinging solar wind is applied to the case of Uranus, in order to predict the detectability of radio and auroral emissions by the planetary radio astronomy (PRA) and UV spectrometer (UVS) instruments of the Voyager spacecraft. The power available for the two energy transfer phenomena cited is a function of Uranus' magnetic moment, which is presently derived for each power source as a function of the date of first detection of radio emissions by the PRA or auroral emissions by the UVS.
Released in 1995, the Trilogy cardiac pacemaker is the fourth generation of a unit developed in the 1970s by NASA, Johns Hopkins Applied Physics Laboratory and St. Jude Medical's Cardiac Rhythm Management Division (formerly known as Pacesetter Systems, Inc.). The new system incorporates the company's PDx diagnostic and programming software and a powerful microprocessor that allows more functions to be fully automatic and gives more detailed information on the patient's health and the performance of the pacing systems. The pacemaker incorporates bidirectional telemetry used for space communications for noninvasive communication with the implanted pacemaker, smaller implantable pulse generators from space microminiaturization, and longer-life batteries from technology for spacecraft electrical power systems.
After 39 years of continuous operation in space, the output of the Voyager 1 & 2 spacecraft RTG power systems has decreased to the point where managing the power margin and maintaining thermal control has become increasingly difficult. As the total power dissipation in the bus has decreased, propellant line temperatures and margin above minimum AFTs have decreased, creating risk of the hydrazine freezing (at 1.6°C). This is further complicated by the lack of existing thermal models that can be used to assess propellant tank and line temperatures. In 2014, an effort was begun to create a Voyager spacecraft thermal model for that purpose.A steady-state Thermal Desktop model has been created from scratch over the past two years. The initial thermal model development was started by Applied Sciences Laboratory (ASL) under contract to JPL. The effort relied primarily on archived manufacturing drawings, limited documentation, interviews of senior engineers who worked on the Voyager design and implementation, and the experience of the Voyager Flight Operations team.Data from the Voyager System Thermal Vacuum tests is no longer available, making it necessary to correlate the model to more recent flight data and small in-flight tests. Correlation was achieved to within ±5°C for a hot case and a cold case (both data sets from 2014). However, the flight system has very few temperature sensors directly on propellant lines. So the task remains to determine how best to use the model, in conjunction with flight data, to make sure the Voyagers can continue to fly successfully. How does one go about creating a thermal model for a spacecraft that is already launched, has limited existing mechanical description files, no thermal model in order to do a maneuver that was never planned when the spacecraft was designed?
Electrical power control and conditioning subsystem for Manned Orbital Research Laboratory /MORL/, noting Brayton cycle, system optimization, alternator and frequency converter selection
The satellite system requires the use of a large satellite antenna and spacecraft array power of about 12 kW or more depending on the operating frequency. Technology developments needed include large offset reflector multibeam antennas, satellite electrical power sybsystems providing greater than 12 kW of power, signal switching hardware, and linearized efficient solid state amplifiers for the satellite-aided mobile band. Presently there is no frequency assignment for this service, and it is recommended that an allocation be pursued. The satellite system appears to be within reasonable extrapolation of the state of the art. It is further recommended that the satellite-aided system spacecraft definition studies and supporting technology development be initiated.
In 1992, the Naval Research Laboratory (NRL) started a program to qualify a large diameter common pressure vessel (CPV) nickel-hydrogen (NiH2) batteries for use on future Navy/NRL spacecraft electrical power subsystems. NRL's involvement with the qualification of CPV NiH2 batteries dates back to 1988 when COMSAT and Johnson Controls, Inc. initiated a joint effort to fly the first ever NiH2 CPV in space. A later NRL-JCI cooperative research and development agreement led to the launch of a space experiment in 1993 and to the use of a single NiH2 CPV battery on the BMDO Clementine spacecraft in 1994. NRL initiated procurement of two, 50 Ah CPV NiH2 batteries in the Fall of 1992. The two batteries were delivered to NRL in June 1994. NiH2 CPV batteries have almost 2x the specific energy (Wh/kg) of nickel cadium batteries and 2x the energy density (Wh/l) of individual pressure vessel NiH2 CPV's. This presentation discusses the results of electrical and mechanical qualification tests conducted at NRL. The tests included electrical characterization, standard capacity, random vibration, peak load, and thermal vacuum. The last slides of the presentation show initial results from the life cycle tests of the second NiH2 CPV battery at 40% depth of discharge and a temperature of 10 C.
The space station module power management and distribution (SSM/PMAD) breadboard models power distribution and management, including scheduling, load prioritization, and a fault detection, identification, and recovery (FDIR) system within a Space Station Freedom habitation or laboratory module. This 120 VDC system is capable of distributing up to 30 kW of power among more than 25 loads. In addition to the power distribution hardware, the system includes computer control through a hierarchy of processes. The lowest level consists of fast, simple (from a computing standpoint) switchgear that is capable of quickly safing the system. At the next level are local load center processors, (LLP's) which execute load scheduling, perform redundant switching, and shed loads which use more than scheduled power. Above the LLP's are three cooperating artificial intelligence (AI) systems which manage load prioritizations, load scheduling, load shedding, and fault recovery and management. Recent upgrades to hardware and modifications to software at both the LLP and AI system levels promise a drastic increase in speed, a significant increase in functionality and reliability, and potential for further examination of advanced automation techniques. The background, SSM/PMAD, interface to the Lewis Research Center test bed, the large autonomous spacecraft electrical power system, and future plans are discussed.
Flight and ground system operability has been a focus area on the Europa Clipper Project since early in its formulation phase. This has given the operations team the opportunity to influence the design, with a goal of increasing overall system operability. This paper presents example operability challenges, opportunities, and solutions arising from the pre-Critical Design Review (CDR) system design. The integrated wing assembly design directly couples a scientific instrument (the REASON sounding radar) to the spacecraft’s power source (solar array wing panels). Impacts to mission operations of this design include: increased slew durations; solar array pointing constraints during inner cruise, Europa flybys, and orbit trim maneuvers; and stray light intrusions into the stellar reference units’ keep out zones. The use of CCSDS File Delivery Protocol (CFDP) Class-2 for reliable downlink of the large volume of Europa Clipper science data is described, along with nominal and off-nominal use cases. The effort to improve post-launch spacecraft visibility by adding a third low-gain antenna to the spacecraft is detailed. The design of the bulk data store has necessitated the implementation of accountable data products (ADPs), accountability identifiers (AIDs), and metadata packets to provide end-to-end science data accountability. To streamline and automate the flight rules generation and checking process, a first order and temporal logic-based solution of expressing flight rules without ambiguity, and whose programmatic implementation can be automated, is proposed. The focus on operability has had a positive influence on Europa Clipper design decisions, although cost, schedule, budget, heritage, and other technical concerns have many times outweighed operability concerns. However, experience to date demonstrates that this approach to operability results in more thorough, balanced consideration of the effect of early design trades and decisions on the operations phase of a mission than seen in many previous missions, and provides operations development insight into prioritizing work to go.