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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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NASA NTRS1987

On sine dwell or broadband methods for modal testing

For large, complex spacecraft structural systems, the objectives of the modal test are outlined. Based on these objectives, the comparison criteria for the modal test methods, namely, the broadband excitation and the sine dwell methods are established. Using the Galileo spacecraft modal test and the Centaur G Prime upper stage vehicle modal test as examples, the relative advantages or disadvantages of each method are examined. The usefulness or shortcoming of the methods are given from a practicing engineer's view point.

Chen, Jay-Chung↗
NASA NTRS1978

Applications of the hybrid coordinate method to the TOPS autopilot

Preliminary results are presented from the application of the hybrid coordinate method to modeling TOPS (thermoelectric outer planet spacecraft) structural dynamics. Computer simulated responses of the vehicle are included which illustrate the interaction of relatively flexible appendages with an autopilot control system. Comparisons were made between simplified single-axis models of the control loop, with spacecraft flexibility represented by hinged rigid bodies, and a very detailed three-axis spacecraft model whose flexible portions are described by modal coordinates. While single-axis system, root loci provided reasonable qualitative indications of stability margins in this case, they were quantitatively optimistic when matched against responses of the detailed model.

Fleischer, G. E.↗
NASA NTRS1983

Observations of charging dynamics

Plasma data from the University of California, San Diego, Auroral Particles experiments on Applied Technology satellites 5 and 6 are used to investigate the dynamics of natural charging events. Both eclipse and daylight charging events are considered, and typical responses are illustrated by data from specific events. Two different physical processes are found to be involved in the charging process. One of these is straightforward: the spacecraft structure potential responds rapidly to changes in the environment, typically changing by hundreds of volts in a few seconds. The other process is more subtle: differential charging and potential barrier formation precede structural charging and determine the time scale; typically, tens of minutes are required for the potential to change by several hundreds of volts. The latter process is found to be predominately responsible for daylight charging on both spacecraft.

Olsen, R. C.↗
NASA NTRS1993

The Submillimeter Wave Astronomy Satellite (SWAS) solar array system

The SWAS (Submillimeter Wave Astronomy Satellite) solar array system is described. It is an innovative approach to meet the missions requirements. The SWAS satellite provides a three axis stabilized platform to survey a variety of galactic cloud structures. This system includes highly reliable, lightweight launch latch, deployment, and lock mechanisms, and solar array panels that provide the maximum solar cell area. The design of the solar arrays are the result of system trades that included instrument and spacecraft thermal constraints, attitude control system maneuvering rates and pointing accuracies, the power system, and the spacecraft structure.

Sneiderman, Gary↗
NASA NTRS2000

Nozomi Cis-Lunar Phase Orbit Determination

Japan's Institute of Space and Astronautical Science (ISAS) launched Nozomi, its first mission to the planet Mars using the newly developed M-V launch vehicle on July 3, 1998. Scientific objectives of the mission are to study the structure and dynamics of the Martian upper atmosphere and its interaction with the solar wind. Nozomi is a cooperative mission between ISAS and the National Aeronautics and Space Administration (NASA). The NASA contribution includes navigation and tracking services provided by the Jet Propulsion Laboratory (JPL). The spacecraft also serves as an engineering demonstration of basic technology for planetary exploration. One of the new technologies was a unique trajectory, developed by ISAS, which used solar gravitational perturbations at the weak stability boundary as an aid to achieve an Earth-Mars transfer orbit. This trajectory saves approximately 120 m/s of Delta V compared to direct hyperbolic insertion and is considered an enabling technology for the mission. Nozomi was the first spacecraft to employ this trajectory and provided on-orbit validation of the technique. The trajectory was achieved by initially placing the spacecraft in a highly elliptical cis-lunar phasing orbit. Six maneuvers were performed during this period to correct injection errors and target an outbound lunar swingby in September 1998. The gravity assist from the lunar swingby raised apogee to the vicinity of the weak stability boundary. After three more targeting maneuvers, Nozomi performed an inbound lunar swingby followed immediately by a powered Earth swingby in late December 1998. A 420 m/s Trans Mars Insertion (TMI) burn at the final Earth periapsis was intended to place the spacecraft on a heliocentric trajectory leading to Mars orbit insertion in October 1999. Orbit determination for Nozomi is performed in parallel by both ISAS and the Multi-Mission Navigation (MMNAV) group at JPL. This was an advantage for the mission because each group would generate solutions based on data collected from their respective tracking networks. Spacecraft events, such as sequence uplinks and maneuvers, were generally scheduled during passes at the Usuda tracking station in Japan. As a result, maneuver design and reconstruction was derived from MMNAV solutions based on JPL tracking data obtained immediately prior to or following maneuvers. Data was also exchanged between ISAS and MMNAV so orbit determination could be performed on joint data sets in support of critical targeting late in the cis-lunar phase. In this paper, information regarding the MMNAV orbit determination effort for the first six months of the mission is presented. The spacecraft trajectory is characterized first, followed by a discussion of the orbit determination estimation procedure and models. Results from selected orbit solutions are presented and compared against reconstructed trajectories. One area of emphasis in this paper is orbit determination in the vicinity of the weak stability boundary. Precise navigation was necessary to target the second lunar swingby and the powered Earth swingby. Delivery accuracy of 150 m was required for these critical encounters, but a number of factors contributed to the general degradation of orbit determination accuracy. This included the fact that the spacecraft was at apogee, at a range of 1.7 million km and moving at less than I km/sec perpendicular to the line of sight. Nozomi was also close to zero degrees declination where there are known limitations on orbit determination performance. Finally, S-band tracking data was acquired through the Nozomi backup low gain antenna. This antenna is offset from the axis of this spin stabilized spacecraft and superimposed large signatures in the Doppler and range data. These difficulties were overcome by combining long data arcs, spanning several maneuvers, with a high fidelity solar pressure model. The model included a physically accurate representation of the spacecraft structure and a high time resolution orientation model. Observation modeling included the removal of the spin induced Doppler bias, spin signature and per pass correction of range calibration errors applied for data leading up to critical events. As a result, all orbit determination goals were met. A second area of emphasis in this paper is the JPL tracking and orbit determination effort in support of the TMI maneuver. TMI occurred out of contact with ground stations and the JPL Goldstone tracking complex had the first pass following the bum. As a result, MMNAV had the responsibility to make a rapid assessment of the maneuver performance. MMNAV made the determination that a 100 m/s under bum had occurred and promptly informed ISAS via voice lines. ISAS immediately began preparations for a correction maneuver (TMIc), which had to be performed during the next Usuda pass. The near real time assessment by MMNAV provided accurate antenna frequency and pointing updates for the spacecraft acquisition at Usuda and the close coordination between the two agencies enabled the design and successful execution of the TMc maneuver. Propellant consumption during the correction burn dictated that the mission be redesigned. ISAS developed a new plan which adds 3 full solar orbits, two Earth swingbys and one lunar swingby with arrival at Mars in January 2004. The final Mars orbit will still enable the mission to achieve all of its science objectives.

Ryne, Mark↗
NASA NTRS1979

Predictions of induced background radiations at gamma/X-ray experiment envelopes in NASA spacecraft

This work seeks to predict secondary radiation levels induced in spacecraft structures by space protons. The radiations analyzed are secondary neutrons from spallation and evaporation reactions and gamma and beta rays from the decay of induced radioactivity, as sources of interfering background to spaceborne measurements of galactic and planetary gamma rays below 10 MeV. The spacecraft considered are the Multi-Mission Spacecraft (MMS) and the Space Shuttle, modeled as spherical shells. The proton environment is that of the South Atlantic Anomaly, as well as cosmic ray protons. The induced radioactivity is analyzed in terms of its interference with various gamma-ray lines of astrophysical interest, as well as its contribution to several spectral regions of the gamma-ray continuum. The buildup of the line and continuum radioactivity background is predicted for a period of nearly 9 months in orbit (approximately 4100 orbits). In addition, background contributions from cosmic ray electron bremsstrahlung and earth gamma-ray albedo are estimated.

Fischbein, W. L.↗
NASA NTRS2019

Review of Habitable Softgoods Inflatable Design, Analysis, Testing, and Potential Space Applications

Inflatable space structures have the potential to significantly reduce the required launch volume of large crewed pressure vessels for space exploration missions. Mass savings can also be achieved via the use of high specific strength softgoods materials, and the reduced design penalty from launching the structure in a densely packaged state. Inflatable softgoods structures have been investigated since the late 1950's, and several major development programs at NASA and in industry have helped advance the state-of-the-art in this technology area. This paper discusses the design, analysis, structural testing, and potential applications for inflatable softgoods structures. In particular, this paper will discuss the design of the multi-layer softgoods shell (inner layer, bladder, structural restraint layer, micrometeoroid orbital debris protection layers, thermal insulation layers, and atomic oxygen layer (for low earth orbit) and the results of material and module-level testing that has been conducted over the past two decades at NASA. Finally, the current utilization of expandable spacecraft structures is discussed, as well as potential future applications including airlocks and habitats on the Lunar Orbital Platform-Gateway, and the surface of the Moon and Mars.

Valle, Gerard↗
NASA NTRS2015

Investigation of HZETRN 2010 as a Tool for Single Event Effect Qualification of Avionics Systems - Part II

An accurate prediction of spacecraft avionics single event effect (SEE) radiation susceptibility is key to ensuring a safe and reliable vehicle. This is particularly important for long-duration deep space missions for human exploration where there is little or no chance for a quick emergency return to Earth. Monte Carlo nuclear reaction and transport codes such as FLUKA can be used to generate very accurate models of the expected in-flight radiation environment for SEE analyses. A major downside to using a Monte Carlo-based code is that the run times can be very long (on the order of days). A more popular choice for SEE calculations is the CREME96 deterministic code, which offers significantly shorter run times (on the order of seconds). However, CREME96, though fast and easy to use, has not been updated in several years and underestimates secondary particle shower effects in spacecraft structural shielding mass. Another modeling option to consider is the deterministic code HZETRN 20104, which includes updates to address secondary particle shower effects more accurately. This paper builds on previous work by Rojdev, et al. to compare the use of HZETRN 2010 against CREME96 as a tool to verify spacecraft avionics system reliability in a space flight SEE environment. This paper will discuss modifications made to HZETRN 2010 to improve its performance for calculating SEE rates and compare results with both in-flight SEE rates and other calculation methods.

Rojdev, Kristina↗
NASA NTRS1991

Development and testing of controller performance evaluation methodology for multi-input/multi-output digital control systems

Described here is the development and implementation of on-line, near real time controller performance evaluation (CPE) methods capability. Briefly discussed are the structure of data flow, the signal processing methods used to process the data, and the software developed to generate the transfer functions. This methodology is generic in nature and can be used in any type of multi-input/multi-output (MIMO) digital controller application, including digital flight control systems, digitally controlled spacecraft structures, and actively controlled wind tunnel models. Results of applying the CPE methodology to evaluate (in near real time) MIMO digital flutter suppression systems being tested on the Rockwell Active Flexible Wing (AFW) wind tunnel model are presented to demonstrate the CPE capability.

Pototzky, Anthony↗
NASA NTRS1978

Comments on the drag-free control of a solar probe relativity mission

A spacecraft going within 4 solar radii of the Sun experiences an acceleration up to .00005 g from solar radiation pressure, and significant (though smaller) accelerations from the solar wind. To obtain satisfactory information about relativistic effects and the Sun's quadrupole mass coefficient, these non-gravitational accelerations are either measured or compensated out by making the spacecraft drag free. A proof mass inside the spacecraft structure is shielded from the external forces, so that it follows a nearly ideal gravitational orbit, and a control system activates gas jets (or other translational forcers) to make the vehicle follow the mass. The problems are mechanizing the control laws and minimizing extraneous effects such as the self gravitational pull of the spacecraft. The extraneous forces can be averaged in one plane by having a spinning vehicle.

Everitt, C. W. F.↗
NASA NTRS2004

The Space Technology 5 Avionics System

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&DH) 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& 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 co-located 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.

Speer, Dave↗
NASA NTRS1983

Time domain response envelope for structural dynamic systems

A transient envelope solution method is developed for payload structural systems. This method requires that the external forcing functions are decomposed into a quasi-static part and a dynamic part. The amplitude of the peak envelope and the corresponding varying frequencies for the generalized forcing functions are required for obtaining the response envelopes. The proposed method is applied to a sample problem as well as a spacecraft structural system and the results are in good agreement with the exact solution.

Chen, J.-C.↗
NASA NTRS2016

Utilizing the ISS Mission as a Testbed to Develop Cognitive Communications Systems

The ISS provides an excellent opportunity for pioneering artificial intelligence software to meet the challenges of real-time communications (comm) link management. This opportunity empowers the ISS Program to forge a testbed for developing cognitive communications systems for the benefit of the ISS mission, manned Low Earth Orbit (LEO) science programs and future planetary exploration programs. In November, 1998, the Flight Operations Directorate (FOD) started the ISS Antenna Manager (IAM) project to develop a single processor supporting multiple comm satellite tracking for two different antenna systems. Further, the processor was developed to be highly adaptable as it supported the ISS mission through all assembly stages. The ISS mission mandated communications specialists with complete knowledge of when the ISS was about to lose or gain comm link service. The current specialty mandated cognizance of large sun-tracking solar arrays and thermal management panels in addition to the highly-dynamic satellite service schedules and rise/set tables. This mission requirement makes the ISS the ideal communications management analogue for future LEO space station and long-duration planetary exploration missions. Future missions, with their precision-pointed, dynamic, laser-based comm links, require complete autonomy for managing high-data rate communications systems. Development of cognitive communications management systems that permit any crew member or payload science specialist, regardless of experience level, to control communications is one of the greater benefits the ISS can offer new space exploration programs. The IAM project met a new mission requirement never previously levied against US space-born communications systems management: process and display the orientation of large solar arrays and thermal control panels based on real-time joint angle telemetry. However, IAM leaves the actual communications availability assessment to human judgment, which introduces unwanted variability because each specialist has a different core of experience with comm link performance. Because the ISS utilizes two different frequency bands, dynamic structure can be occasionally translucent at one frequency while it can completely interdict service at the other frequency. The impact of articulating structure on the comm link can depend on its orientation at the time it impinges on the link. It can become easy for a human specialist to cross-associate experience at one frequency with experience at the other frequency. Additionally, the specialist's experience is incremental, occurring one nine-hour shift at a time. Only the IAM processor experiences the complete 24x7x365 communications link performance for both communications links but, it has no "learning capability." If the IAM processor could be endowed with a cognitive ability to remember past structure-induced comm link outages, based on its knowledge of the ISS position, attitude, communications gear, array joint angles and tracking accuracy, it could convey such experience to the human operator. It could also use its learned communications link behaviors to accurately convey the availability of future communications sessions. Further, the tool could remember how accurately or inaccurately it predicted availability and correct future predictions based on past performance. The IAM tool could learn frequency-specific impacts due to spacecraft structures and pass that information along as "experience." Such development would provide a single artificial intelligence processor that could provide two different experience bases. If it also "knew" the satellite service schedule, it could distinguish structure blockage from schedule or planet blockage and then quickly switch to another satellite. Alternatively, just as a human operator could judge, a cognizant comm system based on the IAM model could "know" that the blockage is not going to last very long and continue tracking a comm satellite, waiting for it to track away from structure. Ultimately, once this capability was fully developed and tested in the Mission Control Center, it could be transferred on-orbit to support development of operations concepts that include more advanced cognitive communications systems. Future applications of this capability are easily foreseen because even more dynamic satellite constellations with more nodes and greater capability are coming. Currently, the ISS fully employs a 300 million bit-per-second (Mbps) return link for harvesting payload science. In the coming eighteen months, it will step up to 600 Mbps. Already there is talk of a 1.2 billion bit-per-second (Gbps) upgrade for the ISS and laser comm links have already been tested from the ISS. Every data rate upgrade mandates more complicated and sensitive communications equipment which implies greater expertise invested in the human operator. Future on-orbit cognizant comm systems will be needed to meet greater performance demands aboard larger, far more complicated spacecraft. In the LEO environment, the old-style one-satellite-per-spacecraft operations concept will give way to a new concept of a single customer spacecraft simultaneously using multiple comm satellites. Much more highly-dynamic manned LEO missions with decades of crew members potentially increase the demand for communications link performance. A cognizant on-board communications system will meet advanced communications demands from future LEO missions and future planetary missions. The ISS has fledgling components of future exploration programs, both LEO and planetary. Further, the Flight Operations Directorate, through the IAM project, has already begun to develop a communications management system that attempts to solve advanced problems ideally represented by dynamic structure impacting scheduled satellite service. With an earnest project to integrate artificial intelligence into the IAM processor, the ISS Program could develop a cognizant communications system that could be adapted and transferred to future on-orbit avionics designs.

Jackson, Dan↗
NASA NTRS2016

Utilizing the ISS Mission as a Testbed to Develop Cognitive Communications Systems

The ISS provides an excellent opportunity for pioneering artificial intelligence software to meet the challenges of real-time communications (comm) link management. This opportunity empowers the ISS Program to forge a testbed for developing cognitive communications systems for the benefit of the ISS mission, manned Low Earth Orbit (LEO) science programs and future planetary exploration programs. In November, 1998, the Flight Operations Directorate (FOD) started the ISS Antenna Manager (IAM) project to develop a single processor supporting multiple comm satellite tracking for two different antenna systems. Further, the processor was developed to be highly adaptable as it supported the ISS mission through all assembly stages. The ISS mission mandated communications specialists with complete knowledge of when the ISS was about to lose or gain comm link service. The current specialty mandated cognizance of large sun-tracking solar arrays and thermal management panels in addition to the highly-dynamic satellite service schedules and rise/set tables. This mission requirement makes the ISS the ideal communications management analogue for future LEO space station and long-duration planetary exploration missions. Future missions, with their precision-pointed, dynamic, laser-based comm links, require complete autonomy for managing high-data rate communications systems. Development of cognitive communications management systems that permit any crew member or payload science specialist, regardless of experience level, to control communications is one of the greater benefits the ISS can offer new space exploration programs. The IAM project met a new mission requirement never previously levied against US space-born communications systems management: process and display the orientation of large solar arrays and thermal control panels based on real-time joint angle telemetry. However, IAM leaves the actual communications availability assessment to human judgement, which introduces unwanted variability because each specialist has a different core of experience with comm link performance. Because the ISS utilizes two different frequency bands, dynamic structure can be occasionally translucent at one frequency while it can completely interdict service at the other frequency. The impact of articulating structure on the comm link can depend on its orientation at the time it impinges on the link. It can become easy for a human specialist to cross-associate experience at one frequency with experience at the other frequency. Additionally, the specialist's experience is incremental, occurring one nine-hour shift at a time. Only the IAM processor experiences the complete 24x7x365 communications link performance for both communications links but, it has no "learning capability." If the IAM processor could be endowed with a cognitive ability to remember past structure-induced comm link outages, based on its knowledge of the ISS position, attitude, communications gear, array joint angles and tracking accuracy, it could convey such experience to the human operator. It could also use its learned communications link behaviors to accurately convey the availability of future communications sessions. Further, the tool could remember how accurately or inaccurately it predicted availability and correct future predictions based on past performance. The IAM tool could learn frequency-specific impacts due to spacecraft structures and pass that information along as "experience." Such development would provide a single artificial intelligence processor that could provide two different experience bases. If it also "knew" the satellite service schedule, it could distinguish structure blockage from schedule or planet blockage and then quickly switch to another satellite. Alternatively, just as a human operator could judge, a cognizant comm system based on the IAM model could "know" that the blockage is not going to last very long and continue tracking a comm satellite, waiting for it to track away from structure. Ultimately, once this capability was fully developed and tested in the Mission Control Center, it could be transferred on-orbit to support development of operations concepts that include more advanced cognitive communications systems. Future applications of this capability are easily foreseen because even more dynamic satellite constellations with more nodes and greater capability are coming. Currently, the ISS fully employs its high-data-rate return link for harvesting payload science. In the coming months, it will double that data rate and is forecast to fully utilize that capability. Already there is talk of an upgrade that quadruples the current data rate allocated to ISS payload science before the end of its mission and laser comm links have already been tested from the ISS. Every data rate upgrade mandates more complicated and sensitive communications equipment which implies greater expertise invested in the human operator. Future on-orbit cognizant comm systems will be needed to meet greater performance demands aboard larger, far more complicated spacecraft. In the LEO environment, the old-style one-satellite-per-spacecraft operations concept will give way to a new concept of a single customer spacecraft simultaneously using multiple comm satellites. Much more highly-dynamic manned LEO missions with decades of crew members potentially increase the demand for communications link performance. A cognizant on-board communications system will meet advanced communications demands from future LEO missions and future planetary missions. The ISS has fledgling components of future exploration programs, both LEO and planetary. Further, the Flight Operations Directorate, through the IAM project, has already begun to develop a communications management system that attempts to solve advanced problems ideally represented by dynamic structure impacting scheduled satellite service. With an earnest project to integrate artificial intelligence into the IAM processor, the ISS Program could develop a cognizant communications system that could be adapted and transferred to future on-orbit avionics designs.

Jackson, Dan↗
NASA NTRS2014

Investigation of HZETRN 2010 as a Tool for Single Event Effect Qualification of Avionics Systems

NASA's future missions are focused on long-duration deep space missions for human exploration which offers no options for a quick emergency return to Earth. The combination of long mission duration with no quick emergency return option leads to unprecedented spacecraft system safety and reliability requirements. It is important that spacecraft avionics systems for human deep space missions are not susceptible to Single Event Effect (SEE) failures caused by space radiation (primarily the continuous galactic cosmic ray background and the occasional solar particle event) interactions with electronic components and systems. SEE effects are typically managed during the design, development, and test (DD&T) phase of spacecraft development by using heritage hardware (if possible) and through extensive component level testing, followed by system level failure analysis tasks that are both time consuming and costly. The ultimate product of the SEE DD&T program is a prediction of spacecraft avionics reliability in the flight environment produced using various nuclear reaction and transport codes in combination with the component and subsystem level radiation test data. Previous work by Koontz, et al.1 utilized FLUKA, a Monte Carlo nuclear reaction and transport code, to calculate SEE and single event upset (SEU) rates. This code was then validated against in-flight data for a variety of spacecraft and space flight environments. However, FLUKA has a long run-time (on the order of days). CREME962, an easy to use deterministic code offering short run times, was also compared with FLUKA predictions and in-flight data. CREME96, though fast and easy to use, has not been updated in several years and underestimates secondary particle shower effects in spacecraft structural shielding mass. Thus, this paper will investigate the use of HZETRN 20103, a fast and easy to use deterministic transport code, similar to CREME96, that was developed at NASA Langley Research Center primarily for flight crew ionizing radiation dose assessments. HZETRN 2010 includes updates to address secondary particle shower effects more accurately, and might be used as another tool to verify spacecraft avionics system reliability in space flight SEE environments.

Rojdev, Kristina↗
NASA NTRS1992

Quantitative ESD Guidelines for Charged Spacecraft Derived from the Physics of Discharges

Quantitative guidelines are proposed for Electrostatic Discharge (ESD) pulse shape on charged spacecraft. The guidelines are based on existing ground test data, and on a physical description of the pulsed discharge process. The guidelines are designed to predict pulse shape for surface charging and internal charging on a wide variety of spacecraft structures. The pulses depend on the area of the sample, its capacitance to ground, and the strength of the electric field in the vacuum adjacent to the charged surface. By knowing the pulse shape, current vs. time, one can determine if nearby circuits are threatened by the pulse. The quantitative guidelines might be used to estimate the level of threat to an existing spacecraft, or to redesign a spacecraft to reduce its pulses to a known safe level. The experiments which provide the data and the physics that allow one to interpret the data will be discussed, culminating in examples of how to predict pulse shape/size. This method has been used, but not confirmed, on several spacecraft.

Frederickson, A. R.↗
NASA NTRS2006

Micronuclei Induction in Human Fibroblasts Exposed In Vitro to Los Alamos High-Energy Neutrons

High-energy secondary neutrons, produced by the interaction of galactic cosmic rays with the atmosphere, spacecraft structure and planetary surfaces, contribute to a significant fraction to the dose equivalent in crew members and passengers during commercial aviation travel, and astronauts in space missions. The Los Alamos Nuclear Science Center (LANSCE) neutron facility#s ICE House 30L beamline is known to generate neutrons that simulate the secondary neutron spectra of earth#s atmosphere. The neutron spectrum is also similar to that measured onboard spacecraft like the MIR and International Space Station (ISS). To evaluate the biological damage, we exposed human fibroblasts in vitro to the LANSCE neutron beams without degrader at an entrance dose rate of 25 mGy/hr and analyzed the micronuclei (MN) induction. The cells were also placed behind a 9.9 cm water column to study effect of shielding in the protection of neutron induced damages. It was found that the dose response in the MN frequency was linear for the samples with and without shielding and the slope of the MN yield behind the shielding was reduced by a factor of 3.5. Compared to the MN induction in human fibroblasts exposed to a gamma source at a low dose rate, the RBE was found to be 16.7 and 10.0 for the neutrons without and with 9.9 cm water shielding, respectively.

Gersey, Brad↗
NASA NTRS2016

Robust Thermal Control of Propulsion Lines for Space Missions

In spacecraft that have propulsion lines that are located externally with open bus architecture, the lines are typically insulated by Multi Layer Insulation (MLI) blankets to protect them thermally from the cold space environment. In addition to heat loss through the insulation, mechanical supports used to attach the lines to the spacecraft structure also create heat leaks from the lines. These lines typically have very low thermal conduction in the axial direction, so the heat balance in the lines tends to be very local without much heat spreading. The typical allowable temperature range for hydrazine-based lines is +15/+50°C. This tight temperature range has to be maintained for every location on these lines. For typical spacecraft, these lines can be several meters long. Temperature control is typically achieved by closed loop monitoring of temperatures along the lines and the corresponding powering of the heaters in a bang-bang approach to maintain the temperatures within the dead band of the control loop. The temperatures of propulsion lines are a function of several parameters with heat loss characteristics of the MLI being the key one. Unfortunately, this same key characteristic (MLI effective emittance) has a large variation along its length due to its dependence on workmanship, which in turn leads to large uncertainties in the propulsion lines’ local temperatures. Because of the poor conduction along the axial direction, heat balance along the length varies dramatically from one location to the next, even few inches apart, depending on the combination of the controlling parameters. This paper describes various robust design and implementation approaches that have been investigated to greatly reduce the randomness associated with predicting the temperature of these propulsion lines.

Bhandari, Pradeep↗