Results for “LAUNCHING DEVICE” — Page 28

NASA NTRSDate not supplied

Modeling and Simulation of Tank Pressure Control using Zero-Boil Off Active Thermal Control for LOXSAT Technology Demonstration Mission

To-date, research and modeling of cryogenic fluid management technologies (CFM) for spaceflight has been limited to ground tests, short-duration zero-g simulations (e.g. drop towers), and small-scale experiments on-orbit. There has not been a large-scale flight demonstration of a flight-like system. As future NASA missions to take humans further from Earth will require large, in-space cryogenic propulsion vehicles, it is imperative to begin collecting flight data for these systems to accurately model and design future vehicles. To meet this goal, NASA awarded tipping point technology demonstration awards to Eta Space, Lockheed Martin, Space Exploration Technologies (SpaceX), and United Launch Alliance (ULA) to demonstrate on-orbit storage and transfer of cryogenic propellant. For its award, Eta Space is developing LOXSAT-1. It is a small satellite that will be launched on a Rocket Lab Electron rocket. The spacecraft consists of a Rocket Lab Photon spacecraft bus with a primary payload of a spherical liquid oxygen (LOX) storage tank with thermodynamic control systems. The primary objective of the mission is to demonstrate zero-boil-off storage of liquid oxygen. To accomplish this objective, the payload is equipped with an active thermal control system fluid loop that consists of propellant management device (PMD), positive-displacement pump, heat exchanger connected to a cryocooler, and a spray bar mixing injector. When the fluid loop is operating, fluid is drawn from the tank by the PMD and pumped through the heat exchanger, lowering the fluid temperature below the fluid temperature in the tank. This subcooled liquid is then injected back into the tank through the spray bar. The subcooled injected liquid has two effects. If it is sprayed into the ullage space, the injected liquid will form into jets or droplets and exchange heat with the ullage gas. This will cool and condense the gas, reducing the pressure in the tank. Additionally, the liquid that is not sprayed through the ullage, as well as any remaining liquid spray from the ullage, will rejoin the liquid mass of the tank, lowering the bulk temperature of the liquid. These combines effects provide for zero-boil off pressure control by lowering the tank pressure and the liquid saturation pressure simultaneously, ensuring the liquid stays subcooled. To model these complex mechanics and predict the performance of the active thermal control system, NASA is providing Eta Space with 3 parallel models of the tank. The approach of providing 3 different models allows for cross-checking and comparisons between the three to better understand how different modeling assumptions and selection semi-empirical factors affects the modeling result. Additionally, developing 3 models provides three different schemes for numerical simulation, providing confidence that results depict real physical phenomenon and not numerical quirks of the program. Within the tank thermodynamics, there are two primary areas of heat transfer we concern ourselves with: the heat transfer between the ullage space and the droplet spray, and between the ullage space and bulk liquid. For the droplet heat transfer, there are multiple sets of assumptions that can be made and correlations that can be used. Currently, two working models - the TankSIM model and Easy5 model – provide for an overview of the different approaches available. The TankSIM model and Easy5 model use two different models for droplet heating and evaporation that illustrate how the models use different types of mechanisms to arrive at the same answer. For the TankSIM model, droplets are treated as spheres of constant radius. Heat is transferred from the ullage to the droplet and warms the droplet until it reaches saturation, then the droplet begins evaporating and reducing its radius and mass. To calculate the heat transfer coefficient between the droplet and ullage, the Ranz-Marshall correlation is used. To determine the number of droplets in the ullage, a resident mass approach is used. This approach averages the number of droplets such that residuals at the start-up and shut-off of the spray bar cancel out. This same approach is used in the Easy5 model. The GFSSP model implements a linked list to track individual droplet “nodes” within the model. For the Easy5 model, the droplet is assumed to have an interface at a temperature equal to the saturation temperature corresponding to the pressure of the gas phase. The heat transfer from the gas to the interface and the interface to the droplet bulk is then calculated, and the net mass transfer between the droplet and interface is determined by performing an energy balance across the interface. For the gas side of the interface, the Ranz-Marshall correlation is used. For the liquid side of the interface, a variety of correlations were tried, including Kronig and Brink (1950) and effective conductivity models. As a result of these assumptions, the Easy5 model currently predicts faster depressurization, as at saturated vapor conditions, the heat transfer coeffect on the liquid side for the Easy5 model is greater than the heat transfer coefficient predicted by Ranz-Marshall used in the TankSIM code. This greater heat flux translates into faster condensation of the saturated ullage gas. At the bulk liquid to ullage interface, the models are in much closer agreement. Both models model the ullage as a sphere centered within the bulk liquid in the tank, and both use the energy-jumping boundary condition to model heat and mass transfer across the interface. There are slight differences in how the interface temperature is calculated, however. The Easy5 model assumes the temperature of the interface is equal to the saturation temperature associated with the pressure of the gas phase. The TankSIM model calculates this temperature with Alabovskii’s equation, which provides a correction factor for interface temperatures. Analysis tasks are focused on determining rates of depressurization within the tank during active cooling operation. To maintain net positive suction head at the pump inlet, the tank pressure cannot fall faster than the saturation pressure associated with the temperature of the bulk liquid. Additionally, there is interest in analyzing the performance of the loop at different pump speeds and cryocooler input powers. Adjusting the flowrate affects both the performance of the heat exchanger between the cryocooler and pumped liquid, and the heat transfer between the droplet spray and the ullage. Ideally, a pump speed and cryocooler power can be selected that will allow the tank to operate in zero-boil-off mode with a very narrow range of storage pressure.

zero boil-off↗

NASA NTRSDate not supplied

Modeling and Simulation of Tank Pressure Control using Zero-Boiloff Active Thermal Control for LOXSAT Technology Demonstration Mission

To-date, research and modeling of cryogenic fluid management technologies (CFM) for spaceflight has been limited to ground tests, short-duration zero-g simulations (e.g. drop towers), and small-scale experiments on-orbit. There has not been a large-scale flight demonstration of a flight-like system. As future NASA missions to take humans further from Earth will require large, in-space cryogenic propulsion vehicles, it is imperative to begin collecting flight data for these systems to accurately model and design future vehicles. To meet this goal, NASA awarded tipping point technology demonstration awards to Eta Space, Lockheed Martin, Space Exploration Technologies (SpaceX), and United Launch Alliance (ULA) to demonstrate on-orbit storage and transfer of cryogenic propellant. For its award, Eta Space is developing LOXSAT-1. It is a small satellite that will be launched on a Rocket Lab Electron rocket. The spacecraft consists of a Rocket Lab Photon spacecraft bus with a primary payload of a spherical liquid oxygen (LOX) storage tank with thermodynamic control systems. The primary objective of the mission is to demonstrate zero-boil-off storage of liquid oxygen. To accomplish this objective, the payload is equipped with an active thermal control system fluid loop that consists of propellant management device (PMD), positive-displacement pump, heat exchanger connected to a cryocooler, and a spray bar mixing injector. When the fluid loop is operating, fluid is drawn from the tank by the PMD and pumped through the heat exchanger, lowering the fluid temperature below the fluid temperature in the tank. This subcooled liquid is then injected back into the tank through the spray bar. The subcooled injected liquid has two effects. If it is sprayed into the ullage space, the injected liquid will form into jets or droplets and exchange heat with the ullage gas. This will cool and condense the gas, reducing the pressure in the tank. Additionally, the liquid that is not sprayed through the ullage, as well as any remaining liquid spray from the ullage, will rejoin the liquid mass of the tank, lowering the bulk temperature of the liquid. These combines effects provide for zero-boil off pressure control by lowering the tank pressure and the liquid saturation pressure simultaneously, ensuring the liquid stays subcooled. To model these complex mechanics and predict the performance of the active thermal control system, NASA is providing Eta Space with 3 parallel models of the tank. The approach of providing 3 different models allows for cross-checking and comparisons between the three to better understand how different modeling assumptions and selection semi-empirical factors affects the modeling result. Additionally, developing 3 models provides three different schemes for numerical simulation, providing confidence that results depict real physical phenomenon and not numerical quirks of the program. Within the tank thermodynamics, there are two primary areas of heat transfer we concern ourselves with: the heat transfer between the ullage space and the droplet spray, and between the ullage space and bulk liquid. For the droplet heat transfer, there are multiple sets of assumptions that can be made and correlations that can be used. Currently, two working models - the TankSIM model and Easy5 model – provide for an overview of the different approaches available. The TankSIM model and Easy5 model use two different models for droplet heating and evaporation that illustrate how the models use different types of mechanisms to arrive at the same answer. For the TankSIM model, droplets are treated as spheres of constant radius. Heat is transferred from the ullage to the droplet and warms the droplet until it reaches saturation, then the droplet begins evaporating and reducing its radius and mass. To calculate the heat transfer coefficient between the droplet and ullage, the Ranz-Marshall correlation is used. To determine the number of droplets in the ullage, a resident mass approach is used. This approach averages the number of droplets such that residuals at the start-up and shut-off of the spray bar cancel out. This same approach is used in the Easy5 model. The GFSSP model implements a linked list to track individual droplet “nodes” within the model. For the Easy5 model, the droplet is assumed to have an interface at a temperature equal to the saturation temperature corresponding to the pressure of the gas phase. The heat transfer from the gas to the interface and the interface to the droplet bulk is then calculated, and the net mass transfer between the droplet and interface is determined by performing an energy balance across the interface. For the gas side of the interface, the Ranz-Marshall correlation is used. For the liquid side of the interface, a variety of correlations were tried, including Kronig and Brink (1950) and effective conductivity models. As a result of these assumptions, the Easy5 model currently predicts faster depressurization, as at saturated vapor conditions, the heat transfer coeffect on the liquid side for the Easy5 model is greater than the heat transfer coefficient predicted by Ranz-Marshall used in the TankSIM code. This greater heat flux translates into faster condensation of the saturated ullage gas. At the bulk liquid to ullage interface, the models are in much closer agreement. Both models model the ullage as a sphere centered within the bulk liquid in the tank, and both use the energy-jumping boundary condition to model heat and mass transfer across the interface. There are slight differences in how the interface temperature is calculated, however. The Easy5 model assumes the temperature of the interface is equal to the saturation temperature associated with the pressure of the gas phase. The TankSIM model calculates this temperature with Alabovskii’s equation, which provides a correction factor for interface temperatures. Analysis tasks are focused on determining rates of depressurization within the tank during active cooling operation. To maintain net positive suction head at the pump inlet, the tank pressure cannot fall faster than the saturation pressure associated with the temperature of the bulk liquid. Additionally, there is interest in analyzing the performance of the loop at different pump speeds and cryocooler input powers. Adjusting the flowrate affects both the performance of the heat exchanger between the cryocooler and pumped liquid, and the heat transfer between the droplet spray and the ullage. Ideally, a pump speed and cryocooler power can be selected that will allow the tank to operate in zero-boil-off mode with a very narrow range of storage pressure.

zero boil-off↗

NASA NTRS1971

The Viking mission to Mars.

In 1975 the National Aeronautics and Space Administration will launch two unmanned scientific payloads to make soft landings on Mars in 1976. They will be placed in orbit around the planet to perform a reconnaissance before landing. The spacecraft will be separated into an orbiter, which is sent to the surface. During the entry into the Martian atmosphere, Viking will make measurements of the composition and atmospheric structure. After landing, two cameras will be used to visually characterize the site. A device will be deployed to obtain a sample of the Mars surface, which will be returned to the lander for biological tests and for chemical analysis. The atmosphere at the surface will be chemically analyzed as well as the wind velocity, the humidity, and changes in temperature and pressure.

Soffen, G. A.↗

NASA NTRS1990

Mission analysis for coincident measurements of the Upper Atmosphere Research Satellite and Space Shuttle experiments

The launch and orbital characteristics of Space Shuttle missions which will maximize the coincident measurement opportunities between the Atmospheric Laboratory for Applications and Science (ATLAS) and the Upper Atmosphere Research Satellite (UARS) experiments are presented. Updated UARS orbit conditions are used and several ATLAS orbit constraints are incorporated, and the study is expanded to include other classes of instruments on the two satellites. The effects of launch time, altitude, and orbit inclination on the number and latitudinal distribution of coincident measurement opportunities for the UARS and Shuttle experiments are investigated. Large numbers of coincident measurements are possible for UARS and ATLAS limb-viewing experiments, but their distribution is greatly dependent on the inclination of the satellites and initial equatorial crossing times. Results indicate that a great number of coincident measurements may be obtained between limb-scanning devices (300-400) and that observation of the extreme latitudes (70-80 degrees) is obtained with an ATLAS inclination of 57 degrees.

Harrison, Edwin F.↗

NASA NTRS2000

Energetic Combustion Devices for Aerospace Propulsion and Power

Chemical reactions have long been the mainstay thermal energy source for aerospace propulsion and power. Although it is widely recognized that the intrinsic energy density limitations of chemical bonds place severe constraints on maximum realizable performance, it will likely be several years before systems based on high energy density nuclear fuels can be placed into routine service. In the mean time, efforts to develop high energy density chemicals and advanced combustion devices which can utilize such energetic fuels may yield worthwhile returns in overall system performance and cost. Current efforts in this vein are being carried out at NASA MSFC under the direction of the author in the areas of pulse detonation engine technology development and light metals combustion devices. Pulse detonation engines are touted as a low cost alternative to gas turbine engines and to conventional rocket engines, but actual performance and cost benefits have yet to be convincingly demonstrated. Light metal fueled engines also offer potential benefits in certain niche applications such as aluminum/CO2 fueled engines for endo-atmospheric Martian propulsion. Light metal fueled MHD generators also present promising opportunities with respect to electric power generation for electromagnetic launch assist. This presentation will discuss the applications potential of these concepts with respect to aero ace propulsion and power and will review the current status of the development efforts.

Litchford, Ron J.↗

NASA NTRS1988

Cryogenic, high speed, turbopump bearing cooling requirements

Although the Space Shuttle Main Engine (SSME) has repeatedly demonstrated the capability to perform during launch, the High Pressure Oxidizer Turbopump (HPOTP) main shaft bearings have not met their 7.5 hour life requirement. A tester is being employed to provide the capability of subjecting full scale bearings and seals to speeds, loads, propellants, temperatures, and pressures which simulate engine operating conditions. The tester design permits much more elaborate instrumentation and diagnostics than could be accommodated in an SSME turbopump. Tests were made to demonstrate the facilities; and the devices' capabilities, to verify the instruments in its operating environment and to establish a performance baseline for the flight type SSME HPOTP Turbine Bearing design. Bearing performance data from tests are being utilized to generate: (1) a high speed, cryogenic turbopump bearing computer mechanical model, and (2) a much improved, very detailed thermal model to better understand bearing internal operating conditions. Parametric tests were also made to determine the effects of speed, axial loads, coolant flow rate, and surface finish degradation on bearing performance.

Dolan, Fred J.↗

NASA NTRS1990

Feasibility study for a Cryogenic On-Orbit Liquid Depot-Storage, Acquisition and Transfer (COLD-SAT) satellite

This feasibility study presents the conceptual design of a spacecraft for performing a series of cryogenic fluid management flight experiments. This spacecraft, the Cryogenic On-Orbit Liquid Depot-Storage, Acquisition, and Transfer (COLD-SAT) satellite, will use liquid hydrogen as the test fluid, be launched on a Delta expendable launch vehicle, and conduct a series of experiments over a two to three month period. These experiments will investigate the physics of subcritical cryogens in the low gravity space environment to characterize their behavior and to correlate the data with analytical and numerical models of in-space cryogenic fluid management systems. Primary technologies addressed by COLD-SAT are: (1) pressure control; (2) chilldown; (3) no-vent fill; (4) liquid acquisition device fill; (5) pressurization; (6) low-g fill and drain; (7) liquid acquisition device expulsion; (8) line chilldown; (9) thermodynamic state control; and (10) fluid dumping.

Rybak, S. C.↗

NASA NTRS1975

The Skylab Parasol

An alternate thermal protection system for the Skylab space station cluster was deemed necessary shortly after launch when severe damage occured to the orbital workshop as it passed through the period of maximum dynamic pressure. Data indicated that the combination micrometeoroid and thermal shield on the sun facing side of the vehicle had been destroyed. The simultaneous design, fabrication, and test of the Skylab Parasol System which was completed in 6 days of around-the-clock effort is described. The Skylab Parasol design was selected as the primary sunshade device based on simplicity of design, ability to fabricate quickly, and early demonstration of a working prototype which featured semiautomatic deployment from within the workshop under shirt sleeve environment conditions.

Kinzler, J. A.↗

NASA NTRS2003

Power Electronics Being Developed for Deep Space Cryogenic Applications

Electronic circuits and systems designed for deep space missions need to operate reliably and efficiently in harsh environments that include very low temperatures. Spacecraft that operate in such cold environments carry a large number of heaters so that the ambient temperature for the onboard electronics remains near 20 C. Electronics that can operate at cryogenic temperatures will simplify system design and reduce system size and weight by eliminating the heaters and their associated structures. As a result, system development and launch cost will be reduced. At the NASA Glenn Research Center, an ongoing program is focusing on the development of power electronics geared for deep space low-temperature environments. The research and development efforts include electrical components design, circuit design and construction, and system integration and demonstration at cryogenic temperatures. Investigations are being carried out on circuits and systems that are targeted for use in NASA missions where low temperatures will be encountered: devices such as ceramic and tantalum capacitors, metal film resistors, semiconductor switches, magnetics, and integrated circuits including dc/dc converters, operational amplifiers, voltage references, and motor controllers. Test activities cover a wide range of device and circuit performance under simple as well as complex test conditions, such as multistress and thermal cycling. The effect of low-temperature conditions on the switching characteristics of an advanced silicon-on-insulator field effect transistor is shown. For gate voltages (VGS) below 2.6 V, drain currents at -190 C are lower than drain currents at room temperature (20 C).

Patterson, Richard L.↗

NASA NTRS1994

Thermal Design of the Cassini Narrow Angle Camera

The Narrow Angle Camera (NAC) is one of two cameras in the Imaging Science Subsystem (ISS) on the Cassini Spacecraft (S/C), the second camera is a Voyager-inherited Wide Angle Camera (WAC). Cassini is currently planned to be launched in October 1997 and will arrive at Saturn for a four year tour in June 2004. The Narrow Angle Optics are a Ritchey Chretien type Optics, has a focal length of 2000 mm, a relative aperture of f/10.5, a spectral range of 200 to 1100 nm, 24 filters, a pixel field of view of 6.0 microradian/pixel, and has a field of view of 3.5 x 3.5 degrees. The sensor is a Charged Couple Device (CCD), 1024 x 1024 pixels with a pixel size of 12 x 12 um, a full well greater than 50,000 e-, on chip processing of up to 800,000 e- pixel summation, a dark current of less than 0.1 e-/pixel/sec at operating temperature, and a charge transfer efficiency of 0.99999 at operating temperature...

Hoffman, Pamela↗

NASA NTRS2003

Lightning Instrumentation at KSC

This report summarizes lightning phenomena with a brief explanation of lightning generation and lightning activity as related to KSC. An analysis of the instrumentation used at launching Pads 39 A&B for measurements of lightning effects is included with alternatives and recommendations to improve the protection system and upgrade the actual instrumentation system. An architecture for a new data collection system to replace the present one is also included. A novel architecture to obtain lightning current information from several sensors using only one high speed recording channel while monitoring all sensors to replace the actual manual lightning current recorders and a novel device for the protection system are described.

Colon, Jose L.↗

NASA NTRS2003

Using CFD as Rocket Injector Design Tool: Recent Progress at Marshall Space Flight Center

The choice of tools used for injector design is in a transitional phase between exclusive reliance on the empirically based correlations and extensive use of computational fluid dynamics (CFD). The Next Generation Launch Technology (NGLT) Program goals emphasizing lower costs and increased reliability have produced a need to enable CFD as an injector design tool in a shorter time frame. This is the primary objective of the Staged Combustor Injector Technology Task currently under way at Marshall Space Flight Center (MSFC). The documentation of this effort begins with a very brief status of current injector design tools. MSFC's vision for use of CFD as a tool for combustion devices design is stated and discussed with emphasis on the injector. The concept of the Simulation Readiness Level (SRL), comprised of solution fidelity, robustness and accuracy, is introduced and discussed. This quantitative measurement is used to establish the gap between the current state of demonstrated capability and that necessary for regular use in the design process. MSFC's view of the validation process is presented and issues associated with obtaining the necessary data are noted and discussed. Three current experimental efforts aimed at generating validation data are presented. The importance of uncertainty analysis to understand the data quality is also demonstrated. First, a brief status of current injector design tools is provided as context for the current effort. Next, the MSFC vision for using CFD as an injector design tool is stated. A generic CFD-based injector design methodology is also outlined and briefly discussed. Three areas where MSFC is using injector CFD analyses for program support will be discussed. These include the Integrated Powerhead Development (IPD) engine which uses hydrogen and oxygen propellants in a full flow staged combustion (FFSC) cycle and the TR-107 and the RS84 engine both of which use RP-1 and oxygen in an ORSC cycle. Finally, an attempt is made to objectively summarize what progress has been made at MSFC in enabling CFD as an injector design tool.

Tucker, Kevin↗

NASA NTRS2020

State of the Art Small Spacecraft Technology

When the first edition of NASA’s Small Spacecraft Technology State-of-the-art report was published in 2013, 247 CubeSats and 105 other non-CubeSat small spacecraft under 50 kilograms (kg) had been launched, and these represented less than 2% of launched mass into orbit. By 2019, small spacecraft with mass less than 180 kg made up almost 7% of all mass launched into orbit. Additionally, 63% of spacecraft under 600 kg had mass less than 180 kg and of those 47% were CubeSats (1). Since 2013, flight heritage for small spacecraft, primarily CubeSats, has nearly doubled and with dedicated smallsat launch capabilities readily available and expanding, opportunities to demonstrate new technologies and systems are expected to increase. The 2020 edition of this report captures and distills the wealth of new information available on small spacecraft systems from NASA and other publicly available sources. Overall, this report is a survey of small spacecraft technologies sourced from open literature; it does not endeavor to be an original source, and only considers literature in the public domain to identify and classify devices. Commonly used sources for data include manufacturer datasheets, press releases, conference papers, journal papers, public filings with government agencies, news articles, presentations, and Small Spacecraft Systems Virtual Institute Federated Search. Data not appropriate for public dissemination, such as proprietary, export controlled, or otherwise restricted data, are not considered. As a result, this report includes many dedicated hours of desk research performed by subject matter experts reviewing resources noted above.

Bruce Yost↗

NASA NTRS2016

Biomolecule Sequencer: Next-Generation DNA Sequencing Technology for In-Flight Environmental Monitoring, Research, and Beyond

On the International Space Station (ISS), technologies capable of rapid microbial identification and disease diagnostics are not currently available. NASA still relies upon sample return for comprehensive, molecular-based sample characterization. Next-generation DNA sequencing is a powerful approach for identifying microorganisms in air, water, and surfaces onboard spacecraft. The Biomolecule Sequencer payload, manifested to SpaceX-9 and scheduled on the Increment 4748 research plan (June 2016), will assess the functionality of a commercially-available next-generation DNA sequencer in the microgravity environment of ISS. The MinION device from Oxford Nanopore Technologies (Oxford, UK) measures picoamp changes in electrical current dependent on nucleotide sequences of the DNA strand migrating through nanopores in the system. The hardware is exceptionally small (9.5 x 3.2 x 1.6 cm), lightweight (120 grams), and powered only by a USB connection. For the ISS technology demonstration, the Biomolecule Sequencer will be powered by a Microsoft Surface Pro3. Ground-prepared samples containing lambda bacteriophage, Escherichia coli, and mouse genomic DNA, will be launched and stored frozen on the ISS until experiment initiation. Immediately prior to sequencing, a crew member will collect and thaw frozen DNA samples, connect the sequencer to the Surface Pro3, inject thawed samples into a MinION flow cell, and initiate sequencing. At the completion of the sequencing run, data will be downlinked for ground analysis. Identical, synchronous ground controls will be used for data comparisons to determine sequencer functionality, run-time sequence, current dynamics, and overall accuracy. We will present our latest results from the ISS flight experiment the first time DNA has ever been sequenced in space and discuss the many potential applications of the Biomolecule Sequencer for environmental monitoring, medical diagnostics, higher fidelity and more adaptable Space Biology Human Research Program investigations, and even life detection experiments for astrobiology missions.

Sequencer↗

NASA NTRS1989

Liquid Angular-Momentum Compensator

Report discusses use of fluid-loop reaction ring as part of system orienting spacecraft. Proposed device imparts angular rotation to spacecraft by reacting against liquid contained in loop. Pump, or pumps, provide impetus to both spacecraft and fluid. Hydraulic accumulators and valves added to control flow. Technique offers better control than attitude-control thrusters. Several advantages in applications otherwise requiring large, rigid reaction wheel: Fluid loop need occupy only peripheral circulation path; does not necessarily require motor sized for maximum torque; does not require difficult-to-make bearings specified to withstand high launch-acceleration loads, provide high stiffness, operate smoothly, and exert minimal fractional torque. Unlike reaction wheel, fluid loop not balanced dynamically.

Iskenderian, Theodore C.↗

NASA NTRS2001

Development of Electronics for Low-Temperature Space Missions

Electronic systems that are capable of operating at cryogenic temperatures will be needed for many future NASA space missions, including deep space probes and spacecraft for planetary surface exploration. In addition to being able to survive the harsh deep space environment, low-temperature electronics would help improve circuit performance, increase system efficiency, and reduce payload development and launch costs. Terrestrial applications where components and systems must operate in low-temperature environments include cryogenic instrumentation, superconducting magnetic energy storage, magnetic levitation transportation systems, and arctic exploration. An ongoing research and development project for the design, fabrication, and characterization of low-temperature electronics and supporting technologies at NASA Glenn Research Center focuses on efficient power systems capable of surviving in and exploiting the advantages of low-temperature environments. Supporting technologies include dielectric and insulating materials, semiconductor devices, passive power components, optoelectronic devices, and packaging and integration of the developed components into prototype flight hardware. An overview of the project is presented, including a description of the test facilities, a discussion of selected data from component testing, and a presentation of ongoing research activities being performed in collaboration with various organizations.

Patterson, Richard L.↗

NASA NTRS2018

The Hayabusa2 Sample Return Mission and Its Small Carry-On Impactor Experiment

On December 3, 2014 the Hayabusa2 Spacecraft was launched by JAXA to investigate and return samples from carbonaceous Asteroid (162173) Ryugu. The spacecraft's payload contains several remote sensing instruments that are designed to characterize the asteroid's surface and to help identify scientifically interesting locations from which samples can be obtained safely. The data collected by these instruments will also help place any obtained samples in context with respect to the global characteristics of Ryugu. In addition to the suite of remote sensing instruments and the sampling device, Hayabusa2 also includes a small carry-on impactor (SCI) payload experiment, which is designed to impact the surface of the asteroid and excavate fresh material for subsequent collection. The SCI consists of a 30 cm diameter 2 kg copper disk that is accelerated by an explosive charge into the surface of Ryugu at approximately 2 km/s. The resultant crater formed during this experiment is estimated to be approximately between 2 - 10 meters in diameter with an ejecta blanket extending to 10 - 100 meters on the asteroid. Hayabusa2 will not be able to observe the impact experiment directly due to considerations of spacecraft safety since there is some risk of collisions with debris from the explosion and ejecta from the crater. However, prior to moving safely away, the spacecraft will deploy a small imaging system called Deployable Camera 3 (DCAM3). DCAM3 will detach from Hayabusa2 just before the detonation of the SCI and will image the moment of impact, crater formation, and ejecta propagation. These data will be transmitted to Hayabusa2 and will help constrain the near-surface properties of Ryugu and also impact scaling laws for small bodies in low gravity environments. These data could also be useful for gaining insights into determining asteroid physical properties from higher velocity impact events (e.g., DART and HERA data of Didymos).

Abell, Paul A.↗

NASA NTRS2023

On-Board Deployment Event Verification for GOES-R Spacecraft

As is common with many spacecraft designs, the GOES-R vehicles require a series of deployment events to transition from the launch configuration to the operational configuration. Rather than implementing additional sensors to verify various deployments, the GOES-R program developed an alternate approach that uses existing gyro rate sensing. The approach includes two pieces: the first is a new onboard shock detection capability to confirm initiation of individual deployment events, and the second is a ground-based dynamics verification step to confirm completion of deployment events. We first present the algorithm that detects shock events for various deployment devices along with the parameter tuning performed during ground tests. We then show inflight performance of the shock detection algorithm. While shock detection is useful for observing initiation of deployment events, completion of some deployment events cannot be determined by shock detection alone. For these events, such as solar panel latch-up and deployable boom extension, the program developed dynamics models for the deployment transient responses. High-rate gyro data were recorded for these events, which allow the ground team to verify that the appendages were fully deployed. We show the predictive models for these events and corresponding flight results.

GeoXO↗