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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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At least 739 records · Page 41
All sourcesNASA NTRSDOE OSTI
NASA NTRS2015

Propulsion Technology Demonstrator

NASA's Pathfinder Technology Demonstrator (PTD) project will test the operation of a variety of novel CubeSat technologies in low- Earth orbit, providing significant enhancements to the performance of these small and effective spacecraft. Each Pathfinder Technology Demonstrator mission consists of a 6-unit (6U) CubeSat weighing approximately 26 pounds (12 kilograms) and measuring 12 inches x 10 inches x 4 inches (30 centimeters x 25 centimeters x 10 centimeters), comparable in size to a common shoebox. CubeSats are a class of nanosatellites that use a standard size and form factor. The standard Cube- Sat size uses a "one unit" or "1U" measuring 4 inches x 4 inches x 4 inches (10x10x10 centimeters) and is extendable to larger sizes by "stacking" a number of the 1U blocks to form a larger spacecraft. Each PTD spacecraft will also be equipped with deployable solar arrays that provide an average of 44 watts of power while in orbit.

CubeSat↗
NASA NTRS2015

Precision Closed-Loop Orbital Maneuvering System Design and Performance for the Magnetospheric Multiscale Formation

NASAs Magnetospheric Multiscale (MMS) mission successfully launched on March 13,2015 (UTC) consists of four identically instrumented spin-stabilized observatories that function as a constellation to study magnetic reconnection in space. The need to maintain sufficiently accurate spatial and temporal formation resolution of the observatories must be balanced against the logistical constraints of executing overly-frequent maneuvers on a small fleet of spacecraft. These two considerations make for an extremely challenging maneuver design problem. This paper focuses on the design elements of a 6-DOF spacecraft attitude control and maneuvering system capable of delivering the high-precision adjustments required by the constellation designers specifically, the design, implementation, and on-orbit performance of the closed-loop formation-class maneuvers that include initialization, maintenance, and re-sizing. The maneuvering control system flown on MMS utilizes a micro-gravity resolution accelerometer sampled at a high rate in order to achieve closed-loop velocity tracking of an inertial target with arc-minute directional and millimeter-per second magnitude accuracy. This paper summarizes the techniques used for correcting bias drift, sensor-head offsets, and centripetal aliasing in the acceleration measurements. It also discusses the on-board pre-maneuver calibration and compensation algorithms as well as the implementation of the post-maneuver attitude adjustments.

Maneuver and Attitude Dynamics↗
NASA NTRS2015

Precision Closed-Loop Orbital Maneuvering System Design and Performance for the Magnetospheric Multi-Scale Mission (MMS) Formation

NASA's Magnetospheric Multi-Scale (MMS) mission successfully launched on March 13, 2015 (UTC) consists of four identically instrumented spin-stabilized observatories that function as a constellation to study magnetic reconnection in space. The need to maintain sufficiently accurate spatial and temporal formation resolution of the observatories must be balanced against the logistical constraints of executing overly-frequent maneuvers on a small fleet of spacecraft. These two considerations make for an extremely challenging maneuver design problem. This paper focuses on the design elements of a 6-DOF spacecraft attitude control and maneuvering system capable of delivering the high-precision adjustments required by the constellation designers---specifically, the design, implementation, and on-orbit performance of the closed-loop formation-class maneuvers that include initialization, maintenance, and re-sizing. The maneuvering control system flown on MMS utilizes a micro-gravity resolution accelerometer sampled at a high rate in order to achieve closed-loop velocity tracking of an inertial target with arc-minute directional and millimeter-per-second magnitude accuracy. This paper summarizes the techniques used for correcting bias drift, sensor-head offsets, and centripetal aliasing in the acceleration measurements. It also discusses the on-board pre-maneuver calibration and compensation algorithms as well as the implementation of the post-maneuver attitude adjustments.

Determination and Control↗
NASA NTRS2014

LADEE Multi-Domain Simulation

The Lunar Atmosphere Dust Environment Explorer (LADEE) was a small explorer class spacecraft that was launched on Sept 7, 2013 and that was de-orbited and successfully impacted the Moons surface on April 17, 2014 after completing all of the mission objectives. The low-cost rapidly prototyped hardware design used for the spacecraft was extend to the development of the software base. To achieve this goal, a Model Based Design approach was utilized to develop the onboard flight software, and out of this development a model based multipurpose simulator was created of the LADEE spacecraft and its mission environment. This simulator extended the traditional function of propagating the vehicle's kinematic and rotational states and included the electrical and thermal states propagation. Traditionally, these domains are handled by domain specific high fidelity simulations that use the states histories from other domains as input. By reducing the fidelity and abstracting the relevant features being monitored and controlled by the flight software, it was possible to model the coupling across these domains resulting in more accurate overall system behavior. A faster than real-time workstation (WSIM) version of the LADEE simulator was used to develop and test the software control algorithms in the Simulink environment. To maximize the performance of the simulation, modeling knobs were introduced to reduce the resolution of the some of the domains models when the effects of that domain were not significant for the scope of that simulation. The automatic code generation feature in Simulink was used to port the simulation to several real-time environments to support Processor-in-the-Loop (PIL) and Hardware-in-the-Loop (HIL) testing, verification and validation. The real-time environment required that the design of each of the domain models be deterministic as possible in the time required to perform all of the calculations to update its states. The simulation interface was designed to be compatible with the command interface employed by the LADEE mission operation team. The WSIM, PIL, and HIL simulators thus used a common interface and thus were used for flight software testing, for mission operations personnel training (nominal and off-nominal operations) prior to the mission and to perform command sequence verification during the mission. This presentation will look at the modeling strategies used to create a common interface to the simulator and to model and couple multiple domains within the simulation, the results of those strategies, and the lessons learned.

Multi-Domain↗
NASA NTRS2016

Electrospray Thruster Performance with Microfabricated Emitter Arrays and Indium Propellant

The feasibility of a microfabricated indium-fueled electrospray thruster with excellent performance was demonstrated. High efficiency electrospray thrusters with microfabricated components are under development for very compact, distributable propulsion systems that can be employed on both very small and large spacecraft with 10X improvement over SOA in mass, volume and specific impulse The critical components of this technology are the microfabricated emitter arrays and the capillary force driven indium feed system. Grey scale electron-beam lithography patterning and reactive ion dry etching provided the required micron-scale etch precision and uniformity across an array of 400 emitters in 1 cm2. Initial tests of single microfabricated silicon emitters demonstrated better performance than industry standard single emitters and the performance required to achieve 200 micronewtons when scaled up to a thruster with 400 emitters in 1 cm2. Arrays of emitters tested in a preliminary prototype Microfluidic Electrospray Propulsion (MEP) thruster assembly demonstrated stable performance at estimated thrust levels of 5, 50, and 120 µN at extraction voltages less than 4 kV. Current stability was within 0.15 % at 120 µN with only 1.5% of the emitter current collected by the extractor electrode. Post-test inspections revealed that more than 99% of the 400 emitters were electrospraying during the test. Specific impulse was estimated to be >3100 s from measurements of total charge and consumed indium mass at an emitter voltage of 1470 V. The results of this investigation suggest that microfabricated indium electrospray thruster technology is both feasible and capable of excellent performance as a highly compact microthruster.

Marrese-Reading, Colleen M.↗
NASA NTRS2019

InSight orbit determination

The InSight mission relied on accurate deep-space navigation for a successful Mars landing on November 26, 2018. In this paper, we discuss the role of the cruise Orbit Determination team, whose responsibilities included determining the spacecraft state, predicting the future trajectory, and quantifying the uncertainty associated with those estimates. In particular, we will focus on spacecraft dynamic modeling, small forces due to attitude control, radiometric tracking data, filter strategies, uncertainty quantification, and responses to unexpected flight situations. We will also provide analysis of reconstructed maneuvers, small forces, and delivery accuracy at Mars arrival.

Seubert, Jill↗
NASA NTRS2024

On the Mesoscale Structure of Coronal Mass Ejections at Mercury’s Orbit: BepiColombo and Parker Solar Probe Observations

On 2022 February 15, an impressive filament eruption was observed off the solar eastern limb from three remote-sensing viewpoints, namely, Earth, STEREO-A, and Solar Orbiter. In addition to representing the most-distant observed filament at extreme ultraviolet wavelengths—captured by Solar Orbiter's field of view extending to above 6 R ⊙ —this event was also associated with the release of a fast (∼2200 km s −1 ) coronal mass ejection (CME) that was directed toward BepiColombo and Parker Solar Probe. These two probes were separated by 2° in latitude, 4° in longitude, and 0.03 au in radial distance around the time of the CME-driven shock arrival in situ. The relative proximity of the two probes to each other and the Sun (∼0.35 au) allows us to study the mesoscale structure of CMEs at Mercury's orbit for the first time. We analyze similarities and differences in the main CME-related structures measured at the two locations, namely, the interplanetary shock, the sheath region, and the magnetic ejecta. We find that, despite the separation between the two spacecraft being well within the typical uncertainties associated with determination of CME geometric parameters from remote-sensing observations, the two sets of in situ measurements display some profound differences that make understanding the overall 3D CME structure particularly challenging. Finally, we discuss our findings within the context of space weather at Mercury's distance and in terms of the need to investigate solar transients via spacecraft constellations with small separations, which has been gaining significant attention during recent years.

Erika Palmerio↗
NASA NTRS1971

Experiment M403: Electrostatic charge

Experiment M403 was flown on the Gemini 4 and 5 spacecraft in an effort to discover whether the spacecraft could accumulate sufficient excess electric charge to constitute a hazard for later Gemini-Agena rendezvous missions. An attempt was made to measure the electric field terminating on a small area of the spacecraft retrograde adapter section and to calculate the total spacecraft charge from this measured field value. The Gemini 5 instrument was inhibited from responding to electric fields that corresponded to a charge on the spacecraft. However, readings were obtained on the Gemini 5 mission that were of as great or greater magnitude than those obtained on the Gemini 4 mission. This could be interpreted as an indication that the contribution of a net spacecraft charge to the Gemini 4 measurements was negligible. An operational measurement that was performed on a later rendezvous mission was indicative of a very small exchange of charge between the Gemini spacecraft and the Agena vehicle.

Lafferty, P. E.↗
NASA NTRS1992

Spacecraft design considerations for an Inner Magnetosphere Imager mission

Imaging the Earth's magnetosphere from space will enable scientists to better understand the global shape of the inner magnetosphere, its components and processes. The proposed Inner Magnetosphere Imager (IMI) mission will obtain the first simultaneous images of the component regions of the inner magnetosphere and will enable scientists to relate these global images to internal and external influences as well as local observations. NASA's Marshall Space Flight Center (MSFC) is performing a concept definition study of the proposed mission. As currently envisioned, the baseline mission calls for an instrument complement of approximately seven imagers to be flown in an elliptical Earth orbit with an apogee of seven Earth Radii (RE). Several spacecraft concepts have been examined for the mission. The baseline concept utilizes a spinning spacecraft with a despun platform, the second uses a three-axis stabilized spacecraft with a spinning platform, while the third option splits the instruments onto two small satellites; a spinning spacecraft and a complementary three-axis stabilized spacecraft. This paper will address the mission objectives, the rationale for using proven spacecraft designs, and the preliminary concept definition study team results for all three options.

Herrmann, Melody C.↗
NASA NTRS2015

Swarm Flyby Gravimetry

This study describes a new technology for discerning the gravity fields and mass distribution of a solar system small body, without requiring dedicated orbiters or landers. Instead of a lander, a spacecraft releases a collection of small, simple probes during a flyby past an asteroid or comet. By tracking those probes from the host spacecraft, one can estimate the asteroid's gravity field and infer its underlying composition and structure. This approach offers a diverse measurement set,equivalent to planning and executing many independent and unique flyby encounters of a single spacecraft. This report assesses a feasible hardware implementation, derives the underlying models,and analyzes the performance of this concept via simulation.In terms of hardware, a small, low mass, low cost implementation is presented, which consists of a dispenser and probes. The dispenser contains roughly 12 probes in a tube and has a total size commensurate with a 6U P-Pod. The probes are housed in disc shaped sabots. When commanded,the dispenser ejects the top-most probe using a linear motor. The ejected probe separates from its sabots and unfolds using internal springs. There are two types of probes, each designed for a particular tracking modality. The reflective probe type, tracked by a telescope, unfolds to forma diffusely reflective sphere. The retroreflector probe type, tracked by a lidar, unfolds to form a corner-cube retroreflector assembly. Both types are designed to spherical so that their attitude doesn't affect the spacecraft's tracking performance.This analysis indicates that the point-mass term of small bodies larger than roughly 500 m in diameter can be observed from a host spacecraft that tracks locally deployed probes throughout a flyby to an uncertainty of better than 5%. The conditions by which this measurement is possible depends on the characteristics of the asteroid (size, type), the flyby velocity, and the type of tracking available (angles-only or angles+ranging). For most encounters, a few (1-3) well placed probes can be very effective, with marginal improvement for additional probes. Given realistic deployment errors, an encounter may require roughly 10-12 probes to ensure that 1-3 achieve their target. Long duration tracking of probes flying by large asteroids (>5 km diameter) can sometimes provide observability of the gravity field's first spherical harmonic, J( sub 2). In summary, this method offers a feasible, affordable approach to enabling or augmenting flyby science.

Atchison, Justin↗
NASA NTRS2011

Self-Rupturing Hermetic Valve

For commercial, military, and aerospace applications, low-cost, small, reliable, and lightweight gas and liquid hermetically sealed valves with post initiation on/off capability are highly desirable for pressurized systems. Applications include remote fire suppression, single-use system-pressurization systems, spacecraft propellant systems, and in situ instruments. Current pyrotechnic- activated rupture disk hermetic valves were designed for physically larger systems and are heavy and integrate poorly with portable equipment, aircraft, and small spacecraft and instrument systems. Additionally, current pyrotechnically activated systems impart high g-force shock loads to surrounding components and structures, which increase the risk of damage and can require additional mitigation. The disclosed mechanism addresses the need for producing a hermetically sealed micro-isolation valve for low and high pressure for commercial, aerospace, and spacecraft applications. High-precision electrical discharge machining (EDM) parts allow for the machining of mated parts with gaps less than a thousandth of an inch. These high-precision parts are used to support against pressure and extrusion, a thin hermetically welded diaphragm. This diaphragm ruptures from a pressure differential when the support is removed and/or when the plunger is forced against the diaphragm. With the addition of conventional seals to the plunger and a two-way actuator, a derivative of this design would allow nonhermetic use as an on/off or metering valve after the initial rupturing of the hermetic sealing disk. In addition, in a single-use hermetically sealed isolation valve, the valve can be activated without the use of potential leak-inducing valve body penetrations. One implementation of this technology is a high-pressure, high-flow-rate rupture valve that is self-rupturing, which is advantageous for high-pressure applications such as gas isolation valves. Once initiated, this technology is self-energizing and requires low force compared to current pyrotechnic-based burst disk hermetic valves. This is a novel design for producing a single-use, self-rupturing, hermetically sealed valve for isolation of pressurized gas and/or liquids. This design can also be applied for single-use disposable valves for chemical instruments. A welded foil diaphragm is fully supported by two mated surfaces that are machined to micron accuracies using EDM. To open the valve, one of the surfaces is moved relative to the other to (a) remove the support creating an unsupported diaphragm that ruptures due to over pressure, and/or (b) produce tension in the diaphragm and rupture it.

Tucker, Curtis E., Jr.↗
NASA NTRS2004

Small Mercury Ion Clock for On-board Spacecraft Navigation

I.Small Ion Clock Approach and Heritage: a) No lasers, uwave cavities, cryogenics, atomic beams, etc. b) Ions are electrically shuttled between separate optical and microwave traps. II. Each trap is optimized for its task: quadrupole for optical state selection; multi-pole for microwave clock. a) Very good stability shown in USNO. Timescale running "open loop." III. "Open loop" operation means no self-measurements of frequency offsets: (Zeeman, ion temperature,... etc.) a) Fewer parts and procedures, produces stable output continuously. IV. Ion clock is not so sensitive to temperature fluctuations a) Measured u:nshielded temperature coefficient of few 10(exp -15) per C. b) No bulky temperature isolation needed.

Prestage, John D.↗
NASA NTRS2004

Flexible Shields for Protecting Spacecraft Against Debris

A report presents the concept of Flexshield a class of versatile, lightweight, flexible shields for protecting spacecraft against impacts by small meteors and orbiting debris. The Flexshield concept incorporates elements of, but goes beyond, prior spacecraft-shielding concepts, including those of Whipple shields and, more recently, multi-shock shields and multi-shock blankets. A shield of the Flexshield type includes multiple outer layers (called bumpers in the art) made, variously, of advanced ceramic and/or polymeric fibers spaced apart from each other by a lightweight foam. As in prior such shields, the bumpers serve to shock an impinging hypervelocity particle, causing it to disintegrate vaporize, and spread out over a larger area so that it can be stopped by an innermost layer (back sheet). The flexibility of the fabric layers and compressibility of the foam make it possible to compress and fold the shield for transport, then deploy the shield for use. The shield can be attached to a spacecraft by use of snaps, hook-and-pile patches, or other devices. The shield can also contain multilayer insulation material, so that it provides some thermal protection in addition to mechanical protection.

Christiansen, Eric L.↗
NASA NTRS1977

Cancellation of residual spacecraft accelerations for zero-G space physics experiments

The Drop Dynamics Module houses an acoustic positioning system which counteracts the effects of small accelerations of a spacecraft and thus allows long-term study of free-floating materials such as liquid drops. The acoustic positioning system provides an acoustic 'well' in the center of the experiment chamber. Data collection is by cinematographic photography. The module subsystems are discussed.

Stephens, J. B.↗
NASA NTRS2003

Quasi-Sun-Pointing of Spacecraft Using Radiation Pressure

A report proposes a method of utilizing solar-radiation pressure to keep the axis of rotation of a small spin-stabilized spacecraft pointed approximately (typically, within an angle of 10 deg to 20 deg) toward the Sun. Axisymmetry is not required. Simple tilted planar vanes would be attached to the outer surface of the body, so that the resulting spacecraft would vaguely resemble a rotary fan, windmill, or propeller. The vanes would be painted black for absorption of Solar radiation. A theoretical analysis based on principles of geometric optics and mechanics has shown that torques produced by Solar-radiation pressure would cause the axis of rotation to precess toward Sun-pointing. The required vane size would be a function of the angular momentum of the spacecraft and the maximum acceptable angular deviation from Sun-pointing. The analysis also shows that the torques produced by the vanes would slowly despin the spacecraft -- an effect that could be counteracted by adding specularly reflecting "spin-up" vanes.

Spilker, Thomas↗