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NASA NTRS2014

Micro Cathode Arc Thruster for PhoneSat: Development and Potential Applications

NASA Ames Research Center and the George Washington University are developing an electric propulsion subsystem that will be integrated into the PhoneSat bus. Experimental tests have shown a reliable performance by firing three different thrusters at various frequencies in vacuum conditions. The interface consists of a microcontroller that sends a trigger pulse to the Pulsed Plasma Unit that is responsible for the thruster operation. A Smartphone is utilized as the main user interface for the selection of commands that control the entire system. The propellant, which is the cathode itself, is a solid cylinder made of Titanium. This simplicity in the design avoids miniaturization and manufacturing problems. The characteristics of this thruster allow an array of μCATs to perform attitude control and orbital correction maneuvers that will open the door for the implementation of an extensive collection of new mission concepts and space applications for CubeSats. NASA Ames is currently working on the integration of the system to fit the thrusters and the PPU inside a 1.5U CubeSat together with the PhoneSat bus. This satellite is intended to be deployed from the ISS in 2015 and test the functionality of the thrusters by spinning the satellite around its long axis and measure the rotational speed with the phone gyros. This test flight will raise the TRL of the propulsion system from 5 to 7 and will be a first test for further CubeSats with propulsion systems, a key subsystem for long duration or interplanetary small satellite missions.

CUBESAT↗

NASA NTRS1992

Spacecraft flight control system design selection process for a geostationary communication satellite

The Earth's first artificial satellite, Sputnik 1, slowly tumbled in orbit. The first U.S. satellite, Explorer 1, also tumbled out of control. Now, as we launch the Mars observer and the Cassini spacecraft, stability and control have become higher priorities. The flight control system design selection process is reviewed using as an example a geostationary communication satellite which is to have a life expectancy of 10 to 14 years. Disturbance torques including aerodynamic, magnetic, gravity gradient, solar, micrometeorite, debris, collision, and internal torques are assessed to quantify the disturbance environment so that the required compensating torque can be determined. Then control torque options, including passive versus active, momentum control, bias momentum, spin stabilization, dual spin, gravity gradient, magnetic, reaction wheels, control moment gyros, nutation dampers, inertia augmentation techniques, three-axis control, reactions control system (RCS), and RCS sizing, are considered. A flight control system design is then selected and preliminary stability criteria are met by the control gains selection.

Barret, C.↗

NASA NTRS2002

Thermal Examination of an Orbiting Cryogenic Fuel Depot

For many years NASA has been interested in the storage and transfer of cryogenic fuels in space. Lunar, L2 and other chemical propulsive space vehicle missions now have staged refueling needs that a fuel depot would satisfy. The depot considered is located in lower earth orbit. Many considerations must go into designing and building such a station. Multi-layer insulation systems, thermal shielding and low conductive structural supports are the principal means of protecting the system from excessive heat loss due to boiloff. This study focuses on the thermal losses associated with storing LH2 in a passively cooled fuel depot in a lower earth equatorial orbit. The corresponding examination looks at several configurations of the fuel depot. An analytical model has been developed to determine the thermal advantages and disadvantages of three different fuel depot configurations. Each of the systems consists of three Boeing rocket bodies arranged in various configurations. The first two configurations are gravity gradient stabilized while the third one is a spin-stabilized concept. Each concept was chosen for self-righting capabilities as well as the fuel settling capabilities, however the purpose of this paper is to prove which of the three concepts is the most efficient passively cooled system. The specific areas to be discussed are the heating time from the fusion temperature to the vaporization temperature and the amount of boiloff for a specific number of orbits. Each of the previous points is compared using various sun exposed surface areas of the tanks.

Hull, Patrick V.↗

NASA NTRS2004

Space Technology 5: Pathfinder for Future Micro-Sat Constellations

The Space Technology 5 (ST-5) Project, currently in the implementation phase, is part of the National Aeronautics and Space Administration (NASA) s New Millennium Program (NMP). ST-5 will consist of a constellation of three miniature satellites, each with mass less than 25 kg and size approximately 60 cm by 30 cm. ST-5 addresses technology challenges, as well as fabrication, assembly, test and operations strategies for future micro-satellite missions. ST-5 will be deployed into a highly eccentric, geo-transfer orbit (GTO). This will expose the spacecraft to a high radiation environment as well as provide a low level magnetic background. A three-month flight demonstration phase is planned to validate the technologies and demonstrate concepts for future missions. Each ST-5 spacecraft incorporates NMP competitively-selected breakthrough technologies. These include Cold Gas Micro-Thrusters for propulsion and attitude control, miniature X-band transponder for space-ground communications, Variable Emittance Coatings for dynamic thermal control, and CULPRiT ultra low power logic chip used for Reed-Solomon encoding. The ST-5 spacecraft itself is a technology that can be infused into future missions. It is a fully functional micro-spacecraft built within tight volume and mass constraints. It is built to withstand a high radiation environment, large thermal variations, and high launch loads. The spacecraft power system is low-power and low-voltage, and is designed to turn on after separation &om the launch vehicle. Some of the innovations that are included in the ST-5 design are a custom spacecraft deployment structure, magnetometer deployment boom, nutation damper, X-band antenna, miniature spinning sun sensor, solar array with triple junction solar cells, integral card cage assembly containing single card Command and Data Handling and Power System Electronics, miniature magnetometer, and lithium ion battery. ST-5 will demonstrate the ability of a micro satellite to perform research-quality science. Each ST-5 spacecraft will deploy a precision magnetometer to be used both for attitude determination and as a representative science instrument. The spacecraft has been developed with a low magnetic signature to avoid interference with the magnetometer. The spacecraft will be able to detect and respond autonomously to science events, i.e. significant changes in the magnetic field measurements. The three spacecraft will be a pathfinder for future constellation missions. They will be deployed to demonstrate an appropriate geometry for scientific measurements as a constellation. They will be operationally managed as a constellation, demonstrating automation and communication strategies that will be useful for future missions. The technologies and future mission concepts will be validated both on the ground and in space. Technologies will be validated on the ground by a combination of component level and system level testing of the flight hardware in a thermal vacuum environment. In flight, specific validation runs are planned for each of the technologies. Each validation run consists of one or more orbits with a specific validation objective. This paper will describe the ST-5 mission, and the applicability of the NMP technologies, spacecraft, and mission concepts to future missions. It will also discuss the validation approach for the ST-5 technologies and mission concepts.

Carlisle, Candace↗

NASA NTRS2009

Misaligned Spin and Orbital Axes Cause the Anomalous Precession of DI Herculis

In this case we applied our Rossiter-McLaughlin methodology to a binary star, rather than a star-planet system. The orbits of binary stars precess as a result of general relativistic effects, forces arising from the asphericity of the stars, and forces from any additional stars or planets in the system. For most binaries, the theoretical and observed precession rates are in agreement. However, one system known as DI Herculis has resisted explanation for 30 years. The observed precession rate is a factor of four slower than the theoretical rate, a disagreement that once was interpreted as evidence for a failure of general relativity. Among the contemporary explanations are the existence of a circumbinary planet and a large tilt of the stellar spin axes with respect to the orbit. In this paper we reported that both stars of DI Herculis rotate with their spin axes nearly perpendicular to the orbital axis (contrary to the usual assumption for close binary stars). The rotationally induced stellar oblateness causes precession in the direction opposite to that of relativistic precession, thereby reconciling the theoretical and observed rates.

Albrecht, Simon↗

NASA NTRS1992

Diabatic dynamic initialization

A generalized dynamical adjustment procedure was applied to a diabatic model to produce balanced initial conditions. Namely, backward adiabatic model integration is followed by forward diabatic model integration, with a high frequency (low pass) filter in the form of the Euler-backward time differencing scheme being applied throughout the whole integration. As a result of the application of such a diabatic dynamic initialization procedure within the Goddard Laboratory for Atmospheres (GLA) 4-D data assimilation system, the following properties of forecasts from initialized fields are achieved right from the beginning of the usual forecast integration: (1) the forecast tendencies (and fields) are free of any noise due to imbalance in initial conditions; (2) the shocks related to an initial imbalance between model physics and dynamics, and especially the substantial initial imbalance of precipitation and evaporation fields, or the initial spin-up effect, are practically removed. Diabatic dynamic initialization has been compared with implicit nonlinear normal mode initialization, and found to be superior in removing the initial spin-up effect, and also in improving the tropical structure. The diabatic dynamic initialization procedure was successfully tested for the GLA system with the use of all conventional data and the GLA satellite data retrievals. It allows a smooth data insertion without any shocks or imbalances, which is highly desirable for efficient functioning of 4-D data assimilation systems. The developed initialization procedure is computationally efficient and in principle easily applicable to different forecast models.

Fox-Rabinovitz, Michael S.↗

NASA NTRS2002

Mission Applicability Assessment of Integrated Power Components and Systems

The need for smaller lightweight autonomous power systems has recently increased with the increasing focus on micro- and nanosatellites. Small area high-efficiency thin film batteries and solar cells are an attractive choice for such applications. The NASA Glenn Research Center, Johns Hopkins Applied Physics Laboratory, Lithium Power Technologies, MicroSat Systems, and others, have been working on the development of autonomous monolithic packages combining these elements or what are called integrated power supplies (IPS). These supplies can be combined with individual satellite components and are capable of providing continuous power even under intermittent illumination associated with a spinning or Earth orbiting satellite. This paper discusses the space mission applicability, benefits, and current development efforts associated with integrated power supply components and systems. The characteristics and several mission concepts for an IPS that combines thin-film photovoltaic power generation with thin-film lithium ion energy storage are described. Based on this preliminary assessment, it is concluded that the most likely and beneficial application of an IPS will be for small "nanosatellites" or in specialized applications serving as a decentralized or as a distributed power source or uninterruptible power supply.

Raffaelle, R. P.↗

NASA NTRS2013

UltraSail CubeSat Solar Sail Flight Experiment

UltraSail is a next-generation, highrisk, high-payoff sail system for the launch, deployment, stabilization, and control of very large (km2 class) solar sails enabling high payload mass fractions for interplanetary and deep space spacecraft. UltraSail is a non-traditional approach to propulsion technology achieved by combining propulsion and control systems developed for formation- flying microsatellites with an innovative solar sail architecture to achieve controllable sail areas approaching 1 km2, sail subsystem area densities approaching 1 g/m2, and thrust levels many times those of ion thrusters used for comparable deep space missions. UltraSail can achieve outer planetary rendezvous, a deep-space capability now reserved for high-mass nuclear and chemical systems. There is a twofold rationale behind the UltraSail concept for advanced solar sail systems. The first is that sail-andboom systems are inherently size-limited. The boom mass must be kept small, and column buckling limits the boom length to a few hundred meters. By eliminating the boom, UltraSail not only offers larger sail area, but also lower areal density, allowing larger payloads and shorter mission transit times. The second rationale for UltraSail is that sail films present deployment handling difficulties as the film thickness approaches one micrometer. The square sail requires that the film be folded in two directions for launch, and similarly unfolded for deployment. The film is stressed at the intersection of two folds, and this stress varies inversely with the film thickness. This stress can cause the film to yield, forming a permanent crease, or worse, to perforate. By rolling the film as UltraSail does, creases are prevented. Because the film is so thin, the roll thickness is small. Dynamic structural analysis of UltraSail coupled with dynamic control analysis shows that the system can be designed to eliminate longitudinal torsional waves created while controlling the pitch of the blades, while using solar photon pressure to slew the spin axis. Vacuum tests have also verified that electrostatic and molecular adhesion forces can substantially be eliminated by making the film electrically conductive, reducing the peel force of the film off the storage roll to levels of 100s of micro-N. The innovation demonstrated the capability of deploying a six-micron aluminum- coated film from a reel through a slit in vacuum. The innovation also demonstrated a spin-stabilized method for deploying a long reel of solar sail film using solar pressure to spin-up and orbit raise the satellite, and also a gravity gradient method for deploying a long reel of solar sail film using solar pressure to orbit raise the satellite. The solar sail mass fraction of 25% is consistent with high specific impulse ion systems, but without the added weight and cost of a power source and processing unit. The large sail area, coupled with low film density, is giving UltraSail a high payload fraction. The UltraSail deployment scheme unrolls a micrometerscale reflection-coated polyimide film from a storage mandrel to a maximum length of several kilometers with the aid of a blade tip satellite.

Carroll, David↗

NASA NTRS1993

A small satellite design for deep space network testing and training

With the continuing exploration of the Solar System and the reemphasis on Earth focused missions, the need for faster data transmission rates has grown. Ka-band could allow a higher data delivery rate over the current X-band, however the adverse effects of the Earth's atmosphere on Ka are as yet unknown. The Deep Space Network and Jet Propulsion Lab have proposed to launch a small satellite that would simultaneously transmit X and Ka signals to test the viability of switching to Ka-band. The Mockingbird Design Team at the University of Texas at Austin applied small satellite design principles to achieve this objective. The Mockingbird design, named BATSAT, incorporates simple, low-cost systems designed for university production and testing. The BATSAT satellite is a 0.64 m diameter, spherical panel led satellite, mounted with solar cells and omni-directional antennae. The antennae configuration negates the need for active attitude control or spin stabilization. The space-frame truss structure was designed for 11 g launch loads while allowing for easy construction and solar-panel mounting. The communication system transmits at 1 mW by carrying the required Ka and X-band transmitters, as well as an S band transmitter used for DSN training. The power system provides the 8.6 W maximum power requirements via silicon solar arrays and nickel-cadmium batteries. The BATSAT satellite will be lofted into an 1163 km, 70 deg orbit by the Pegasus launch system. This orbit fulfills DSN dish slew rate requirements while keeping the satellite out of the heaviest regions of the Van Allen radiation belts. Each of the three DSN stations capable of receiving Ka-band (Goldstone, Canberra, and Madrid) will have an average of 85 minutes of view-time per day over the satellites ten year design life. Mockingbird Designs hopes that its small satellite design will not only be applicable to this specific mission scenario, but that it could easily be modified for instrument capability for university, government, and/or commercial research.

Mcwilliams, Dennis↗

NASA NTRS1974

Approach guidance for outer planet Pioneer missions

An overall parametric system analysis of the use of an optical approach guidance measurement system (using a V-slit electrooptical 'star-pipper') onboard a Pioneer-type spin-stabilized spacecraft for outer planet missions. Optical measurements of satellites of Saturn and Uranus against the star background are considered. Measurement geometry and spacecraft attitude are evaluated in depth with respect to two Pioneer missions: Saturn-Uranus and Jupiter-Uranus atmospheric entry probe missions with departure from Earth in November/December, 1980. A preliminary evaluation of long-term attitude stability measurements of Pioneer 10 is presented in the Appendix. The major conclusions are: (1) at Saturn flyby, viewing some medium period satellites can provide the best measurement results. (2) At Uranus approach, viewing the planet may give best measurements if early bus-probe separation is required. (3) Final measurement accuracy depends on the data format transmitted to Earth and the applied data interpretation techniques.-

Bejczy, A. K.↗

NASA NTRS1978

Electric propulsion for communications satellites

Electric propulsion systems derive their low overall mass, relative to chemical propulsion systems, from an optimized mix of independently controlled power and mass flow rate. A significant reduction in mass can be secured by substituting electric propulsion for conventional hydrazine systems for all the major propulsion tasks of a communications satellite. Additional advantages for all electric propulsion maneuvers are precision location and higher pointing accuracy. Three auxiliary electric propulsion systems are now available for application on communications satellites: the electrically augmented hydrazine system, the Teflon pulsed plasma system, and the mercury ion thruster system. Primary electric propulsion will be available in the mid to late 1980s as a spin-off of NASA's program of interplanetary exploration.

Free, B. A.↗

NASA NTRS1984

General Maneuver Program

General Maneuver Program, GMAN, computes both orbital and spin-axis reorientation maneuver parameters for various spacecraft. Specifically, GMAN computes detailed maneuver scenarios necessary to achieve desired orbit and attitude maneuvers. Program supports both spinning and inertially-fixed despun spacecraft and includes modeling for either hydrazine or Freon propulsion system.

Erickson, J. A.↗

NASA NTRS1984

Muon spin rotation studies

The bulk of the muon spin rotation research work centered around the development of the muon spin rotation facility at the Alternating Gradient Synchrotron (AGS) of Brookhaven National Laboratory (BNL). The collimation system was both designed and fabricated at Virginia State University. This improved collimation system, plus improvements in detectors and electronics enabled the acquisition of spectra free of background out to 15 microseconds. There were two runs at Brookhaven in 1984, one run was devoted primarily to beam development and the other run allowed several successful experiments to be performed. The effect of uniaxial strain on an Fe(Si) crystal at elevated temperature (360K) was measured and the results are incorporated herein. A complete analysis of Fe pulling data taken earlier is included.

Source record↗

NASA NTRS1988

Radar investigation of asteroids and planetary satellites

The aim is to make radar reconnaissance of near-Earth asteroids, mainbelt ateroids, the Galilean satellites, the Martian satellites, and the largest Saturnian satellites, using the Arecibo 13-cm and the Goldstone 3.5-cm systems. Measurements of echo strength, polarization, and delay/Doppler distribution of echo power provide information about dimensions, spin vector, large-scale topography, cm-to-m-scale morphology, and surface bulk density. The observations also yield refined estimates of target orbital elements. Radar signatures were measured for 31 mainbelt asteroids and 16 near-Earth asteroids since this task began eight years ago. The dispersion in asteroid radar albedoes and circular polarization ratios is extreme, revealing huge differences in surface morphologies, bulk densities, and metal concentration. For the most part, correction between radar signature and VIS/IR class is not high. Many near-Earth asteroids have extremely irregular, nonconvex shapes, but some have polar silhouettes that appear only slightly noncircular. The signatures of 1627 Ivar, 1986 DA, and the approximately 180-km mainbelt asteroid 216 Kleopatra suggest bifurcated shapes. Observational milestones during 1987 and 1988 are noted.

Ostro, Steven J.↗

NASA NTRS1989

Orbit and attitude determination results during launch support operations for SBS-5

Presented are orbit and attitude determination results from the launch of Satellite Business Systems (SBS)-5 satellite on September 8, 1988 by Arianespace. SBS-5 is a (HS-376) spin stabilized spacecraft. The launch vehicle injected the spacecraft into a low inclination transfer orbit. Apogee motor firing (AMF) attitude was achieved with trim maneuvers. An apogee kick motor placed the spacecraft into drift orbit. Postburn, reorientation and spindown maneuvers were performed during the next 25 hours. The spacecraft was on-station 19 days later. The orbit and attitude were determined by both an extended Kalman filter and a weighted least squares batch processor. Although the orbit inclination was low and the launch was near equinox, post-AMF analysis indicated an attitude declination error of 0.034 deg., resulting in a saving of 8.5 pounds of fuel. The AMF velocity error was 0.4 percent below nominal. The post-AMF drift rate was determined with the filter only 2.5 hours after motor firing. The filter was used to monitor and retarget the reorientation to orbit normal in real time.

Hartman, K. R.↗

NASA NTRS1988

Attitude Determination Error Analysis System (ADEAS) mathematical specifications document

The mathematical specifications of Release 4.0 of the Attitude Determination Error Analysis System (ADEAS), which provides a general-purpose linear error analysis capability for various spacecraft attitude geometries and determination processes, are presented. The analytical basis of the system is presented. The analytical basis of the system is presented, and detailed equations are provided for both three-axis-stabilized and spin-stabilized attitude sensor models.

Nicholson, Mark↗

NASA NTRS2012

A Common Proper Motion Stellar Companion to HAT-P-7

We report that HAT-P-7 has a common proper motion stellar companion. The companion is located at approx. 3.9 arcsec to the east and estimated as an M5.5V dwarf based on its colors. We also confirm the presence of the third companion, which was first reported by Winn et al. (2009), based on long-term radial velocity measurements. We revisit the migration mechanism of HAT-P-7b given the presence of those companions, and propose sequential Kozai migration as a likely scenario in this system. This scenario may explain the reason for an outlier in the discussion of the spin-orbit alignment timescale for HAT-P-7b by Albrecht et al. (2012).

Grady, Carol A.↗

NASA NTRSDate not supplied

Hold Down and Release Mechanism Hardware Simulator for Ground Deployment Testing

Hold Down and Release Mechanisms (HDRMs) are a common piece of spacecraft hardware which restrain components during launch, transit, and later allow them to be extended, jettisoned, separated, etc. at the spacecraft’s destination. Due to prohibitively high cost, long lead times, and limited reset capability, testing of mechanisms that require HDRMs is often difficult. This project explored the possibility of designing HDRM-analogue devices that would replicate the function of flight HDRMs during ground-based testing for a fraction of the cost of a flight HDRM. These devices were designed with the requirements of the Spin Eject Mechanism to be used on the Mars Sample Return (MSR) Capture, Containment, and Return System (CCRS) in mind. A trade study was conducted to determine the optimal internal mechanism design of the device, comparing several mechanisms from existing HDRMs to several simpler, more novel mechanisms. CAD software was used to create a preliminary design of the device utilizing a mix of commercial off-the-shelf parts and custom-manufactured parts. Several straightforward analyses were conducted to determine the most effective models for the off-the-shelf parts. A prototype largely composed of 3D-printed components was also created to proof the design concept. Although this HDRM simulator was designed for a particular mechanism for the CCRS mission, its general design is useful for ground testing of a variety of mechanisms requiring similar HDRM simulators.

Kian Vilhauer↗