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NASA NTRS1991

Sunsynchronous low Earth orbit spacecraft concepts and technology requirements for global change monitoring

The Global Change Technology Initiative listing of instruments for operation in low Earth, sunsynchronous orbits contain 21 entries, of which 20 are carried aboard multi-instrument spacecraft. This list identifies the temporal requirements for repetition of measurements and also includes groups of instruments that make complementing measurements. Definitions for individual spacecraft follows the temporal and grouping requirements to establish constellations which will provide the measurement data. The definitions of constellations for multi-instrument spacecraft show two alternatives: a constellation of 10 spacecraft, each compatible with launch by a Delta booster; a constellation of 4 spacecraft, each requiring a Titan booster. Operating subsystems for the individual spacecraft can use modular concepts that are adaptations based upon current plans for improving the performance of the NASA-Goddard Multimission Modular units. The descriptions of the spacecraft and constellations begins with a compilation of instrument related requirements that define the principal system performance parameters and operating capabilities.

Garrett, L. Bernard↗
NASA NTRS1991

Hoop column soil moisture spacecraft in low Earth orbit for global change monitoring

A subset of the total Global Change Technology Initiative instruments are required to be in low Earth, sunsynchronous orbits. There is one instrument, however, that requires its own specialized spacecraft; the Soil Moisture Microwave Radiometer (SMMR). The characteristic structure of the instrument is the 118 m hoop column support structure. The hoop is supported by an axially placed column. Tension cables support and shape an electromagnetically reflective mesh surface. The instrument is capable of detecting frequencies in the 1.4 GHz range (Soil Moisture and Sea Salinity). Three apertures are used to reduce the degree of paraboloid offset and improve the beam quality. The spacecraft configuration is determined by the instrument support requirements and the requirement that it can fit into the Titan IV cargo bay. The configuration is derived by cross referencing the instrument performance requirements with the performance of the spacecraft. The spacecraft design is similar with the Multi-mission Modular Spacecraft in terms of size and packaging. A description of the spacecraft's features will yield a summary of the technologies needed for the SMMR spacecraft.

Ferebee, Melvin J., Jr.↗
NASA NTRS1992

Attitude stabilization of a rigid spacecraft using momentum wheel actuators operating in a failure mode

In this paper, we demonstrate that, in fact, two momentum wheel actuators can be used to control the attitude of a rigid spacecraft and that arbitrary reorientation maneuvers of the spacecraft can be accomplished in a specific sense. The complete spacecraft dynamics cannot be stabilized to an equilibrium attitude. However, the spacecraft equations are small time locally controllable in a reduced nonlinear sense. The reduced spacecraft dynamics cannot be asymptotically stabilized to an equilibrium attitude using a time-invariant continuous feedback control law, but a discontinuous feedback control strategy is constructed which asymptotically stabilizes an equilibrium attitude of the spacecraft. Consequently, arbitrary reorientation of the spacecraft can be accomplished using this discontinuous feedback control strategy.

Krishnan, Hariharan↗
NASA NTRS1990

Integrated controls-structures design methodology development for a class of flexible spacecraft

Future utilization of space will require large space structures in low-Earth and geostationary orbits. Example missions include: Earth observation systems, personal communication systems, space science missions, space processing facilities, etc., requiring large antennas, platforms, and solar arrays. The dimensions of such structures will range from a few meters to possibly hundreds of meters. For reducing the cost of construction, launching, and operating (e.g., energy required for reboosting and control), it will be necessary to make the structure as light as possible. However, reducing structural mass tends to increase the flexibility which would make it more difficult to control with the specified precision in attitude and shape. Therefore, there is a need to develop a methodology for designing space structures which are optimal with respect to both structural design and control design. In the current spacecraft design practice, it is customary to first perform the structural design and then the controller design. However, the structural design and the control design problems are substantially coupled and must be considered concurrently in order to obtain a truly optimal spacecraft design. For example, let C denote the set of the 'control' design variables (e.g., controller gains), and L the set of the 'structural' design variables (e.g., member sizes). If a structural member thickness is changed, the dynamics would change which would then change the control law and the actuator mass. That would, in turn, change the structural model. Thus, the sets C and L depend on each other. Future space structures can be roughly divided into four mission classes. Class 1 missions include flexible spacecraft with no articulated appendages which require fine attitude pointing and vibration suppression (e.g., large space antennas). Class 2 missions consist of flexible spacecraft with articulated multiple payloads, where the requirement is to fine-point the spacecraft and each individual payload while suppressing the elastic motion. Class 3 missions include rapid slewing of spacecraft without appendages, while Class 4 missions include general nonlinear motion of a flexible spacecraft with articulated appendages and robot arms. Class 1 and 2 missions represent linear mathematical modeling and control system design problems (except for actuator and sensor nonlinearities), while Class 3 and 4 missions represent nonlinear problems. The development of an integrated controls/structures design approach for Class 1 missions is addressed. The performance for these missions is usually specified in terms of (1) root mean square (RMS) pointing errors at different locations on the structure, and (2) the rate of decay of the transient response. Both of these performance measures include the contributions of rigid as well as elastic motion.

Maghami, P. G.↗
NASA NTRS1993

Low power arcjet system spacecraft impacts

Application of electrothermal arcjets on communications satellites requires assessment of integration concerns identified by the user community. Perceived risks include plume contamination of spacecraft materials, induced arcing or electrostatic discharges between differentially charged spacecraft surfaces, and conducted and radiated electromagnetic interference (EMI) for both steady state and transient conditions. A Space Act agreement between Martin Marietta Astro Space, the Rocket Research Company, and NASA's Lewis Research Center was established to experimentally examine these issues. Spacecraft materials were exposed to an arcjet plume for 40 hours, representing 40 weeks of actual spacecraft life, and contamination was characterized by changes in surface properties. With the exception of the change in emittance of one sample, all measurable changes in surface properties resulted in acceptable end of life characteristics. Charged spacecraft samples were benignly and consistently reduced to ground potential during exposure to the powered arcjet plume, suggesting that the arcjet could act as a charge control device on spacecraft. Steady state EMI signatures obtained using two different power processing units were similar to emissions measured in a previous test. Emissions measured in UHF, S, C, Ku and Ka bands obtained a null result which verified previous work in the UHF, S, and C bands. Characteristics of conducted and radiated transient emissions appear within standard spacecraft susceptibility criteria.

Pencil, Eric J.↗
NASA NTRS1993

A packet-based concept for spacecraft command planning

The current generation of spacecraft being developed and operated by the Applied Physics Laboratory provides users with access to a broad spectrum of scientific instruments on maneuverable platforms that can be oriented for observation of both moving and stationary targets of interest. The capability of these increasingly complex spacecraft to perform data collection operations is approaching one observation per orbit. To enable both rapid configuration and generation of complex spacecraft command sequences, as well as reusability of command sequences among data collection operations, a packet-based concept for spacecraft command planning has been developed. The configuration of the spacecraft for any operation is designed using 'packets' where a packet represents a set of commands that is reusable. The packets can be combined in varying levels of functionality, and in varying time relationships, to create an observation timeline. At the lowest packet level are primitives. Primitives relate the details of command generation for a particular spacecraft to a 'message template.' Thus the packet concept itself is reusable from one spacecraft to the next.

Barnes, Valerie B.↗
NASA NTRS1994

Vulnerability of manned spacecraft to crew loss from orbital debris penetration

Orbital debris growth threatens the survival of spacecraft systems from impact-induced failures. Whereas the probability of debris impact and spacecraft penetration may currently be calculated, another parameter of great interest to safety engineers is the probability that debris penetration will cause actual spacecraft or crew loss. Quantifying the likelihood of crew loss following a penetration allows spacecraft designers to identify those design features and crew operational protocols that offer the highest improvement in crew safety for available resources. Within this study, a manned spacecraft crew survivability (MSCSurv) computer model is developed that quantifies the conditional probability of losing one or more crew members, P(sub loss/pen), following the remote likelihood of an orbital debris penetration into an eight module space station. Contributions to P(sub loss/pen) are quantified from three significant penetration-induced hazards: pressure wall rupture (explosive decompression), fragment-induced injury, and 'slow' depressurization. Sensitivity analyses are performed using alternate assumptions for hazard-generating functions, crew vulnerability thresholds, and selected spacecraft design and crew operations parameters. These results are then used to recommend modifications to the spacecraft design and expected crew operations that quantitatively increase crew safety from orbital debris impacts.

Williamsen, J. E.↗
NASA NTRS1994

A geometric model of a V-slit Sun sensor correcting for spacecraft wobble

A V-Slit sun sensor is body-mounted on a spin-stabilized spacecraft. During injection from a parking or transfer orbit to some final orbit, the spacecraft may not be dynamically balanced. This may result in wobble about the spacecraft spin axis as the spin axis may not be aligned with the spacecraft's axis of symmetry. While the widely used models in Spacecraft Attitude Determination and Control, edited by Wertz, correct for separation, elevation, and azimuthal mounting biases, spacecraft wobble is not taken into consideration. A geometric approach is used to develop a method for measurement of the sun angle which corrects for the magnitude and phase of spacecraft wobble. The algorithm was implemented using a set of standard mathematical routines for spherical geometry on a unit sphere.

Mcmartin, W. P.↗
NASA NTRS1994

Safety aspects of spacecraft commanding

The commanding of spacecraft is a potentially hazardous activity for the safety of the spacecraft. Present day control systems contain safety features in their commanding subsystem and in addition, strict procedures are also followed by operations staff. However, problems have occurred on a number of missions as a result of erroneous commanding leading in some cases to spacecraft contingencies and even to near loss of the spacecraft. The problems of checking commands in advance are increased by the tendency in modern spacecraft to use blocked/time-tagged commands and the increased usage of on-board computers, for which commands changing on-board software tables can radically change spacecraft or subsystem behavior. This paper reports on an on-going study. The study aims to improve the approach to safety of spacecraft commanding. It will show how ensuring 'safe' commanding can be carried out more efficiently, and with greater reliability, with the help of knowledge based systems and/or fast simulators. The whole concept will be developed based on the Object-Oriented approach.

Peccia, N.↗
NASA NTRS1995

An LDEF follow-on spacecraft concept

The successful flight, retrieval, and analyses of the Long Duration Exposure Facility (LDEF) experiments demonstrated the value of long duration space exposure for a broad spectrum of science and engineering investigations. The original LDEF was an excellent gravity gradient spacecraft, but because of its 9 m length and 9,700 kg mass it was difficult to manifest on the Shuttle, for either launch or retrieval, in conjunction with other payloads. This paper discusses an LDEF follow-on spacecraft concept whose short stowed length (approximately 3 m) greatly improves Shuttle manifesting opportunities while still providing very large surface area exposure for experiments. Deployable 'wings' on each end of the short, 'cylindrical' main body of this new spacecraft provide the means for gravity gradient stabilization while greatly increasing the spacecraft surface area. The center section of the spacecraft is oriented with the end faces of the twelve sided, 4.2 m diameter 'cylinder' perpendicular to the velocity vector thus providing large areas for experiments in both the ram and anti-ram directions as well as additional exposure area around the periphery of the cylinder. When deployed and properly oriented with the Shuttle's Remote Manipulator System (RMS), both wings of the spacecraft are oriented edge on to the direction of motion and lie in the plane which contains the local gravity vector. The relatively thin wings readily accommodate dual side exposure of glass plate stacks for cosmic ray detection. Flat surfaces mounted normal to and on the periphery of the wings provide additional areas in both the ram and anti-ram directions for cosmic dust, micrometeoroid, and orbital debris collection free of contamination from 'splatter' off secondary surfaces. The baseline concept provides enhancements not available on the original LDEF such as solar array generated electrical power and data telemetry. Status of the efforts to promote support for and ultimately space flight of this concept will be presented. Suggestions for improvements in the spacecraft design and proposed utilization are solicited.

Keller, Vernon↗
NASA NTRS1995

Development of a Robust star identification technique for use in attitude determination of the ACE spacecraft

The Advanced Composition Explorer (ACE) spacecraft is designed to fly in a spin-stabilized attitude. The spacecraft will carry two attitude sensors - a digital fine Sun sensor and a charge coupled device (CCD) star tracker - to allow ground-based determination of the spacecraft attitude and spin rate. Part of the processing that must be performed on the CCD star tracker data is the star identification. Star data received from the spacecraft must be matched with star information in the SKYMAP catalog to determine exactly which stars the sensor is tracking. This information, along with the Sun vector measured by the Sun sensor, is used to determine the spacecraft attitude. Several existing star identification (star ID) systems were examined to determine whether they could be modified for use on the ACE mission. Star ID systems which exist for three-axis stabilized spacecraft tend to be complex in nature and many require fairly good knowledge of the spacecraft attitude, making their use for ACE excessive. Star ID systems used for spinners carrying traditional slit star sensors would have to be modified to model the CCD star tracker. The ACE star ID algorithm must also be robust, in that it will be able to correctly identify stars even though the attitude is not known to a high degree of accuracy, and must be very efficient to allow real-time star identification. The paper presents the star ID algorithm that was developed for ACE. Results from prototype testing are also presented to demonstrate the efficiency, accuracy, and robustness of the algorithm.

Woodard, Mark↗
NASA NTRS1994

Solar activity and coronal mass ejections on the western hemisphere of the Sun in mid-August 1989: Association with interplanetary observations at the ICE and IMP 8 spacecraft

During the 10-day period from August 12 to 21, 1989, a sequence of coronal mass ejections (CMEs) was observed above the west limb of the Sun by the Solar Maximum Mission (SMM) coronagraph. Most of these CMEs apparently originated in the vicinity of one particularly active region during its passage from near central meridian to behind the west limb of the Sun. We present observations made at 1 AU during this period by the International Cometary Explorer (ICE) (formerly International Sun Earth Explorer-3 (ISEE 3)) and Interplanetary Monitoring Platform (IMP 8) spacecraft which were separated by approximately 75 deg in heliolongitude. Following CMEs on August 12 associated with solar events at approximately W40 deg, IMP 8 (in Earth orbit) detected a strong shock followed by signatures in magnetic field, solar wind plasma, and energetic ion data which suggest that CME-related material ('ejecta') forming the shock driver engulfed the spacecraft. This spacecraft only observed weak shocks, and no ejecta, from later CMEs originating further west of the spacecraft. In contrast ICE, off the west limb at approximately W75 deg, observed the shock from the W40 deg event but failed to encounter the shock driver, whereas clear ejecta signatures were observed following events further west, closer to the spacecraft heliolongitude. The disappearance of these signatures (which include bidirectional energetic ion flows, bidirectional solar wind heat fluxes, quiet, enhanced magnetic fields and anomalously cool plasma) at IMP 8 and their emergence at ICE as the solar source region moved westward supports the association of such signatures with ejecta related to CMEs. The dual-spacecraft observations are also consistent with the conclusion of Richardson and Cane (1993) that ejecta at 1 AU typically extend approximately 50 degs in longitude from the solar source. Some plausible associations between particular intervals of ejecta signatures at ICE and individual CMEs are made. However, these associations are complicated by the large number of CMEs present, by intermittent ICE data coverage, and by uncertainties in the ejecta propagation speeds to the spacecraft.

Richardson, I. G.↗
NASA NTRS1994

Radio occultation studies of the Venus atmosphere with the Magellan spacecraft. 1: Experimental description and performance

While primarily designed for radar studies of the Venus surface, the high effective isotropic radiated power (EIRP) from the Magellan spacecraft makes it an ideal transmitter for use in radio occultation measurements of the refractivity and absorptivity of the Venus atmosphere. Such experiments have been conducted involving transmissions at 2.3 GHz and 8.4 GHz (13 cm and 3.6 cm, respectively), during spacecraft ingress. Since the stability of the spacecraft transmitter is critical for accurately determining the Doppler shift and amplitude attenuation created as the ray penetrates the atmosphere, the spacecraft transmitter was locked to a 2.1 GHz uplink from a 70-meter DSN station which also received the signals. Because of the high directivity of the spacecraft antenna, and the significant ray bending in the deep Venus atmosphere, a spacecraft tracking maneuver was designed to keep the spacecraft antenna pointed in the direction of the refracted ray path back to Earth. This tracking maneuver, plus the high EIRP of the Magellan transmitter has yielded 3.6 cm refractivity and absorptivity profiles down to the 35 km altitude and 13 cm profiles down to the altitude of critical refraction (approximately 33 km). The statistical uncertainties in the derived profiles are significantly lower than those previously obtained, resulting in extremely accurate profiles of H2SO4 (g) abundance as discussed in an accompanying paper.

Steffes, Paul G.↗
NASA NTRS1995

Trace chemical contaminant generation rates for spacecraft contamination control system design

A spacecraft presents a unique design challenge with respect to providing a comfortable environment in which people can live and work. All aspects of the spacecraft environmental design including the size of the habitable volume, its temperature, relative humidity, and composition must be considered to ensure the comfort and health of the occupants. The crew members and the materials selected for outfitting the spacecraft play an integral part in designing a habitable spacecraft because material offgassing and human metabolism are the primary sources for continuous trace chemical contaminant generation onboard a spacecraft. Since these contamination sources cannot be completely eliminated, active control processes must be designed and deployed onboard the spacecraft to ensure an acceptably clean cabin atmosphere. Knowledge of the expected rates at which contaminants are generated is very important to the design of these processes. Data from past spacecraft missions and human contaminant production studies have been analyzed to provide this knowledge. The resulting compilation of contaminants and generation rates serve as a firm basis for past, present, and future contamination control system designs for space and aeronautics applications.

Perry, J. L.↗
NASA NTRS1996

Spacecraft Maximum Allowable Concentrations for Selected Airborne Contaminants

As part of its efforts to promote safe conditions aboard spacecraft, NASA requested the National Research Council (NRC) to develop guidelines for establishing spacecraft maximum allowable concentrations (SMACs) for contaminants, and to review SMACs for various spacecraft contaminants to determine whether NASA's recommended exposure limits are consistent with the guidelines recommended by the subcommittee. In response to this request, the NRC first developed criteria and methods for preparing SMACs for spacecraft contaminants, published in its 1992 report Guidelines for Developing Spacecraft Maximum Allowable Concentrations for Space Station Contaminants . Since then, the Subcommittee on Spacecraft Maximum Allowable Concentrations has been reviewing NASA's documentation of chemical-specific SMACs as described in the Introduction to this volume. This report is the third volume in the series Spacecraft Maximum Allowable Concentrations for Space Station Contaminants . The first volume was published in 1994 and the second in 1996.

National Research Council↗
NASA NTRS1998

MIDEX Advanced Modular and Distributed Spacecraft Avionics Architecture

MIDEX (Medium Class Explorer) is the newest line in NASA's Explorer spacecraft development program. As part of the MIDEX charter, the MIDEX spacecraft development team has developed a new modular, distributed, and scaleable spacecraft architecture that pioneers new spaceflight technologies and implementation approaches, all designed to reduce overall spacecraft cost while increasing overall functional capability. This resultant "plug and play" system dramatically decreases the complexity and duration of spacecraft integration and test, providing a basic framework that supports spacecraft modularity and scalability for missions of varying size and complexity. Together, these subsystems form a modular, flexible avionics suite that can be modified and expanded to support low-end and very high-end mission requirements with a minimum of redesign, as well as allowing a smooth, continuous infusion of new technologies as they are developed without redesigning the system. This overall approach has the net benefit of allowing a greater portion of the overall mission budget to be allocated to mission science instead of a spacecraft bus. The MIDEX scaleable architecture is currently being manufactured and tested for use on the Microwave Anisotropy Probe (MAP), an inhouse program at GSFC.

Ruffa, John A.↗
NASA NTRS1998

Spacecraft Potential Control by the Plasma Source Instrument on the POLAR Satellite

This increasingly recognized that the low-energy core plasma is a critically important part of magnetospheric plasma transport, yet this plasma cannot be accurately measured from spacecraft at potentials much different from that of the ambient plasma. In low-density regions such as the polar cap and lobes, spacecraft charge positively, excluding core ions from the spacecraft and accelerating core electrons so much that their velocities cannot be measured with any accuracy. In regions of high electron pressure and temperature, spacecraft charge negatively, excluding the ambient core electrons and accelerating the core ions so much that their velocity cannot be accurately measured. Plasma contactors have been used on a number of spacecraft operating in low-plasma-density regions to prevent charging of spacecraft to high potentials, particularly when exposed to high fluxes of energetic particles. This concern has prompted extensive studies by NASA for use of plasma contractors on the international space station where solar arrays may significantly affect the spacecraft potential.

Comfort, R. H.↗
NASA NTRS1995

Analysis of Static Spacecraft Floating Potential at Low Earth Orbit (LEO)

Spacecraft floating potential is the charge on the external surfaces of orbiting spacecraft relative to the space. Charging is caused by unequal negative and positive currents to spacecraft surfaces. The charging process continues until the accelerated particles can be collected rapidly enough to balance the currents at which point the spacecraft has reached its equilibrium or floating potential. In low inclination. Low Earth Orbit (LEO), the collection of positive ion and negative electrons. in a particular direction. are typically not equal. The level of charging required for equilibrium to be established is influenced by the characteristics of the ambient plasma environment. by the spacecraft motion, and by the geometry of the spacecraft. Using the kinetic theory, a statistical approach for studying the interaction is developed. The approach used to study the spacecraft floating potential depends on which phenomena are being applied. and on the properties of the plasma. especially the density and temperature. The results from kinetic theory derivation are applied to determine the charging level and the electric potential distribution at an infinite flat plate perpendicular to a streaming plasma using finite-difference scheme.

Herr, Joel L.↗