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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 235 records · Page 13
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NASA NTRS2011

Compact, Controlled Resistance Exercise Device

Spaceflight leads to muscle and bone atrophy. Isoinertial (free-weight) exercises provide a sufficient stimulus to elicit increases in both muscle strength and bone mineral density in Earth-based studies. While exercise equipment is in use on the International Space Station for crewmember health maintenance, current devices are too large to place in a transport vehicle or small spacecraft. Therefore, a portable computer controlled resistance exercise device is being developed that is able to simulate the inertial loading experienced when lifting a mass on Earth. This portable device weighs less than 50 lb and can simulate the resistance of lifting and lowering up to 600 lb of free-weights. The objective is to allow crewmembers to perform resistance exercise with loads capable of maintaining muscle and bone health. The device is reconfigurable and allows for the performance of typical Earth-based free-weight exercises. Forces exerted, volume of work, range of motion, time-under-tension, and speed/ acceleration of movement are recorded and can be remotely monitored to track progress and modify individual protocols based on exercise session data. A performance evaluation will be completed and data will be presented that include ground-reaction force comparisons between the device and free-weight dead-lifts over a spectrum of resistance levels. Movement biomechanics will also be presented.

Paulus, David C.↗
NASA NTRS2004

Modular, Reconfigurable, and Rapid Response Space Systems: The Remote Sensing Advanced Technology Microsatellite

Modular, Reconfigurable, and Rapid-response (MR(sup 2)) space systems represent a paradigm shift in the way space assets of all sizes are designed, manufactured, integrated, tested, and flown. This paper will describe the MR(sup 2) paradigm in detail, and will include guidelines for its implementation. The Remote Sensing Advanced Technology microsatellite (RSAT) is a proposed flight system test-bed used for developing and implementing principles and best practices for MR(sup 2) spacecraft, and their supporting infrastructure. The initial goal of this test-bed application is to produce a lightweight (approx. 100 kg), production-minded, cost-effective, and scalable remote sensing micro-satellite capable of high performance and broad applicability. Such applications range from future distributed space systems, to sensor-webs, and rapid-response satellite systems. Architectures will be explored that strike a balance between modularity and integration while preserving the MR(sup 2) paradigm. Modularity versus integration has always been a point of contention when approaching a design: whereas one-of-a-kind missions may require close integration resulting in performance optimization, multiple and flexible application spacecraft benefit &om modularity, resulting in maximum flexibility. The process of building spacecraft rapidly (< 7 days), requires a concerted and methodical look at system integration and test processes and pitfalls. Although the concept of modularity is not new and was first developed in the 1970s by NASA's Goddard Space Flight Center (Multi-Mission Modular Spacecraft), it was never modernized and was eventually abandoned. Such concepts as the Rapid Spacecraft Development Office (RSDO) became the preferred method for acquiring satellites. Notwithstanding, over the past 30 years technology has advanced considerably, and the time is ripe to reconsider modularity in its own right, as enabler of R(sup 2), and as a key element of transformational systems. The MR2 architecture provides a competitive advantage over the old modular approach in its rapid response to market needs that are difficult to predict both from the perspectives of evolving technology, as well as mission and application requirements.

Esper, Jaime↗
NASA NTRS1988

Final (Tier 1) environmental impact statement for the Galileo and Ulysses Missions

Presented here is a Final (Tier 1) Environmental Impact Statement (EIS) addressing the potential environmental consequences associated with continuing the modifications of the Galileo and Ulysses spacecraft for launch using a booster/upper stage combination that is different from the one planned for use prior to the Challenger accident, while conducting the detailed safety and environmental analysis in order to preserve the October 1989 launch opportunity for Galileo and an October 1990 launch opportunity for Ulysses. While detailed safety and environmental analyses associated with the missions are underway, they currently are not complete. Nevertheless, sufficient information is available to enable a choice among the reconfiguration alternatives presented. Relevant assessments of the potential for environmental impacts are presented.

Source record↗
NASA NTRS2019

Multi-Band and Multi-Function Reconfigurable Gallium Nitride Based Fully Solid-State Microwave Power Module for Cognitive Radio/Radar Platforms

This paper presents as a proof-of-concept (POC) the design, integration, and performance of a novel reconfigurable S-/X-band Gallium Nitride (GaN) based fully solid-state microwave power module (SSMPM) for the role as the transmit module in a cognitive radio (CR). The SSMPM synergistically integrates multiple amplifiers through diplexing and high power switches to enable a single SSMPM capable of functioning as both S-/X-band amplifiers for telemetry, tracking, and command (TT&C), telecommunications, and science data downlink or as X-band radar for proximity sensing onboard a planetary exploration spacecraft. Integration of an electric field shaping field plate (FP) onto the GaN high electron mobility transistors (HEMTs) in this SSMPM provides increased performance and reliability for operation in the harsh conditions of space. This SSMPM is capable of delivering saturated power (Psat) of 39 dBm (8 W continuous wave (CW)) at S-band, P(sub sat) of 43 dBm (20 W CW) at X-band, and P(sub sat) of >50 dBm (>100 W Pulsed) at X-band.

Simons, Rainee N.↗
NASA NTRS2010

Computation of Earth Science Products on Spaceborne Platforms

Spaceborne sensors like NASA's Hyperion hyperspectral imager generate huge data volumes, and several near-term trends indicate that data volumes will only increase. Next-generation hyperspectral missions, such as NASA's Hyperspectral Infrared Imager (HyspIRI), will operate at higher duty cycles and higher data rates, and their users will expect products to be generated from the data in near real time [1]. Barring a sudden advance in satellite downlink capacity, these trends point to a need to process data and generate products onboard the spacecraft. Rather than downlink an entire hyperspectral image cube, onboard processing enables satellites to downlink partial or completed scientific data products, which are often one to two orders of magnitude smaller than the original image. In addition, a satellite with onboard data processing resources and direct broadcast transmission equipment could send data products directly to first responders, research scientists or other users on the ground. Next-generation space-capable data processors will have a combination of reconfigurable gate arrays, digital signal processors and general-purpose CPUs. Correctly programmed and configured, these resources are sufficient to run sophisticated data analysis programs, including hyperspectral image processing algorithms that commonly run on desktop computers [2]. This paper describes how we implemented one such program, the HSEG hierarchical image segmentation algorithm, software commonly used on desktop and parallel processors, on a hardware platform designed to mimic a next-generation space-capable data processor [3]. We also describe our approach to porting the algorithm to and optimizing it for the new platform, and determine the expected performance gains enabled by our design. This extended abstract will describe the HSEG algorithm and hardware platform in greater detail, provide an analysis of the key function within the algorithm that required hardware acceleration, and describe our implementation of that function in hardware.

Fisher, Kevin↗
NASA NTRS2023

Partially Erupted Prominence Material as a Diagnostic of Coronal Mass Ejection Trajectory

Coronal mass ejections (CMEs) are energetic releases of large-scale magnetic structures from the Sun. CMEs can have impacts on spacecraft and at Earth. This trajectory is typically assumed to be radial, but often the CME moves outward with some spatial off set from the source region where the eruption initially occurred. A CME is frequently accompanied by a prominence eruption, a movement of cool, dense material up into the corona that can be ejected or fall back down. We investigate eruptions in which some portion of the prominence material falls back to the Sun along field lines which have reconfigured in the eruption, rather than draining back to the source or escaping with the CME. Using a method called persistence mapping, 304 ˚A images from the Solar Dynamics Observatory (SDO), and coronagraph images from the Solar and Heliospheric Observatory (SOHO), we measure and compare the offsets in latitude of 20 CMEs and their respective prominences with respect to the source region. The 20 events were chosen to sample over the first 10 years of the SDO mission. We find that the offsets are correlated. We find no difference between eruptions offset towards the equator or the poles, suggesting that the offset is a result of local changes in the eruptive field, rather than of the Sun’s global magnetic field structure. These findings help us contextualize individual eruptions and highlight changes in the local magnetic field associated with the prominence eruption.

B A Hovis-Afflerbach↗
NASA NTRSDate not supplied

Dynamic Characterization of NASA Armstrong Test Facility Mechanical Vibration Facility

In preparation for Sierra Nevada Corporation’s (SNC) Dream Chaser spacecraft vibration test campaign at NASA Armstrong Test Facility (ATF) in Sandusky, Ohio, a dynamic characterization of the Mechanical Vibration Facility (MVF) is currently ongoing. The Mechanical Vibration Facility is comprised of an 18 ft diameter annulus table that is driven with sixteen hydraulic vertical actuator assemblies and four hydraulic horizontal actuator assemblies. During a test campaign this allows for single axis vibration testing in the vertical axis and in each of the two orthogonal horizontal axes without the need for reconfiguring the test article. The dynamic characterization of the facility was defined to follow a building-block approach requiring three vibration test configurations: (1) MVF bare table (2) MVF table and head expander (3) MVF table, head expander, and a dummy test article. Following the completion of the facility dynamic characterization, a facility finite element model (FEM) will be test-verified allowing for more accurate pretest analysis used for determining response limits and abort levels. To date two of the three test configurations have been completed. Test results have been used to verify the facility FEM that includes both the 18 ft diameter annulus table and the aluminum head expander that fills in the center opening of the annulus to provide a continuous flat mounting surface. Both the MVF table and the head expander models were previously correlated using test results from free-free boundary condition modal tests so model updates to the facility FEM focused on the uncorrelated vertical actuator assemblies and horizontal actuator assemblies. The accuracy of the test-verified facility FEM will be further determined based on comparison of analytical and test results from the vibration test of the dummy test article. Efforts will also be made to determine fixed base modes of the test article attached to the flexible vibration table using fixed base correction techniques. This paper will discuss test and analytical results from the MVF dynamic characterization test sequence and corresponding model updating effort.

Emma L Pierson↗
NASA NTRS1998

The Software Design for the Wide-Field Infrared Explorer Attitude Control System

The Wide-Field Infrared Explorer (WIRE), currently scheduled for launch in September 1998, is the fifth of five spacecraft in the NASA/Goddard Small Explorer (SMEX) series. This paper presents the design of WIRE's Attitude Control System flight software (ACS FSW). WIRE is a momentum-biased, three-axis stabilized stellar pointer which provides high-accuracy pointing and autonomous acquisition for eight to ten stellar targets per orbit. WIRE's short mission life and limited cryogen supply motivate requirements for Sun and Earth avoidance constraints which are designed to prevent catastrophic instrument damage and to minimize the heat load on the cryostat. The FSW implements autonomous fault detection and handling (FDH) to enforce these instrument constraints and to perform several other checks which insure the safety of the spacecraft. The ACS FSW implements modules for sensor data processing, attitude determination, attitude control, guide star acquisition, actuator command generation, command/telemetry processing, and FDH. These software components are integrated with a hierarchical control mode managing module that dictates which software components are currently active. The lowest mode in the hierarchy is the 'safest' one, in the sense that it utilizes a minimal complement of sensors and actuators to keep the spacecraft in a stable configuration (power and pointing constraints are maintained). As higher modes in the hierarchy are achieved, the various software functions are activated by the mode manager, and an increasing level of attitude control accuracy is provided. If FDH detects a constraint violation or other anomaly, it triggers a safing transition to a lower control mode. The WIRE ACS FSW satisfies all target acquisition and pointing accuracy requirements, enforces all pointing constraints, provides the ground with a simple means for reconfiguring the system via table load, and meets all the demands of its real-time embedded environment (16 MHz Intel 80386 processor with 80387 coprocessor running under the VRTX operating system). The mode manager organizes and controls all the software modules used to accomplish these goals, and in particular, the FDH module is tightly coupled with the mode manager.

Anderson, Mark O.↗
NASA NTRS2018

Development of Spacecraft Radiated Susceptibility RS103 Requirements From Modeling Methods

In defining radiated susceptibility requirements for a spacecraft with multiple number of receivers and transmitters in close proximity to each other, the main objective of an RF coupling analysis is to determine if the mechanical configuration of the receiver and transmitter antennas presents a risk to the functionality and safety of sensitive science instruments on the payload as a result of unintended RF coupling. Where there is a potential risk for interference or permanent damage, further analysis is required to evaluate the feasibility of mitigation schemes, such as mechanical reconfiguration of antennas or additional RF filtering. From these coupling analysis results, radiated susceptibility RS103 requirements are derived to better reflect actual requirement levels with adequate test margins.

Gonzalez, Edward C.↗
NASA NTRS2008

Semi-Automated Diagnosis, Repair, and Rework of Spacecraft Electronics

NASA's Constellation Program for Exploration of the Moon and Mars places human crews in extreme isolation in resource scarce environments. Near Earth, the discontinuation of Space Shuttle flights after 2010 will alter the up- and down-mass capacity for the International Space Station (ISS). NASA is considering new options for logistics support strategies for future missions. Aerospace systems are often composed of replaceable modular blocks that minimize the need for complex service operations in the field. Such a strategy however, implies a robust and responsive logistics infrastructure with relatively low transportation costs. The modular Orbital Replacement Units (ORU) used for ISS requires relatively large blocks of replacement hardware even though the actual failed component may really be three orders of magnitude smaller. The ability to perform in-situ repair of electronics circuits at the component level can dramatically reduce the scale of spares and related logistics cost. This ability also reduces mission risk, increases crew independence and improves the overall supportability of the program. The Component-Level Electronics Assembly Repair (CLEAR) task under the NASA Supportability program was established to demonstrate the practicality of repair by first investigating widely used soldering materials and processes (M&P) performed by modest manual means. The work will result in program guidelines for performing manual repairs along with design guidance for circuit reparability. The next phase of CLEAR recognizes that manual repair has its limitations and some highly integrated devices are extremely difficult to handle and demand semi-automated equipment. Further, electronics repairs require a broad range of diagnostic capability to isolate the faulty components. Finally repairs must pass functional tests to determine that the repairs are successful and the circuit can be returned to service. To prevent equipment demands from exceeding spacecraft volume capacity and skill demands from exceeding crew time and training limits, the CLEAR project is examining options provided by non-real time tele-operations, robotics, and a new generation of diagnostic equipment. This paper outlines a strategy to create an effective repair environment where, with the support of ground based engineers, crewmembers can diagnose, repair and test flight electronics in-situ. This paper also discusses the implications of successful tele-robotic repairs when expanded to rework and reconfiguration of used flight assets for building Constellation infrastructure elements.

Struk, Peter M.↗
NASA NTRS1991

Proximity operations considerations affecting spacecraft design

Proximity operations can be defined as the maneuvering of two or more spacecraft within 1 nautical mile range, with relative velocity less than 10 feet per second. The passive vehicle is nontranslating and should provide for maintenance of the desired approach attitude. It must accommodate the active (translating) vehicle induced structural loads and performance characteristics (mating hardware tolerances), and support sensor compatibility (transponder, visual targets, etc.). The active vehicle must provide adequate sensor systems (relative state information, field-of-view, redundancy), flight control hardware (thruster sizing, minimal cross-coupling, performance margins, redundancy) and software (reconfigurable, attitude/rate modes, translation and rotation fine control authority) characteristic, and adequate non-propulsive consumables such as power. Operational concerns must be considered. These include the following: (1) the desired approach trajectory and relative orientation; (2) the active vehicle thruster plume effects (forces, torques, contamination) on the passive vehicle; and (3) procedures for contingencies such as loss of communications, sensor or propulsion failures, and target vehicle loss of control.

Staas, Steven K.↗
NASA NTRS2011

Mission Operations Planning and Scheduling System (MOPSS)

MOPSS is a generic framework that can be configured on the fly to support a wide range of planning and scheduling applications. It is currently used to support seven missions at Goddard Space Flight Center (GSFC) in roles that include science planning, mission planning, and real-time control. Prior to MOPSS, each spacecraft project built its own planning and scheduling capability to plan satellite activities and communications and to create the commands to be uplinked to the spacecraft. This approach required creating a data repository for storing planning and scheduling information, building user interfaces to display data, generating needed scheduling algorithms, and implementing customized external interfaces. Complex scheduling problems that involved reacting to multiple variable situations were analyzed manually. Operators then used the results to add commands to the schedule. Each architecture was unique to specific satellite requirements. MOPSS is an expert system that automates mission operations and frees the flight operations team to concentrate on critical activities. It is easily reconfigured by the flight operations team as the mission evolves. The heart of the system is a custom object-oriented data layer mapped onto an Oracle relational database. The combination of these two technologies allows a user or system engineer to capture any type of scheduling or planning data in the system's generic data storage via a GUI.

Wood, Terri↗
NASA NTRS1994

A new communication protocol family for a distributed spacecraft control system

In this paper we describe the concepts behind and architecture of a communication protocol family, which was designed to fulfill the communication requirements of ESOC's new distributed spacecraft control system SCOS 2. A distributed spacecraft control system needs a data delivery subsystem to be used for telemetry (TLM) distribution, telecommand (TLC) dispatch and inter-application communication, characterized by the following properties: reliability, so that any operational workstation is guaranteed to receive the data it needs to accomplish its role; efficiency, so that the telemetry distribution, even for missions with high telemetry rates, does not cause a degradation of the overall control system performance; scalability, so that the network is not the bottleneck both in terms of bandwidth and reconfiguration; flexibility, so that it can be efficiently used in many different situations. The new protocol family which satisfies the above requirements is built on top of widely used communication protocols (UDP and TCP), provides reliable point-to-point and broadcast communication (UDP+) and is implemented in C++. Reliability is achieved using a retransmission mechanism based on a sequence numbering scheme. Such a scheme allows to have cost-effective performances compared to the traditional protocols, because retransmission is only triggered by applications which explicitly need reliability. This flexibility enables applications with different profiles to take advantage of the available protocols, so that the best rate between sped and reliability can be achieved case by case.

Baldi, Andrea↗
NASA NTRS2013

Alternating Twist Along an Erupting Prominence

Triangulation measurements using observations from the two Solar Terrestrial Relations Observatory (STEREO) spacecraft, combined with observations from the Solar Dynamics Observatory (SDO), are used to characterize the behavior of a prominence involved in two successive coronal mass ejections 6-7 December 2010. The STEREO separation at the time was 171.6 degrees, which was functionally equivalent to a separation of 8.4 degrees, and thus very favorable for feature co-identification above the limb. The first eruption at approximately 14:16 UT (Universal Time) on 6 December of the middle branch of the prominence starts off a series of magnetic reconfigurations in the right branch, which itself erupts at approximately 2:06 UT the next day, about 12 hours after the first eruption. The cool prominence material seen at 304 angstroms drains back down to the surface, but a flux-rope-like magnetic structure is seen to erupt in both 195 angstroms by the STEREO/Extreme Ultraviolet Imager (EUVI), and in white light by the STEREO/COR1 (Coronagraph 1) inner coronagraph. In between the two eruptions, two different signs of helicity are seen in the measured twist of the right branch. This is interpreted to be caused by the overall prominence channel being composed of different segments with alternating helicity signs. The erupting parts on 6 and 7 December both show positive twist, but negative twist is seen in between these positive sections. Negative twist is consistent with the dextral chirality signs seen in the He II line at 304 angstroms prior to both eruptions. However, during the period between the eruptions, a region of positive twist grows and replaces the region of negative twist, and finally erupts. We interpret these observations in the light of models that predict that helicity cancellation can be an important factor in the triggering of flares and coronal mass ejections.

Prominences↗
NASA NTRS1995

Average motion, structure and orientation of the distant magnetotail determined from remote sensing of the edge of the plasma sheet boundary layer with E greater than 35 keV ions

We study gradients of the energetic ion intesity observed at the edge of the plasma sheet boundary layer (PSBL) by the energetic ion anisotropy spectrometer (EPAS) on International Sun Earth Explorer 3 (ISEE 3). In particular, we have determined the velocity of the boundary relative to the spacecraft in the direction perpendicular to the tail axis and the angle which the boundary normal makes to the spacecraft spin axis for 1160 PSBL encounters at X(sub GSM) greater than -240 R(sub E). By asuming that, on average, the edge of the PSBL is parallel to the cross-tail current sheet, we are then able to determine a number of properties of the structure, orientation and motion of the deep geomagnetic tail. We conclude the following: (1) Most crossings of the edge of the PSBL are caused by transverse motuion of the entire tail induced by solar wind direction variations, although some are caused by reconfiguration of the tail due to geomagnetic activity. (2) The typical velocity of the PSBL (and hence of the tail) in the direction perpendicular to the tail axis is 50-85 km/s. (3) The average twist of the tail is near zero, with the edge of the PSBL (and by inference the cross-tail current sheet) lying parallel to the ecliptic plane (however, large twists are found in individual events and the distribution of twists is broad, with one standard deviation of approximately 50 deg. (4) The width of the distribution decreases with downtail distance. (5) The variation of the distributions with cross-tail position reveals that this decreas in width is most likely due to the edge of the PSBL being concave, or significantly flared at the tail flanks, in the near-Earth region. (6) During days on which the Interplanetary Magnetic Field (IMF) has 'away' sector structure, the north lobe of the trail is twisted on average towards dawn by 7.0 +/-2.4 deg. (7) During days on which the IMF has 'toward' sector structure, the north lobe is tilted towards duskby 3.8 +/- 2.3. (8) A subset of events for which IMP 8 solar wind data are available show that, for southward IMF BH(sub z) the tail has a mean twist of -12.3 +/- 5.0 deg for IMF B(sub Y) greater than 0 and 5.5 +/- 3.8 deg for IMF B(sub Y) less than 0 (positive twist angles correspond to a tilt of the northern lobe towards dusk). (9) For northward IMF B(sub z) the tail has a twist of -23.9 +/- 5.0 deg for IMF B(sub Y) greater than 0 and 13.4 +/- 6.0 deg for IMF B(sub y) less than 0. Hence the tail appears more twisted on average for the IMF B(sub Z) northward case. (10) The distribution of tail twist is wider for lower levels of geomagnetic activity, indicating that the tail is able to twist more at lower levels of activity. (11) The data set reveals no evident effect of the earth's dipole wobble; tail orientation appears to be controlled by the solar wind and IMF, such that the GSE coordinate system may be appropriate for the study of field and plasma structures in the distant tail region.

Owen, C. J.↗
NASA NTRS2001

Langmuir Probe Spacecraft Potential End Item Specification Document

This document describes the Langmuir Probe Spacecraft Potential (LPSP) investigation of the plasma environment in the vicinity of the ProSEDS Delta II spacecraft. This investigation will employ a group of three (3) Langmuir Probe Assemblies, LPAs, mounted on the Delta II second stage to measure the electron density and temperature (n(sub e) and T(sub e)), the ion density (n(sub i)), and the spacecraft potential (V(sub s)) relative to the surrounding ionospheric plasma. This document is also intended to define the technical requirements and flight-vehicle installation interfaces for the design, development, assembly, testing, qualification, and operation of the LPSP subsystem for the Propulsive Small Expendable Deployer System (ProSEDS) and its associated Ground Support Equipment (GSE). This document also defines the interfaces between the LPSP instrument and the ProSEDS Delta II spacecraft, as well as the design, fabrication, operation, and other requirements established to meet the mission objectives. The LPSP is the primary measurement instrument designed to characterize the background plasma environment and is a supporting instrument for measuring spacecraft potential of the Delta II vehicle used for the ProSEDS mission. Specifically, the LPSP will use the three LPAs equally spaced around the Delta II body to make measurements of the ambient ionospheric plasma during passive operations to aid in validating existing models of electrodynamic-tether propulsion. These same probes will also be used to measure Delta II spacecraft potential when active operations occur. When the electron emitting plasma contractor is on, dense neutral plasma is emitted. Effective operation of the plasma contactor (PC) will mean a low potential difference between the Delta II second stage and the surrounding plasma and represents one of the voltage parameters needed to fully characterize the electrodynamic-tether closed circuit. Given that the LP already needs to be well away from any near-field disturbances around the Delta II, it is possible to use the same probe with a simple reconfiguration of the electronics to measure potential with respect to the ambient plasma. The LP measurement techniques are outlined in the following text and discussed in detail in the Appendix. The scientific goals of the investigation, the physical and electrical characteristics of the instrument, and the on-orbit measurement requirements are also discussed in this document.

Gilchrist, Brian↗
NASA NTRS1994

David Florida Laboratory Thermal Vacuum Data Processing System

During 1991, the Space Simulation Facility conducted a survey to assess the requirements and analyze the merits for purchasing a new thermal vacuum data processing system for its facilities. A new, integrated, cost effective PC-based system was purchased which uses commercial off-the-shelf software for operation and control. This system can be easily reconfigured and allows its users to access a local area network. In addition, it provides superior performance compared to that of the former system which used an outdated mini-computer and peripheral hardware. This paper provides essential background on the old data processing system's features, capabilities, and the performance criteria that drove the genesis of its successor. This paper concludes with a detailed discussion of the thermal vacuum data processing system's components, features, and its important role in supporting our space-simulation environment and our capabilities for spacecraft testing. The new system was tested during the ANIK E spacecraft test, and was fully operational in November 1991.

Choueiry, Elie↗
NASA NTRS1996

ACTS Multibeam Antenna On-Orbit Performance

The Advanced Communications Technology Satellite (ACTS) launched in September 1993 introduces several new technologies including a multibeam antenna (MBA) operating at Ka-band. The MBA with fixed and rapidly reconfigurable spot beams serves users equipped with small aperture terminals within the coverage area. The antenna produces spot beams with approximately 0.3 degrees beamwidth and gains of approximately 50 dBi. A number of MBA performance evaluations have been performed since the ACTS launch. These evaluations were designed to assess MBA performance (e.g., beam pointing stability, beam shape, gain, etc.) in the space environment. The on-orbit measurements found systematic environmental perturbation to the MBA beam pointing. These perturbations were found to be imposed by satellite attitude control system, antenna and spacecraft mechanical alignments, on-orbit thermal effects, etc. As a result, the footprint coverage of the MBA may not exactly cover the intended service area at all times. This report describes the space environment effects on the ACTS MBA performance as a function of time of the day and time of the year and compensation approaches for these effects.

Acosta, R.↗