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The IN-STEP Inflatable Antenna Experiment was intended to demonstrate high payoff technology for large, inflatable space antenna structures, in a realistic operational environment.
A system level design and analysis has been conducted on an Earth Observation Satellite (EOS) system using the Interactive Design and Evaluation of Advanced Spacecraft (IDEAS) computer-aided design and analysis program. The IDEAS program consists of about 40 user-friendly technical modules and an interactive graphics display. The reflector support system and feed mast of the EOS spacecraft are constructed with a box-truss structural concept, a lattice configuration which can be packaged for delivery in a single Shuttle flight and deployed in orbit. The deployed spacecraft consists of a 120-m by 60-m parabolic reflector, a 120-m-long support structure, and a 30-m feed arc beam on the focal axis. The spacecraft was modeled for structural, thermal, and control systems analysis and structural elements were designed. On-orbit dynamic and thermal loading analyses were conducted and spacecraft weights were determined.
The objective of this paper is to explore possibilities for continued use of certain deployable spacecraft antennas when full deployment fails. Four (4) deployables are used as examples, including 1) a reflectarray, 2) a parabolic reflector, 3) a loop, and 4) a set of 4 monopoles. The measure of “continued usability” in each case is relative antenna gain as a function of partial deployment. Note, types (and levels) of partial deployment presented here are strictly hypothetical. Also, discussions of actual deployment mechanisms are considered outside the scope of this paper.
A deployable antenna is described. The antenna comprises a mesh attached to foldable ribs, a hub and a sub-reflector. The antenna can be stowed in a tight space for launching in space, and later deployed by extending out of its container. The antenna is designed to work in the Ka band or other bands and can increase data rates and function as a radio antenna.
This paper summarizes some of the future civil missions requiring large space systems technologies. Antenna, collector, and reflector missions are generalized to define a similar set of system requirements and characteristics. Although many concepts exist for both deployable and space assemblable large structures, four technically mature deployable concepts are reviewed. Two of these concepts are probably applicable to only antenna/collector missions, whereas the other two employ continuous trusses which can be configured for a broad range of planar, linear, or curved structures. Finally, technology problems or needs associated with large deployable systems are reviewed to highlight additional research and development, both analytical and experimental, required to reduce mission risk.
Practical approaches for establishing large, precise antenna reflectors in space are described. Reflector surfaces consisting of either solid panels or knitted mesh are considered. The approach using a deep articulated truss structure to support a mesh reflector is selected for detailed investigations. A new sequential deployment concept for the tetrahedral truss is explained. Good joint design is discussed, and examples are described both analytically and by means of demonstration models. The influence of curvature on the design and its vibration characteristics are investigated.
Spaceborne experiments to test the deployment reliability of large space antennas are discussed. Retraction, reflector surface tolerance, thermal distortion, electromagnetic performance, and dynamic controls are discussed.
A NASA R&D program, the Large Deployable Antenna program, was initiated to investigate and demonstrate the availability of critical technologies for passive microwave imagers. The science objectives, current state-of-the-art, a number of electromagnetic configurations under consideration, and the mechanical systems development effort are presented. The program team conducted a detailed technology review, evaluated the feasibility and technology readiness for a large dual-reflector radiometer, and developed a system concept for a 25-meter deployable radiometer. The study approach involved determining basic operational parameters and configurations for a geosynchronous wide-scanning radiometer from which specific structural requirements were utilized as goals (rather than specifications) with which specific technologies could be evaluated.
The overall mission and spacecraft systems, testing, and operations are summarized. The mechanical subsystems are reviewed, encompassing mechanical design requirements; separation and deployment mechanisms; design and performance evaluation; and the television camera reflector monitor. Thermal control and contamination are discussed in terms of thermal control subsystems, design validation, subsystems performance, the advanced flight experiment, and the quartz-crystal microbalance contamination monitor.
The wrap rib antenna design of a parabolic reflector large space antenna is discussed. Cost estimates, design/mission compatibility, deployment sequence, ground based tests, and fabrication are discussed.
The science objectives of the HYDROS mission are to provide frequent, global measurements of surface soil moisture and surface freeze/thaw state. In order to adequately measure these geophysical quantities, the key instrument requirements were determined by the HYDROS science team to be: 1) Dual-polarization L-Band passive radiometer measurements at 40 km resolution, 2) Dual-polarization L-Band active radar measurements at 3 km resolution, and 3) A wide swath to insure global three day refresh time for these measurements (1000 km swath at the selected orbit altitude of 670 km). As a solution to this challenging set of instrument requirements, a relatively large, 6 meter, conically-scanning reflector antenna architecture was selected for the instrument design. The deployable mesh antenna is shared by both the radiometer and radar electronics by employing a single L-Band feed.
Surface errors on parabolic reflector antennas degrade the overall performance of the antenna. Space antenna structures are difficult to build, deploy and control. They must maintain a nearly perfect parabolic shape in a harsh environment and must be lightweight. Electromagnetic compensation for surface errors in large space reflector antennas can be used to supplement mechanical compensation. Electromagnetic compensation for surface errors in large space reflector antennas has been the topic of several research studies. Most of these studies try to correct the focal plane fields of the reflector near the focal point and, hence, compensate for the distortions over the whole radiation pattern. An alternative approach to electromagnetic compensation is presented. The proposed technique uses pattern synthesis to compensate for the surface errors. The pattern synthesis approach uses a localized algorithm in which pattern corrections are directed specifically towards portions of the pattern requiring improvement. The pattern synthesis technique does not require knowledge of the reflector surface. It uses radiation pattern data to perform the compensation.
The Naval Research Laboratory (NRL) has designed, built, and fully qualified a low cost, low Passive Intermodulation (PIM) 12-foot (3.66-m) diameter deployable ultra high frequency (UHF) antenna for the Tacsat-4 program. The design utilized novel approaches in reflector material and capacitive coupling techniques. This paper discusses major design trades, unique design characteristics, and lessons learned from the development of the Tacsat 4 deployable antenna. This antenna development was sponsored by the Office of Naval Research.
The design and test details of a high precision hinge and locking mechanism created principally in support of the Lockheed space deployable Flexrib Parabolic Antenna concept are presented. These developed improvements in the detail design of ultra-precise rib hinge and contour adjustment mechanisms will allow radio frequency antenna reflectors to expand into the 15-50 meter diameter size class and to support the .013 mm. deployment repeatability tolerance required to support the 12+ GHz frequency range.
The design technology is described for manufacturing a 20 m or larger space erectable antenna with high thermal stability, high dynamic stiffness, and minimum stowed size. The selected approach includes a wrap rib design with a cantilever beam basic element and graphite-epoxy composite lenticular cross section ribs. The rib configuration and powered type operated deploying mechanism are described and illustrated. Other features of the parabolic reflector discussed include weight and stowed diameter characteristics, structural dynamics characteristics, orbit thermal aperture limitations, and equivalent element and secondary (on axis) patterns. A block diagram of the multiple beam pattern is also presented.
The preliminary analysis of the C-band cross-track interferometric data (XTI) acquired during the MAC Europe 1991 campaign over the Matera test site, in Southern Italy is presented. Twenty three passive calibrators (Corner Reflector, CR) and 3 active calibrators (Active Radar Calibrator, ARC) were deployed over an area characterized by homogeneous background. Contemporaneously to the flight, a ground truth data collection campaign was carried out. The research activity was focused on the development of motion compensation algorithms, in order to improve the height measurement accuracy of the TOPSAR system.
NASA Langley Research center is developing key structures and mechanisms technologies for micron-accuracy, in-space deployment of future space instruments. Achieving micron-accuracy deployment requires significant advancements in deployment mechanism design such as the revolute joint presented herein. The joint presented herein exhibits a load-cycling response that is essentially linear with less than two percent hysteresis, and the joint rotates with less than one in.-oz. of resistance. A prototype reflector metering truss incorporating the joint exhibits only a few microns of kinematic error under repeated deployment and impulse loading. No other mechanically deployable structure found in literature has been demonstrated to be this kinematically accurate.
A major source of uncertainty in climate models is the presence, shape and distribution of ice particles in the uppermost layers of the clouds. The effects of this component are poorly constrained, turning ice particles into an almost-free variable in many climate models.NASA-GSFC is developing a new instrument aimed at measuring the size and shape of ice particles. The instrument consists of two sub-mm polarimeters (at 220 and 670 GHz) coupled with a long-wave infrared polarimeter at 10 micron. Each polarimeter has identical V-pol and H-pol channels; the axes of polarization are defined geometrically by the orientation of the waveguide elements, and the purity has been measured in the lab. The instrument is configured as a conical scanner, suitable for deployment as a payload on a small satellite or on a high-altitude sub-orbital platform. From a 400 km orbit, the instrument has a 3dB spatial resolution of 20 (10) km at 220 (670) GHz and a swath of 600 km over 180 degrees of view.The BAPTA (Bearing And Power Transfer Assembly) carries heritage from the SSMIS design, now in its 22nd year of on-orbit operation, but with a much reduced SWaP (Size Weight and Power) footprint, suitable for a small satellite.The main components of the instrument have been fabricated and are undergoing final testing prior to their integration as a single unit. The sub-mm channels have dedicated secondary reflectors which illuminate a shared primary reflector. The receiving units are placed behind the focal point of the optical arrangement, so that all beams equally illuminate the primary reflector and are almost co-located on the ground (within a single 220 GHz footprint). Primary and secondary beam patterns have been measured and verified to match the as-designed expectations. A Zytex (TM) window is deployed to protect the secondary reflectors and the feed horns from debris and other contaminants, and to reduce the heat load from the active (hot) IR calibration unit. The insertion loss of Zytex has been measured and is accounted in the calibration equation of the sub-mm channels.The radiometric performance of the sub-mm receivers has been characterized in the lab and under operational conditions of temperature and pressure.This paper discusses the design constraints on the sub-mm components, details of the scientific goals and their flowdown, and describes the characterization of the polarimeters. Options to optimize the layout and distribution of the masses within the assembly, with the goal of making the instrument even more compact and fully-compatible with cubesat-class satellites will be presented.