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Proposed microstrip reflectarray antenna provides two or more beams in different directions, without power-dividing circuitry. Because antenna flat, mass and cost of fabrication less than paraboloidal reflector with aperture of equal size, and mounted easier with less supporting structure. Designed to be mounted on ground or surfaces of ships, land vehicles, and buildings.
The primary mirror assembly (PMA) requirements and concepts for the Space Infrared Telescope Facility (SIRTF) program are discussed. The PMA studies at NASA/ARC resulted in the design of two engineering test articles, the development of a mirror mount cryogenic static load testing system, and the procurement and partial testing of a full scale spherical mirror mounting system. Preliminary analysis and testing of the single arch mirror with conical mount design and the structured mirror with the spherical mount design indicate that the designs will meet all figure and environmental requirements of the SIRTF program.
The Next Generation Space Telescope is currently envisioned as a eight meter diameter cryogenic deployable telescope that will operate at the earth sun libration point L2. A number of different designs are being examined within NASA and under industry studies by Ball Aerospace, Lockheed-Martin and TRW. Although these designs differ in many respects, they all require significant advancements in the state-of-the-art with respect to large diameter, ultra-lightweight, mirrors. The purpose of this paper is to provide insight into the current status of the mirror development program NGST is a tremendously ambitious undertaking that sets the mark for new NASA missions. In order to achieve the weight, cost and performance requirements of NGST, the primary mirror must be made lighter, cheaper and better than anything that has ever been done. In order to accomplish this an aggressive technology program has been put in place. The scope of the program was determined by examining historically what has been accomplished; assessing recent technological advances in fabrication and testing; and evaluating the effect of these advances relative to enabling the manufacture of lightweight mirrors that meet NGST requirements. As it is currently envisioned, the primary mirror for NGST is on the order of eight meters in diameter, it is to be diffraction limited at a wave length of 2 microns and has an overall weight requirement of 15 kilograms per square meter. Two large scale demonstration projects are under way along with a number of smaller scale demonstrations on a variety of mirror materials and concepts. The University of Arizona (UA) mirror concept is based around a 2mm thick Borosilicate glass face sheet mounted to a composite backplane structure via actuators for mirror figure correction. The Composite Optics Inc.(COI) concept consists of a 3.2mm thick Zerodur face sheet bonded to a composite support structure which in turn is mounted to a composite backplane structure via actuators for mirror phasing. These mirrors are due to be performance tested in ambient conditions in the fall of '98, and cryogenically tested in the spring of '99. The smaller scale efforts include the following: Beryllium is being investigated at Ball Aerospace, Electroform nickel is being investigated in-house at MSFC, Chemical Vapor Deposition (CVD) Silicon Carbide (SiC) is being investigated at Morton International Silicon mirrors are being investigated at Schafer, Carbon Fiber Reinforced Silicon Carbide (CSIC) is being investigated at IABG. SiC at SSG, Composite mirrors at COI, pyrolyzed graphite mirrors at Ultramet, reaction bonded SiC mirrors at Xinetics, along with techniques for lightweighting using waterjets at Waterjet Technology Inc. are all being investigated under the Small Business innovative Research Program SBIR program. A procurement for a third large scale demonstration (nominally 1.5m in diameter) is being planned for release this fall.
A mechanism for the docking of a space vehicle to a space station where a connection for transfer of personnel and equipment is desired. The invention comprises an active docking structure on a space vehicle 10 and a passive docking structure on a station 11. The passive structure includes a docking ring 50 mounted on a tunnel structure 35 fixed to the space station. The active structure including a docking ring 18 carried by actuator-attenuator devices 20, each attached at one end to the ring 18 and at its other end in the vehicle's payload bay 12. The devices 20 respond to command signals for moving the docking ring 18 between a stowed position in the space vehicle to a deployed position suitable for engagement with the docking ring 50. The devices 20 comprise means responsive to signals of sensed loadings to absorb impact energy and retraction means for drawing the coupled space vehicle and station into final docked configuration and moving the tunnel structure to a berthed position in the space vehicle 10. Latches 60 couple the space vehicle and space station upon contact of docking rings 18 and 50 and latches 41-48 establish a structural tie between the spacecraft when retracted.
Since mid-1980s NASA has launched several first-generation interplanetary observatory missions. A salient feature of these missions has been the mounting of the optical payload on top of a large flexible space structure. This imposes challenging control-structure interaction problems (CSI) that have been studied over several decades. The Dec. 2021 launch of the James Webb Telescope marks the latest example of such a mission. One way to encourage development of novel control methods is to provide a benchmark problem for researchers. By the mid-2000s, there were a few well-known benchmark problems developed by the control community with a focus on CSI applications. However, in the past 15 years, the control challenges in optical observation missions have shifted from CSI to line-of-sight (LOS) precision pointing. This is because popular designs of the second-generation (Gen II) space earth missions have eliminated the large structure interface by mounting the optical payload directly on top of a rigid bus platform. The focus now is to achieve accurate pointing with maximum rejection of the surrounding disturbances. The NASA Safety and Engineering Center (NESC) has tasked T he Aerospace Corporation to develop a Benchmark Problem to study the challenging control issues that Gen II observatory missions encounter. This software tool is to be released as a public domain platform aimed at government, academic, and industry researchers. The focus of this benchmark problem is for researchers to design a set of innovative payload control laws with the maximum Optical Disturbance Rejection (ODR) capability to meet a set of pre-defined μrad-level of LOS jitter requirements. This paper describes the development of the benchmark problem. Aerospace/NASA will provide an integrated LOS plant model and disturbance/command profiles for users to run their design and simulation with as well as a user’s guide describing the required interfaces. The Aerospace Corp is also working to develop a hardware fast steering mirror testbed for users to demonstrate their innovative design in real hardware, if so desired.
Fiber-optic strain sensing (FOSS) provides an opportunity for additional awareness and can enable other mitigations such as flutter suppression or gust load alleviation but requires an accurate computational model as part of a design process that integrates structures and controls. A modeling approach for fiber-optic strain sensors has been developed using a shell element formulation in a detailed finite element model. The new FOSS modeling approach is designed to better scale to more complex realistic structures and fiber layouts than previous approaches. The modeling method was applied to the X-56A vehicle and results were compared to experimental static ground-test data. Additionally, the experimental sensor noise of these fiber optic sensors was assessed for the correlations in time and correlations along the fiber. The noise does show significant correlations both in time and along the fibers. These correlations of the noise should also be considered in modeling and assessing algorithms that will use FOSS.
Fiber-optic strain sensing (FOSS) provides an opportunity for additional awareness and can enable other mitigations such as flutter suppression or gust load alleviation but requires an accurate computational model as part of a design process that integrates structures and controls. A modeling approach for fiber-optic strain sensors has been developed using a shell element formulation in a detailed finite element model. The new FOSS modeling approach is designed to better scale to more complex realistic structures and fiber layouts than previous approaches. The modeling method was applied to the X-56A vehicle and results were compared to experimental static ground-test data. Additionally, the experimental sensor noise of these fiber optic sensors was assessed for the correlations in time and correlations along the fiber. The noise does show significant correlations both in time and along the fibers. These correlations of the noise should also be considered in modeling and assessing algorithms that will use FOSS.
According to one implementation a mount pad is provided that includes a plurality of arcuate flexure members disposed between and connected to top and bottom plates. According to some implementations a majority or all of the arcuate flexure members have a semi-circular cross-section and are arranged to form a plurality of radially spaced-apart concentric rings. Concave surfaces of the arcuate flexure members face radially inward toward a center of the concentric rings. Circumferential adjacent arcuate flexure members in each of the concentric rings are circumferentially spaced apart from one another so that a gap exists between them. According to some implementations the mount pad further includes a coupling unit prolonging from one of the top and bottom plates, the coupling unit facilitating a connection of the mount pad to a support structure. According to some implementations the mount pad is a monolithic structure. According to some implementations the monolithic structure is made using an additive manufacturing method.
The results of a 3 month preliminary design and analysis effort is presented. The configuration that emerged consists of a very stiff deployable truss structure with an overall triangular cross section having universal modules attached at the apexes. Sufficient analysis was performed to show feasibility of the configuration. An evaluation of the structure shows that desirable attributes of the configuration are: (1) the solar cells, radiators, and antennas will be mounted to stiff structure to minimize control problems during orbit maintenance and correction, docking, and attitude control; (2) large flat areas are available for mounting and servicing of equipment; (3) Large mass items can be mounted near the center of gravity of the system to minimize gravity gradient torques; (4) the trusses are lightweight structures and can be transported into orbit in one Shuttle flight; (5) the trusses are expandable and will require a minimum of EVA; and (6) the modules are anticipated to be structurally identical except for internal equipment to minimize cost.
The structure, physical properties of ejecta, ballistics, and growth of Northeast crater, a young pyroclastic cone that originated in 1911 near the summit of Mount Etna, Sicily, were studied in order to form a model of volcano cinder cone growth. Four stages of growth were discerned: (1) a simple cone; (2) a cone with an outward-dipping talus slope; (3) destruction of rounded rim by the inward migration of the upper edge of the talus pile; and (4) extension of limits of talus pile beyond the ballistic limit of ejecta trajectories. The model is used to predict the features of lunar and Martian cones, assuming that they erupted under conditions qualitatively similar to Etna's Northeast crater.
The Integrated Health Management (IHM) for the future aerospace systems requires to interface models of multiple subsystems in an efficient and accurate information environment at the earlier stages of system design. The complexity of modern aeronautic and aircraft systems (including e.g. the power distribution, flight control, solid and liquid motors) dictates employment of hybrid models and high-level reasoners for analysing mixed continuous and discrete information flow involving multiple modes of operation in uncertain environments, unknown state variables, heterogeneous software and hardware components. To provide the information link between key design/performance parameters and high-level reasoners we rely on development of multi-physics performance models, distributed sensors networks, and fault diagnostic and prognostic (FD&P) technologies in close collaboration with system designers. The main challenges of our research are related to the in-flight assessment of the structural stability, engine performance, and trajectory control. The main goal is to develop an intelligent IHM that not only enhances components and system reliability, but also provides a post-flight feedback helping to optimize design of the next generation of aerospace systems. Our efforts are concentrated on several directions of the research. One of the key components of our strategy is an innovative approach to the diagnostics/prognostics based on the real time dynamical inference (DI) technologies extended to encompass hybrid systems with hidden state trajectories. The major investments are into the multiphysics performance modelling that provides an access of the FD&P technologies to the main performance parameters of e.g. solid and liquid rocket motors and composite materials of the nozzle and case. Some of the recent results of our research are discussed in this chapter. We begin by introducing the problem of dynamical inference of stochastic nonlinear models and reviewing earlier results. Next, we present our analytical approach to the solution of this problem based on the path integral formulation. The resulting algorithm does not require an extensive global search for the model parameters, provides optimal compensation for the effects of dynamical noise, and is robust for a broad range of dynamical models. In the following Section the strengths of the algorithm are illustrated illustrated by inferring the parameters of the stochastic Lorenz system and comparing the results with those of earlier research. Next, we discuss a number of recent results in application to the development of the IHM for aerospace system. Firstly, we apply dynamical inference approach to a solution of classical three tank problems with mixed unknown continuous and binary parameters. The problem is considered in the context of ground support system for filling fuel tanks of liquid rocket motors. It is shown that the DI algorithm is well suited for successful solution of a hybrid version of this benchmark problem even in the presence of additional periodic and stochastic perturbation of unknown strength. Secondly, we illustrate our approach by its application to an analysis of the nozzle fault in a solid rocket motor (SRM). The internal ballistics of the SRM is modelled as a set of one-dimensional partial differential equations coupled to the dynamics of the propellant regression. In this example we are specifically focussed on the inference of discrete and continuous parameters of the nozzle blocking fault and on the possibility of an application of the DI algorithm to reducing the probability of "misses" of an on-board FD&P for SRM. In the next section re-contact problem caused by first stage/upper stage separation failure is discussed. The reaction forces imposed on the nozzle of the upper stage during the re-contact and their connection to the nozzle damage and to the thrust vector control (TVC) signal are obtained. It is shown that transient impact induced torquean be modelled as a response of an effective damped oscillator. A possible application of the DI algorithm to the inference of damage parameters and predicting fault dynamics ahead of time using the actuator signal is discussed. Finally, we formulate Bayesian inferential framework for development of the IHM system for in-flight structural health monitoring (SHM) of composite materials. We consider the signal generated by piezoelectric actuator mounted on composite structure generating elastic waves in it. The signal received by the sensor is than compared with the baseline signal. The possibility of damage inference is discussed in the context of development of the SHM.
A modular electronics packaging system includes multiple packaging slices that are mounted horizontally to a base structure. The slices interlock to provide added structural support. Each packaging slice includes a rigid and thermally conductive housing having four side walls that together form a cavity to house an electronic circuit. The chamber is enclosed on one end by an end wall, or web, that isolates the electronic circuit from a circuit in an adjacent packaging slice. The web also provides a thermal path between the electronic circuit and the base structure. Each slice also includes a mounting bracket that connects the packaging slice to the base structure. Four guide pins protrude from the slice into four corresponding receptacles in an adjacent slice. A locking element, such as a set screw, protrudes into each receptacle and interlocks with the corresponding guide pin. A conduit is formed in the slice to allow electrical connection to the electronic circuit.
The JWST Integrated Science Instrument Module (ISIM) includes a large metering structure (approx. 2m x 2m x 1.5m) that houses the science instruments and guider. Stringent dimensional stability and repeatability requirements combined with mass limitations led to the selection of a composite bonded frame design comprised of biased laminate tubes. Even with the superb material specific stiffness, achieving the required frequency for the given mass allocations in conjunction with severe spatial limitations imposed by the instrument complement has proven challenging. In response to the challenge, the ISIM structure team considered literally over 100 primary structure topology and kinematic mount configurations, and settled on a concept comprised of over 70 m of tubes, over 50 bonded joint assemblies, and a "split bi-pod" kinematic mount configuration. In this paper, we review the evolution of the ISIM primary structure topology and kinematic mount configuration to the current baseline concept.
A conceptual design and program plan for an Orbital Construction Demonstration Article (OCDA) was developed that can be used for evaluating and establishing practical large structural assembly operations. A flight plan for initial placement and continued utility is presented as a basic for an entirely new shuttle payload line-item having great future potential benefit for space applications. The OCDA is a three-axis stabilized platform in low-earth orbit with many structural nodals for mounting large construction and fabrication equipments. This equipment would be used to explore methods for constructing the large structures for future missions. The OCDA would be supported at regular intervals by the shuttle. Construction experiments and consumables resupply are performed during shuttle visit periods. A 250 kw solar array provides sufficient power to support the shuttle while attached to the OCDA and to run construction experiments at the same time. Wide band communications with a Telemetry and Data Relay Satellite compatible high gain antenna can be used between shuttle revisits to perform remote controlled, TV assisted construction experiments.
In sounding rocket experiment CLASP, I have placed a slit a mirror-finished around the focal point of the telescope. The light reflected by the mirror surface surrounding the slit is then imaged in Slit-jaw optical system, to obtain the a-ray Lyman secondary image. This image, not only to use the real-time image in rocket flight rocket oriented direction selection, and also used as a scientific data showing the spatial structure of the Lyman alpha emission line intensity distribution and solar chromosphere around the observation area of the polarimetric spectroscope. Slit-jaw optical system is a two off-axis mirror unit part including a parabolic mirror and folding mirror, Lyman alpha transmission filter, the optical system magnification 1x consisting camera. The camera is supplied from the United States, and the other was carried out fabrication and testing in all the Japanese side. Slit-jaw optical system, it is difficult to access the structure, it is necessary to install the low place clearance. Therefore, influence the optical performance, the fine adjustment is necessary optical elements are collectively in the form of the mirror unit. On the other hand, due to the alignment of the solar sensor in the US launch site, must be removed once the Lyman alpha transmission filter holder including a filter has a different part from the mirror unit. In order to make the structure simple, stray light measures Aru to concentrate around Lyman alpha transmission filter. To overcome the difficulties of performing optical alignment in Lyman alpha wavelength absorbed by the atmosphere, it was planned 'following four steps in order to reduce standing time alignment me. 1. is measured in advance refractive index at Lyman alpha wavelength of Lyman alpha transmission filter (121.567nm), to prepare a visible light Firuwo having the same optical path length in the visible light (630nm).2. The mirror structure CLASP before mounting unit standing, dummy slit and camera standing prescribed position in leading frame is, to complete the internal alignment adjustment. 3. CLASP structure F mirror unit and by attaching the visible light filter, as will plague the focus is carried out in standing position adjustment visible flight products camera. 4. Replace the Lyman alpha transmission filter, it is confirmed by Lyman alpha wavelength (under vacuum) the requested optical performance have come. Currently, up to 3 of the steps completed, it was confirmed in the visible light optical performance that satisfies the required value sufficiently extended. Also, put in Slit-jaw optical system the sunlight through the telescope of CLASP, it is also confirmed that and that stray light rejection no vignetting is in the field of view meets request standing.
In sounding rocket experiment CLASP, I have placed a slit a mirror‐finished around the focal point of the telescope. The light reflected by the mirror surface surrounding the slit is then imaged in Slit‐jaw optical system, to obtain the alpha‐ray Lyman secondary image. This image, not only to use the real‐time image in rocket flight rocket oriented direction selection, and also used as a scientific data showing the spatial structure of the Lyman alpha emission line intensity distribution and solar chromosphere around the observation area of the polarimetric spectroscope. Slit‐jaw optical system is a two off‐axis mirror unit part including a parabolic mirror and folding mirror, Lyman alpha transmission filter, the optical system magnification 1x consisting camera. The camera is supplied from the United States, and the other was carried out fabrication and testing in all the Japanese side. Slit‐jaw optical system, it is difficult to access the structure, it is necessary to install the low place clearance. Therefore, influence the optical performance, the fine adjustment is necessary optical elements are collectively in the form of the mirror unit. On the other hand, due to the alignment of the solar sensor in the US launch site, must be removed once the Lyman alpha transmission filter holder including a filter has a different part from the mirror unit. In order to make the structure simple, stray light measures Aru to concentrate around Lyman alpha transmission filter. To overcome the difficulties of performing optical alignment in Lyman alpha wavelength absorbed by the atmosphere, it was planned following four steps in order to reduce standing time alignment me. 1: is measured in advance refractive index at Lyman alpha wavelength of Lyman alpha transmission filter (121.567nm), to prepare a visible light Firuwo having the same optical path length in the visible light (630nm). 2: The mirror structure CLASP before mounting unit standing, dummy slit and camera standing prescribed position in leading frame is, to complete the internal alignment adjustment. 3: CLASP structure F mirror unit and by attaching the visible light filter, as will plague the focus is carried out in standing position adjustment visible flight products camera. 4: Replace the Lyman alpha transmission filter, it is confirmed by Lyman alpha wavelength (under vacuum) the requested optical performance have come. Currently, up to 3 of the steps completed, it was confirmed in the visible light optical performance that satisfies the required value sufficiently extended. Also, put in Slit‐jaw optical system the sunlight through the telescope of CLASP, it is also confirmed that and that stray light rejection no vignetting is in the field of view meets request standing.
Many studies are being made to determine the most efficient method for the transfer of large space structures from low earth orbits (LEO) to higher altitudes, principally geosynchronous orbits. In this paper, towing is proposed as a possible means of orbital transfer of large space structures. A potential towing arrangement is identified and some of the systems sized for the transfer of a 2460-ft (750 meter) diameter antenna weighing approximately 154,000 lbs (70 metric tons). A savings of approximately 23 metric tons in propellants and structure resulted from the use of a towing arrangement using six towlines in parallel when compared to the use of a single tug hard mounted to the structure being towed. The towing concept appears practical and affords a means of applying loads using multiple towlines to a wide variety of structures without special adaptors.
The Long Duration Exposure Facility (LDEF) is being developed to accommodate experiments which require a free-flying exposure in space and which benefit from postflight laboratory analysis of the retrieved experiment hardware. The first LDEF mission, which is planned for a 6- to 12-month stay in space is scheduled in the late 1970s during the Shuttle Orbital Flight Test Program. The LDEF is a simple reusable structure which is three-axis gravity gradient stabilized when free flying in space. The experiments on the LDEF are totally self-contained in trays mounted to the structure. After the LDEF is either revisited or retrieved by the Shuttle, the trays with experiments will be returned to the experimenters for postflight inspection. Attention is given to the fabrication of a second LDEF for a long-term mission planned in 1979, aspects of LDEF orbit life time, the radiation environment, questions of solar exposure, the thermal environment, and the vacuum (atomic particle) environment.