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All sourcesNASA NTRSDOE OSTI
NASA NTRS2016

Design and Development of NEA Scout Solar Sail Deployer Mechanism

The 6U (approx.10 cm x 20 cm x 30 cm) cubesat Near Earth Asteroid (NEA) Scout1, projected for launch in September 2018 aboard the maiden voyage of the Space Launch System, will utilize a solar sail as its main method of propulsion throughout its approx.3-year mission to a Near Earth Asteroid. Due to the extreme volume constraints levied onto the mission, an acutely compact solar sail deployment mechanism has been designed to meet the volume and mass constraints, as well as provide enough propulsive solar sail area and quality in order to achieve mission success. The design of such a compact system required the development of approximately half a dozen prototypes in order to identify unforeseen problems, advance solutions, and build confidence in the final design product. This paper focuses on the obstacles of developing a solar sail deployment mechanism for such an application and the lessons learned from a thorough development process. The lessons presented will have significant applications beyond the NEA Scout mission, such as the development of other deployable boom mechanisms and uses for gossamer-thin films in space.

Sobey, Alexander R.↗
NASA NTRS2004

Static Testing of an Inflatable/Rigidizable Hexapod Structure

Scientific application of large space antennas and telescopes is fostering the study of ultralightweight and inflatable structures. These so-called gossamer structures are tightly packaged for launch and subsequent deployment. Of particular interest is a hexapod configuration often found in vehicle ride simulators, ground-based telescopes, and antennas. One such configuration was modeled and tested at NASA Langley. Discrepancies between dynamic test data and predictions from a finite element model prompted additional static testing. This paper discusses static tests that were conducted to update stiffness parameters in the finite element model.

Berger, K. T.↗
NASA NTRS2005

Stable Polyimides for Terrestrial and Space Uses

Polyimides of a recently developed type have an attractive combination of properties, including low solar absorptivity (manifested as low color) when cast into thin films, resistance to atomic oxygen and ultraviolet radiation, solubility in organic solvents, high glass-transition temperatures, and high thermal stability. The focus of the development work was on polymers that can endure the space environment and that have specific combinations of properties for use on Gossamer spacecraft. Because of their unique combination of properties, these polymers are also expected to find use in a variety of other applications on Earth as well as in space. Examples of other space applications include membranes on antennas, second-surface mirrors, thermal optical coatings, and multilayer thermal insulation. For both terrestrial and space applications, these polyimides can be processed into various forms, including films, fibers, foams, threads, adhesives, and coatings.

Connell, John W.↗
NASA NTRS2002

Conformal Membrane Reflectors for Deployable Optics

This presentation reports the Phase I results on NASA's Gossamer Spacecraft Exploratory Research and Technology Program. Cornerstone Research Group, Inc., the University of Rochester, and International Photonics Consultants collaborated to investigate the feasibility of free-standing, liquid-crystal-polymer (LCP) reflectors for integration into space-based optical systems. The goal of the program was to achieve large-diameter, broadband. reflective membranes that are resistant to the effects of space, specifically cryogenic environments and gamma-ray irradiation. Additionally, we assessed the applicability of utilizing the technology as tight sails, since, by their very nature, these films offer high-reflectivity at specified wavelengths. Previous research programs have demonstrated all-polymer, narrow-band Specular reflectors and diffuse membrane reflectors. The feasibility of fabricating an all-polymer broadband specular reflector and a narrow-band specular membrane reflector was assessed in the Phase I Gossamer program. In addition, preliminary gamma irradiation studies were conducted to determine the stability of the polymer reflectors to radiation. Materials and process technology were developed to fabricate coupon-scale reflectors of both broad- and narrow-band specular reflectors in Phase 1. This presentation will report the results of these studies, including, the performance of a narrow-band specular membrane. Gamma irradiation exposures indicate limited impact on the optical performance although additional exposure studies are warranted. Plans to scale up the membrane fabrication process will be presented.

Hood, Patrick J.↗
NASA NTRS2003

Dot-Projection Photogrammetry and Videogrammetry of Gossamer Space Structures

This paper documents the technique of using hundreds or thousands of projected dots of light as targets for photogrammetry and videogrammetry of gossamer space structures. Photogrammetry calculates the three-dimensional coordinates of each target on the structure, and videogrammetry tracks the coordinates versus time. Gossamer structures characteristically contain large areas of delicate, thin-film membranes. Examples include solar sails, large antennas, inflatable solar arrays, solar power concentrators and transmitters, sun shields, and planetary balloons and habitats. Using projected-dot targets avoids the unwanted mass, stiffness, and installation costs of traditional retroreflective adhesive targets. Four laboratory applications are covered that demonstrate the practical effectiveness of white-light dot projection for both static-shape and dynamic measurement of reflective and diffuse surfaces, respectively. Comparisons are made between dot-projection videogrammetry and traditional laser vibrometry for membrane vibration measurements. The paper closes by introducing a promising extension of existing techniques using a novel laser-induced fluorescence approach.

Pappa, Richard S.↗
NASA NTRS2002

Compositionally-Graded Shape Memory Film for Self-Deployment of Membrane Reflectors and Optics

The next generation of space systems will require large apertures in order to image faint targets or cover large areas of Earth. These large apertures must be able to fit inside a launch vehicle fairing, be light enough for launch into orbit, and deploy on orbit with repeatability and reliability. The current state-of-the-art in flight optics is represented by the 4 meter LAMP telescope, with an areal density of 10 km sq m. Development of a Beryllium mirror demonstration article for NGST (Next Generation Space Telescope) at the University of Arizona indicate areal densities of 0.5 kg sq m with flight hardware in the 12 meter range. With progressive improvements in existing deployment, packaging, and structural technologies, the size of optics and reflectors will continue to increase, while mass is reduced. However, without a breakthrough in materials, packaging and/or deployment technologies, the goal for Gossamer structures of 0.1 kg sq m is unachievable for the near and mid-term NASA missions. Membrane technology provides the best hope of achieving such low areal densities. In combination with advances in membrane materials and structures, development of revolutionary techniques for deployment systems can provide significant improvements in large aperture technology. In this paper, the results of a six-month Phase I research effort to demonstrate the application of thin film NiTi to aerospace-qualified membrane and mesh materials are presented. Deposition of shape memory thin film was achieved Astromesh (trademark) metal mesh and CP-1, and optical-quality polymer membrane. Not only was full-coating deposition demonstrated, but also small segment deposition which holds potential for local surface control. Deployment of these materials was also demonstrated, setting the stage for the development of a larger test article.

Hill, Lisa↗
NASA NTRS2006

Material Interactions with Solar Wind Ion Environments

Solar wind composition is dominated by hydrogen (approx.96%) and helium (approx.3 to 4%) with a minor fraction (less than or equal to 1%) of heavy ions. Hydrogen (helium) ions impact spacecraft surfaces with energies from 0.5 to 5 keV (1.8 to 21 keV) due to variations in solar wind velocity from 300 km/s to 1000 km/sec with extremes of a few 10 s keV during periods of extremely high solar wind velocity exceeding 1000 km/sec. Mean impact energies are typically on the order of approximately 1 keV and 4 keV for hydrogen ions and helium ions, respectively. These energies are typically of the peak of the energy dependent light ion sputter yields for hydrogen and helium on many metals. In addition, light ions with kilovolt energies have been shown to produce blister (or exfoliation) damage to metal surfaces due to formation of high pressure gas bubbles within the materials when exposed to ion fluences on the order of 10(exp 16 to (10(exp 17 ions/sq cm. A number of spacecraft designs for current and future missions include gossamer polymer structures with thin metallic reflection coatings to shield instruments from the Sun or solar sail propulsion systems for use in a variety of locations in the inner solar system from 0.5 to 1 AU. In addition, there is interest in designing spacecraft for solar physics missions requiring operations as close to the Sun as 0.16 to 0.2 AU. Integrity of the metallic coatings is critical in many of these applications since degradation will result in modification of material thermal properties or exposure of polymers to solar UV photons which can compromise mission requirements. This paper will evaluate the relative contributions of sputtering and blister formation to material degradation in solar wind environments over a range of radial distances from the Sun to demonstrate where solar wind environments become important for materials selection. We will first review the physics and results from laboratory measurements of light ion sputtering, blistering, and exfoliation of metallic surfaces to establish the order of magnitude ion fluence required for significant surface damage. Solar wind ion fluence environments will then be evaluated due to variations in solar wind conditions as a function of solar cycle for varying distances from the Sun using models for radial variations in solar wind ion number density, temperature, and velocity to determine where sputtering and blistering is most likely to be an issue. Finally, ion fluence statistics for varying radial distances from the Sun will be shown to establish the mission duration and radial distances from the Sun where missions will encounter sufficient ion fluence to exhibit damage to metallic surfaces.

Minow, Joseph I.↗
NASA NTRS2004

Development of Modal Test Techniques for Validation of a Solar Sail Design

This paper focuses on the development of modal test techniques for validation of a solar sail gossamer space structure design. The major focus is on validating and comparing the capabilities of various excitation techniques for modal testing solar sail components. One triangular shaped quadrant of a solar sail membrane was tested in a 1 Torr vacuum environment using various excitation techniques including, magnetic excitation, and surface-bonded piezoelectric patch actuators. Results from modal tests performed on the sail using piezoelectric patches at different positions are discussed. The excitation methods were evaluated for their applicability to in-vacuum ground testing and to the development of on orbit flight test techniques. The solar sail membrane was tested in the horizontal configuration at various tension levels to assess the variation in frequency with tension in a vacuum environment. A segment of a solar sail mast prototype was also tested in ambient atmospheric conditions using various excitation techniques, and these methods are also assessed for their ground test capabilities and on-orbit flight testing.

Gaspar, James L.↗
NASA NTRS2006

Stretched Lens Array (SLA) for Collection and Conversion of Infrared Laser Light: 45% Efficiency Demonstrated for Near-Term 800 W/kg Space Power System

For the past 2% years, our team has been developing a unique photovoltaic concentrator array for collection and conversion of infrared laser light. This laser-receiving array has evolved from the solar-receiving Stretched Lens Array (SLA). The laser-receiving version of SLA is being developed for space power applications when or where sunlight is not available (e.g., the eternally dark lunar polar craters). The laser-receiving SLA can efficiently collect and convert beamed laser power from orbiting spacecraft or other sources (e.g., solar-powered lasers on the permanently illuminated ridges of lunar polar craters). A dual-use version of SLA can produce power from sunlight during sunlit portions of the mission, and from beamed laser light during dark portions of the mission. SLA minimizes the cost and mass of photovoltaic cells by using gossamer-like Fresnel lenses to capture and focus incoming light (solar or laser) by a factor of 8.5X, thereby providing a cost-effective, ultra-light space power system.

O'Neill, Mark↗
NASA NTRS2001

Development of an inflatable SAR engineering model

Deployable and inflatable structures are successfully used in a wide variety of earth-bound applications. Their use in space, however, requires not only assurance that the structure will survive a launch to space, but also certainly that the desired deployed configuration and that the processes of deployment, inflation and rigidization in space will be successful and without posing a threat to the hosting satellite or vessel.

gossamer↗
NASA NTRS2004

Study of ripple formation in unidirectionally-tensioned membranes

The study of membrane behavior is one of the areas of interest in the development of ultralightweight and lightweight structures for space applications. Utilization of membranes as loadcarrying components or support structure for antenna patch-arrays, collectors, sun-shades and solar-sail reflective surfaces brings about a variety of challenges that require understanding of the ripple-formation phenomenology, development of reliable test and analysis techniques, and solution methods for challenges related to the intended applications. This paper presents interim results from a study on the behavior of unidirectionally tensioned flat and singly-curved membranes. It focuses on preliminary experimental work to explore formation of ripples' and on finite element analysis (FEA) to correlate and predict their formation on thin polyimide membrane models.

radar↗
NASA NTRS2010

Self-Deployable Membrane Structures

Currently existing approaches for deployment of large, ultra-lightweight gossamer structures in space rely typically upon electromechanical mechanisms and mechanically expandable or inflatable booms for deployment and to maintain them in a fully deployed, operational configuration. These support structures, with the associated deployment mechanisms, launch restraints, inflation systems, and controls, can comprise more than 90 percent of the total mass budget. In addition, they significantly increase the stowage volume, cost, and complexity. A CHEM (cold hibernated elastic memory) membrane structure without any deployable mechanism and support booms/structure is deployed by using shape memory and elastic recovery. The use of CHEM micro-foams reinforced with carbon nanotubes is considered for thin-membrane structure applications. In this advanced structural concept, the CHEM membrane structure is warmed up to allow packaging and stowing prior to launch, and then cooled to induce hibernation of the internal restoring forces. In space, the membrane remembers its original shape and size when warmed up. After the internal restoring forces deploy the structure, it is then cooled to achieve rigidization. For this type of structure, the solar radiation could be utilized as the heat energy used for deployment and space ambient temperature for rigidization. The overall simplicity of the CHEM self-deployable membrane is one of its greatest assets. In present approaches to space-deployable structures, the stow age and deployment are difficult and challenging, and introduce a significant risk, heavy mass, and high cost. Simple procedures provided by CHEM membrane greatly simplify the overall end-to-end process for designing, fabricating, deploying, and rigidizing large structures. The CHEM membrane avoids the complexities associated with other methods for deploying and rigidizing structures by eliminating deployable booms, deployment mechanisms, and inflation and control systems that can use up the majority of the mass budget

Sokolowski, Witold M.↗
NASA NTRSDate not supplied

Design and Testing of the BionicWingSat in a Zero-g Flight Campaign - A 2U-CubeSat with Deployable, Biologically-Inspired Wings

In this paper, recent developments in the design, manufacturing, and testing of a novel deployable structure with several potential applications in space will be described. Through a cooperative effort of the German Aerospace Center (DLR) and the National Aeronautics and Space Administration (NASA), a biologically inspired structurally integrated membrane featuring distributed functional elements has been developed and tested in a 2U CubeSat called BionicWingSat. Such a membrane structure could be useful for several applications in which a relatively flat area is desirable such as solar sails, drag sails, or solar shades. For SmallSats and CubeSats, the design proposed also has the desirable property of being self-deploying without the need for powered deployment mechanisms. Building on previous work inspired by the wings of earwigs, the research presented in this paper includes structural design of self-deploying hinges, a survey of various advanced additive layer manufacturing (ALM) methods for making hinges, mechanical characterization of the hinges, and finite element analysis (FEA) of the hinges. In this work, the conflicting goals of maximizing deployed structural stiffness, maximizing deployed area, maximizing stowed packaging efficiency, and maximizing resistance to creep when stowed must be considered. The resulting design concept is a gossamer structure that cannot support its own weight in gravity. For this reason, a focus in this paper is on a parabolic flight test campaign in which 24 fully integrated wings on two BionicWingSats were tested in a microgravity environment. From this test campaign, several lessons were learned regarding the wing design and procedures for carrying out microgravity tests of this manner.

CubeSat↗
NASA NTRSDate not supplied

Design and Testing of the BionicWingSat in a Zero-g Flight Campaign - A 2U-CubeSat with Deployable, Biologically-Inspired Wings

In this paper, recent developments in the design, manufacturing, and testing of a novel deployable structure with several potential applications in space will be described. Through a cooperative effort of the German Aerospace Center (DLR) and the National Aeronautics and Space Administration (NASA), a biologically inspired structurally integrated membrane featuring distributed functional elements has been developed and tested in a 2U CubeSat called BionicWingSat. Such a membrane structure could be useful for several applications in which a relatively flat area is desirable such as solar sails, drag sails, or solar shades. For SmallSats and CubeSats, the design proposed also has the desirable property of being self-deploying without the need for powered deployment mechanisms. Building on previous work inspired by the wings of earwigs, the research presented in this paper includes structural design of self-deploying hinges, a survey of various advanced additive layer manufacturing (ALM) methods for making hinges, mechanical characterization of the hinges, and finite element analysis (FEA) of the hinges. In this work, the conflicting goals of maximizing deployed structural stiffness, maximizing deployed area, maximizing stowed packaging efficiency, and maximizing resistance to creep when stowed must be considered. The resulting design concept is a gossamer structure that cannot support its own weight in gravity. For this reason, a focus in this paper is on a parabolic flight test campaign in which 24 fully integrated wings on two BionicWingSats were tested in a microgravity environment. From this test campaign, several lessons were learned regarding the wing design and procedures for carrying out microgravity tests of this manner.

CubeSat↗
NASA NTRS1988

Flight test results for the Daedalus and Light Eagle human powered aircraft

The results of the flight test program of the Daedalus and Light Eagle human powered aircraft in the winter of 1987/88 are given. The results from experiments exploring the Light Eagle's rigid body and structural dynamics are presented. The interactions of these dynamics with the autopilot design are investigated. Estimates of the power required to fly the Daedalus aircraft are detailed. The system of sensors, signal conditioning boards, and data acquisition equipment used to record the flight data is also described. In order to investigate the dynamics of the aircraft, flight test maneuvers were developed to yield maximum data quality from the point of view of estimating lateral and longitudinal stability derivatives. From this data, structural flexibility and unsteady aerodynamics have been modeled in an ad hoc manner and are used to augment the equations of motion with flexibility effects. Results of maneuvers that were flown are compared with the predictions from the flexibility model. To extend the ad hoc flexibility model, a fully flexible aeroelastic model has been developed. The model is unusual in the approximate equality of many structural natural frequencies and the importance of unsteady aerodynamic effects. the Gossamer Albatross. It is hypothesized that this inverse ground effect is caused by turbulence in the Earth's boundary layer. The diameters of the largest boundary layer eddies (which represent most of the turbulent kinetic energy) are proportional to altitude; thus, closer to the ground, the energy in the boundary layer becomes concentrated in eddies of smaller and smaller diameter. Eventually the eddies become sufficiently small (approximately 0.5 cm) that they trip the laminar boundary layer on the wing. As a result, a greater percentage of the wing area is covered with turbulent flow. Consequently the aircraft's drag and the pow er required both increase as the aircraft flies closer to the ground. The results of the flight test program are examined critically, especially with respect to future applications. Maneuvers that allow the observation of stability derivatives for flexible aircraft are detailed. Considerations for the design of autopilots for future human power aircraft and high-altitude RPV's are discussed. Techniques useful in estimating the power required to fly aircraft of very high lift to drag ratio are described.

Sullivan, R. Bryan↗
NASA NTRS2010

Gossamer Technology to Deorbit LEO Non-Propulsion Fitted Satellite

Since 2004, CNES has decided to apply the end of life Code of Conduct rules to debris mitigation. Originally drawn up by the main European space agencies, it contains basic rules to be applied in space in order to limit the increase of orbital debris. In low Earth orbit, the rule is to limit in-orbit lifetime to 25 years after the end of the operational mission, or else to transfer to a graveyard orbit above 2000 km. In order to follow these instructions, a task force was set up in 2005 to find the best way to implement them on MICROSCOPE and CNES microsatellite family (MYRIADE). This 200-kg spacecraft should be launched in 2014 on a 790-km high circular orbit. Without targeted action, its natural re-entry would occur in 67 years. Two strategies to reduce this time period were compared: propulsive maneuvers at the end of the mission or the deployment of large surfaces to increase significantly the ballistic coefficient. At the end of the trade off, it was recommended: .. For the non-propulsive system fitted satellites, to use passive aerobraking by deployment of added surface, .. For satellites having propulsive subsystem in baseline for mission purposes, to keep sufficient propellant and implement specific maneuvers. The poster gives an overview of the process that led to the development of a deployable aerobraking wing using a lightweight aluminized Kapton membrane and an inflatable aluminum laminate boom. The main requirements; The trade off among various aerobraking solutions; The development plan. This technology presents a very attractive potential and it could be a first step in using of inflatable technology on spaces vehicles, before to deal with others more exigent applications.

Dupuy, C.↗