Thesis (Aeronautical and Astronautical Engineer)--Naval Postgraduate School, December 2002. / Thesis advisor(s): I. Michael Ross, Dennis Byrnes. Includes bibliographical references (p. 119-120). Also available online.

Report (M. Eng.)--Virginia Polytechnic Institute and State University, 1991. / Vita. Abstract. Includes bibliographical references (leaves 75-76). Also available via the Internet.

Thesis (M.App.Sc.) - Carleton University, 2002. / Includes bibliographical references (p. 124-130). Also available in electronic format on the Internet.

Gossamer membranes are large, ultra light weight, highly flexible thin films. They have been proposed for use as elements in systems such as solar sails and optical apertures, whose large areas require a low-mass material that can be launched in a compact package and then deployed to operational configuration upon reaching orbit. Many of the proposed applications require that the film possess a flat, wrinkle-free surface. Surface wrinkle configuration is determined, in part, by the method used to support the film. One configuration that has not been studied in detail involves the formation of vertical wrinkles oriented along the direction of a tensile force applied at the upper and lower horizontal film supports. An experiment was designed to allow known forces to be applied to a 9 inch by 9 inch square sample of film supported at its upper and lower boundaries. Four films 7.6 m and 12.7 m thick samples of Kapton (polymide), and 12.2 m and 23.4 m thick samples of Mylar were loaded at levels of applied tensile force ranging from 1.446 to 4.388 N. The out-of-plane surface contours that resulted were measured using close-range photogrammetry, a non-contact, optical measurement technique. Experimental results indicate that both wrinkle wavelength and amplitude decrease as a function of applied force magnitude. These trends matched those obtained using numerical techniques, which also showed that lateral border strain, not measured during the experiment, may be a more important factor in determining surface wrinkle configuration. After presentation of the results, the photogrammetry technique is further considered as a tool for use in the manufacturing industry, in similar close-range applications, for the measurement of both dimensions and displacements.

Thesis. 1977. M.S.--Massachusetts Institute of Technology. Dept. of Aeronautics and Astronautics. / MICROFICHE COPY AVAILABLE IN ARCHIVES AND AERO. / Bibliography : leaves 51-53. / M.S.

Thesis (Ph. D.)--University of Hawaii at Manoa, 1991. / Includes bibliographical references (leaves 152-154) / Microfiche. / xvi,154 leaves, bound ill. 29 cm

Solar sails have the potential to benefit many future space exploration missions, but they lack the heritage required for present-day use. To grow confidence in solar sail technology, they could be deployed on LEO satellites higher than 600 km to help de-orbit the satellite within 25 years upon mission termination. To determine how atomic oxygen would affect the solar sail, material from Lightsail-2 was tested in a thermal-energy, isotropic, atomic oxygen vacuum chamber based in the space environments laboratory in California Polytechnic State University. The sail material, aluminized Mylar, was tested for its survivability on both the coated and uncoated side, as well as tested for the optical degradation of the coated side. The uncoated side was found to be completely eroded after a fluence of 2.27 x1020 atoms/cm2, or ~40 days in International Space Station orbit. The coated side experienced no mass loss, but signs of significant undercutting were found with a fluence of 1.19 x1021 atoms/cm2, or ~200 days at station orbit. The stitches present on the coated side, meant to prevent tear propagation, eroded before the sample experienced a fluence of 4.13 x1020 atoms/cm2, or ~70 days at station orbit. The average total reflectivity of the material dropped by ~5% after atomic oxygen exposure, however no correlation with fluence was found. Average specular reflectivity remained unchanged after atomic oxygen exposure. The reflectivity results were impacted by wrinkling in the material, which was found to have a much larger impact than atomic oxygen exposure. These results were paired with an optimal de-orbit trajectory algorithm, developed in this thesis, to determine how atomic oxygen would affect a solar sail deployed to de-orbit an 800 km LEO satellite with a ballistic coefficient of 0.1. Using a simplified 2D orbit case, it was found that the satellite would de-orbit within 12-18 years, depending primarily on the solar activity level. The measured worst-case for optical degradation increased de-orbit time by ~6 months. Additionally, assuming that the sail material was perfectly reflecting decreased de-orbit time by 2-4 years. The amount of fluence required to erode the uncoated Mylar, and the amount required to erode the stitches, were both reached long before the satellite re-entered. It is therefore recommended that the solar sail minimize uncoated side exposure to atomic oxygen, and a more atomic oxygen-resistant stitch material be found. The fluence required to produce significant material undercutting was reached only once the satellite’s orbit had degraded to below 400 km. But the undercutting was observed to structurally compromise the material; thus, future LEO solar sail mission designers must take care when balancing added performance with higher failure risk when considering the tension in the deployed sail.

Thesis (MScEng (Electrical and Electronic Engineering))--University of Stellenbosch, 2011. / ENGLISH ABSTRACT: CubeSats are small, lightweight satellites which are often used by academic institutions due to their application potential and low cost. Because of their size and weight, less powerful attitude controllers, such as solar sails, can be used. In 2010, the Japanese satellite, Ikaros, was launched to illustrate the usage of solar sails as a propulsion system. Similarly, by exploiting the solar radiation pressure, it is possible to use a solar sail, together with three magnetorquers, to achieve 3-axis attitude control of a 3-unit CubeSat. Simulations are required to demonstrate the attitude control of a sun-synchronous, low Earth orbit CubeSat using a solar sail. To allow the adjustment of the solar sail, and its resulting torque, a mechanical structure is required which can be used to position the sail within two orthogonal axes. Although the magnetorquers and solar sail are sufficient to achieve 3-axis attitude control, the addition of a reaction wheel can be implemented in an attempt to improve this control. / AFRIKAANSE OPSOMMING: CubeSats is klein, ligte satelliete wat dikwels deur universiteite gebruik word weens hul lae koste en groot toepassings potensiaal. As gevolg van hulle gewig en grootte, kan minder kragtige posisie beheerders, soos byvoorbeeld sonseile, gebruik word. Die Japannese satelliet, Ikaros, was in 2010 gelanseer om die gebruik van ’n sonseil as aandrywingstelsel te illustreer. Net so is dit moontlik om die bestraling van die son te gebruik, met behulp van ’n sonseil, en drie magneetstange om 3-as posisiebeheer op ’n 3-eenheid CubeSat te bekom. Simulasies word benodig om die posisie beheer van ’n sonsinkrone, lae-aard wentelbaan CubeSat met ’n sonseil te demonstreer. ’n Meganiese struktuur word benodig vir die posisionering van die sonseil in twee ortogonale asse sodat die sonseil, en dus die geassosieerde draaimoment, verskuif kan word. Alhoewel die magneetstange en sonseil voldoende is om 3-as posisiebeheer te bekom, kan ’n reaksiewiel bygevoeg word om hierdie beheer te probeer verbeter.

Future solar sail and solar power satellite missions require deployment of large and lightweight flexible structures in space. One option is to spin the assembly and use the centrifugal force for deployment, stiffening and stabilization. Some of the main advantages with spin deployment are that the significant forces are in the plane of rotation, a relatively simple control can be used and the tension in the membrane or web can be adjusted by the spin rate to meet the mission requirements. However, a successful deployment requires careful development of new control schemes. The deployment rate can be controlled by a torque, applied either to a satellite in the center or by thrusters in the corners, or by deployment rate control, obtained by tether, spool braking or folding properties. Analytical models with only three degrees of freedom were here used to model the deployment of webs and membranes for various folding patterns and control schemes, with focus on space webs folded in star-like arms coiled around a center hub. The model was used to investigate control requirements and folding patterns and to obtain optimal control laws for centrifugal deployment. New control laws were derived from the optimal control results and previously presented control strategies. Analytical and finite element simulations indicate that the here developed control laws yield less oscillations, and most likely more robustness, than existing controls. Rotation-free (RF) shell elements can be used to model inflation or centrifugal deployment of flexible memebrane structures by the finite element method. RF elements approximate the rotational degrees of freedom from the out-of-plane displacements of a patch of elements, and thus avoid common singularity problems for very thin shells. The performance of RF shell elements on unstructured grids is investigated in the last article of this thesis, and it is shown that a combination of existing RF elements performs well even for unstructured grids. / QC 20100729

Thesis (MScEng (Electrical and Electronic Engineering))--University of Stellenbosch, 2011. / ENGLISH ABSTRACT: This thesis describes the development of attitude sensors required for the Attitude Determination and Control System (ADCS) for a Cubesat. The aim is to find the most suitable sensors for use on a small picosatellite by implementing miniaturised sensors with available commercial-off-the-shelf (COTS) technology. Specifically, the algorithms, hardware prototypes, software and filters required to create accurate sensors to determine the 3-axis orientation of a CubeSat are discussed. / AFRIKAANSE OPSOMMING: Hierdie tesis beskryf die ontwikkeling van oriëntasiesensors wat benodig word vir die oriëntasiebepaling en -beheerstelsel (Engels: ADCS) van ’n CubeSat. Die doelwit is om sensors te vind wat die geskikste is om in ’n klein picosatelliet te gebruik, deur miniatuursensors met kommersiële maklik verkrygbare tegnologie (Engels: COTS technology) te implementeer. Daar word in die bespreking veral aandag geskenk aan die algoritmes, hardewareprototipes, programmatuur en filters wat benodig word om akkurate sensors te skep wat op hul beurt 3-as oriëntasie van die CubeSat kan bepaal.