NASA's “Expanding the Horizons of Mars Science” report emphasizes the need for low-cost, smaller satellites for future Mars missions, accompanied by enhanced Mars-Earth telecommunications. Current Mars telecom relay orbiters cost 400 M-800 M (adjusted for inflation), with rising costs outpacing budgets. Leveraging deployable antennas, such as Tendeg's perimeter truss reflector (PTR), could enable small spacecraft (<500 kg) to match the capabilities of larger relays at a fraction of the cost (>100 M per spacecraft). To reuse a 3 -meter GEO PTR for Mars, it is important to develop the requirements for the Martian application of the antenna, and understand what differences exist between GEO and Mars. Key requirements include packaging and spacecraft interface, optical prescription and geometric accuracy, and environmental. Environmental requirements include those related to vibration, radiation, thermal and aerobraking. Future work will validate the PTR design against Mars-specific requirements defined here, focusing on cold-temperature performance.

To access the interstellar medium with current approaches requires 30 to 40 years, significantly longer than most mission lifetimes. The goal of this study was to explore mission concepts that will reach the interstellar medium in a primary mission’s lifetime (15 years or less). Faster access to the interstellar medium would allow high-capability science probes, with many relevant instruments, to explore the galaxy beyond our solar system in-situ. Science targets include the hydrogen wall structure, bow wave/shock, gravitational lens, foreground emissions and interstellar dust, just to name a few. Further, such a capability would enable rapid exploration of Kuiper belt objects in a much shorter time frame than current methods. Finally, distant targets include the solar gravitational lens, which may enable direct imaging of exoplanets.Figure 1: The Interstellar Medium Science TargetsWe examine a solar thermal propulsion (STP) system to rapidly access the local interstellar medium via a solar perihelion burn. This approach uses several Venus and Earth gravity assists to fly out to Jupiter and then would dive towards the Sun. Approaching within 3 solar radii a perihelion burn would be performed, maximizing the spacecraft’s ΔV to achieve high Solar System escape velocities. A unique aspect of the STP mission concept is that the Sun is not only used as a gravity well for an Oberth maneuver, but also to heat the fuel to ultra-high temperatures (>3000 K), enabling a monopropellant burn with high specific impulse (Isp). An in-depth modeling exercise found this approach to be preliminarily feasible, with escape velocities of around 9 AU/yr achievable with current technology, and up to 16 AU/yr with significant future technological advances.While the baseline STP design is capable of providing just under 9±1 AU/yr, Figure 2 highlights areas of key technological improvements that could be explored. Ultimately, if all technological paths could be implemented, the overall performance as a best-case scenario could reach approximately 16 AU/yr. Figure 2 also qualitatively ranks these improvements from most likely to least likely when reading the graph bottom up. For example, implementing turbopumps in the system is likely more readily feasible than reducing the liner thickness in the near future. It is assumed that these upgrades can be implemented in the future without incurring any additional mass penalty over the baseline design. Thus, it predicts best case performance, and actual values would likely be lower. The improvements could be the result of a single point improvement, or a propagation of several developments.Figure 2: Overview of Solar Thermal Propulsion PerformanceAfter reviewing the STP approach, and comparing it to a solid rocket motor (SRM), it was found that with currently available technology, SRM outperforms STP with an escape velocity of approximately 12 Au/yr. However, future advances in heat exchanger lining materials, turbo pumps, and advanced heat exchanger geometries may enable solar thermal propulsion to provide higher escape velocities, which would provide one of the fastest ways to exit the solar system. Of particular importance is heating the hydrogen to 3,500K. Using a perihelion burn as a kick stage for a nuclear electric propulsion system was found to be particularly effective for achieving even higher escape velocities, up to 19.5 Au/yr.

This paper explores the critical role of deployable mechanisms in enabling spacecraft missions that demand large radio frequency (RF) aperture systems. Our focus centers on deployable antenna elements, specifically reflectors, booms, and feeds, for which JPL held responsibility in the last decade, addressing their impact on past, present, and future missions. The deployable systems are divided by maturity, starting with successful past deployments, upcoming missions nearing launch, projects in the funded phase without fabricated hardware, and pioneering technologies under development. This paper highlights the innovative accomplishments on deployable RF systems at JPL, from CubeSats to flagship missions. A key theme emerges from the study of these deployable systems: a constant decrease in stowed volume available, and a constant increase in frequency at which the antennas must work and antenna size. This increases the mechanical demands on stowing efficiency and deployment accuracy.

This paper introduces a design framework for elastically deformable hinges consisting of two parallel flat flexures, focusing on the case where the hinge is folded by 180 deg in the stowed configuration. The authors first consider architectures in which the flexures are staggered to avoid contact so that they can be modeled independently using Euler's elastica. They next focus on the nonstaggered case when folding can result in contact between the flexures, which they study through finite element simulations. The paper provides a set of design guidelines by rationalizing the relationship between hinge dimensions and the allowable curvature in the flexure material. For most hinge geometries applicable to deployable structures, the authors find a simple requirement for the minimum flexure length as a function of the allowable curvature in the material, L-h>10.7/kappa(m). Their analysis also provides insight into the reaction forces necessary to keep the hinge in the folded configuration, which is useful in determining the deployment dynamics and the constraints necessary to secure the panel in place during stowage. Experimental prototypes show good agreement with the numerical predictions.

In the past decade, CubeSats have undergone a revolution, moving from university research projects to enabling industry opportunities and government missions. In 2014, the Jet Propulsion Laboratory, California Institute of Technology (JPL/Caltech) initiated a research and technology development effort to advance CubeSat communication capabilities. One of the critical thrusts was the Ka-band parabolic deployable antenna (KaPDA). This antenna started with the ambitious goal of fitting a 42 dB, 0.5 m, 35 GHz antenna in a 1.5U canister. At that time, there had been minimal development in high-gain CubeSat antennas, critical for high-data-rate communications and remote sensing science. A Ka-band high-gain antenna would provide a 10,000 times increase in data communication rates over an X-band patch antenna and a 100 times increase over state-of-the-art S-band parabolic antennas. This paper discusses designing, building, integrating, and operating the flight antenna from a mechanical perspective, its final performance, and lessons learned. KaPDA enabled the RainCube mission, the first Earth Science CubeSat to have an active instrument. RainCube was launched in May 2018, making KaPDA the second deployable parabolic antenna to fly on a CubeSat and the first to operate in Ka-band, enabling follow-on opportunities for high-rate antenna communications and remote sensing science.

Elastically deformable hinges or deployable booms commonly follow architectures with face-skin layers only, cutting out the core to facilitate folding. This significantly reduces the shear stiffness of the hinge, which might reduce the first resonance frequency of the structure. We rationalize the free vibration of the system and the interplay between shearing and bending deformation by deriving an analytical formulation using Timoshenko’s beam theory. The framework is derived for a general beam and applied to a case of two flexures of arbitrary geometry with no core. Investigating the specific geometry of flat flexures reveals the existence of two nondimensional length ratios that capture the interplay between inertia, bending stiffness, and shear stiffness of the hinge. Our model explains the dependence of the frequency on the parameters of the system, as well as the dimensions that determine the transition between bending- and shear-dominated vibration modes. Comparison with finite element simulations shows that our analytic framework is able to predict the first natural frequency with less than 1% error. Additionally, an experimental validation was carried out using steel flexures and acrylic panels, showing good agreement with the predictions.

This paper presents Starburst, a novel architecture for deployable structures that has the potential to revolutionize the way rigid panel deployables are approached in future missions. The core elements of the Starburst architecture consist of rigid segments, high strain rods for rough actuation and guidance during deployment, a cable for pulling the system together and preloading, and kinematic mounts for precise location. Unlike traditional deployable structures that are constrained throughout deployment, Starburst uses an under-constrained approach that allows for greater flexibility in where segments can be stored on the spacecraft. The Solid Underconstrained Multi-Frequency (SUM) deployable antenna is an instantiation of the Starburst technology and would enable Earth science missions to better predict storm patterns. While early prototypes have shown initial feasibility, several key areas of future work remain, including sequencing multiple segment deployments, developing an assembly process for alignment, refining composite layup processes, and modeling thermal distortions. Overall, Starburst's adaptable and flexible design has the potential to enable a myriad of high accuracy deployables, not just for Earth science, but for missions from astrophysics to planetary science.

AbstractBecause of the miniaturization of small satellites, most of them have deployables to expand effective areas. However, Small Satellites are not only required to miniaturize systems, but often have a reduced budget, timeline, and employ teams with less experience. The goal of this paper is to provide a starting point for those new to deloyables, and working on small satellites, to understand the approaches available for deployable mechanisms and provide design practices which can improve success rates. To do so, this paper develops a framework for small satellite deployable structures, categorizing them into distinct deployment stages. It investigates the approaches that can be utilized for each stage, focusing on the stow, restrain, actuate, and locate stages. This review paper discusses the advantages and disadvantages of each approach, supported by examples provided in the references. It then highlights best practices for deployable mechanisms, and describes key challenges and future directions. By offering a comprehensive analysis of small deployable systems, this paper aims to guide engineers and researchers in implementing successful design practices for small satellite deployable structures.

A solution for wide bandwidth high frequency (up to 240GHz) deployable reflector is needed to enable a new type of low-cost and platform-versatile Earth Science missions. Deployable mesh reflectors are limited to a maximum frequency of 60GHz due to the mesh limitations (i.e. surface accuracy and reflectivity). Leveraging on a revolutionary deployable solution, the proposed antenna can operate at frequencies up to 240GHz with an overall surface root mean squared error of 31 microns. In this paper, we are assessing the feasibility of this reflector to operate at 238.5GHz. A Ka- and W-band dual-band deployable reflector is proposed achieving excellent efficiency of 69% and 63% at Ka- and W-band.

Two multi-beam Ku-band reflector antennas are designed for a NASA's Earth Science mission concept. This satellite could carry two antennas: a center-fed Cassegrain fixed reflector and an offset-fed deployable mesh reflector. Each support 7 near-nadir beams at Ku-band with matched patterns between the two antennas. The feeds are arranged to achieve a desired cross-track overlap at the 1.5 dB point of the antenna pattern. This solution takes advantage of the platform motion to achieve the desired scan pattern with a simple switched-horn solution. This paper describes the design approach of the two antenas to achieve identical beam pointing and beamwidth despite using different optical prescription.

This paper presents the fabrication and testing of a novel hinge design using high strain composite (HSC) flexures. Rather than more conventional curved shells such as tape springs, we use flat HSC sheets for the hinge. Although these flat flexures sacrifice stiffness, they offer the significant benefit of reducing the design size to the centimeter scale. We investigated fabrication techniques using unidirectional and plain-weave plies that result in minimum thickness for a required natural frequency when deployed. We observed that applying external pressure during the curing reduces surface defects of the HSC sheet. Next, we built several prototypes consisting of two panels and a hinge. We explored different fabrication techniques of co-curing the HSC sheets together and separately with panel bonding. All the prototypes survived folding tests and were able to deploy for at least several cycles. We observed cracks in one of the prototypes that we attribute to the misalignment of one of the panels, which in turn modifies the boundary conditions for the flexure. We experimentally characterized the natural frequency of the deployed panels using impact hammer tests. Our analytic predictions, which use calculated laminate stiffness, are within 3% accuracy for unidirectional plies. The plain weave laminates show significant deviation from the predictions, and we hope to experimentally estimate the stiffness to improve the estimation. We analyzed the deployment dynamics of the hinge and observed maximum accelerations in the order of 10g for the panel.

Herein is presented a design for a lightweight 5 m x 1 m radio-frequency (RF) reflector that can be flattened, folded, and coiled for compact stowage on board an ESPA-class spacecraft. The reflector is a reflectarray that operates at 3.2 GHz. Compared to state-of-the-art technologies for RF reflectors, the presented design has advantages in terms of areal density, stiffness, deployed stability, and scalability. Thermal and structural analysis is presented to demonstrate deployed stiffness and thermoelastic stability of the proposed design. Thermal analysis is used to predict in-space deployed temperatures in an operational condition, and structural finite element analysis is used to predict deployed vibration modes and frequencies, and the thermoelastic deformation of the deployed reflector. Also presented are the fabrication, assembly, and testing of two one-third-scale-length full-scale-width 1.7 m x 1 m test articles. These test articles are used to experimentally demonstrate RF functioning, stowage, deployment, and RF performance after deployment.

This paper describes Caltech's Space Solar Power Demonstration One (SSPD-1) payload and upcoming mission on Momentus Space Vigoride 5. SSPD-1 is comprised of three experiments each of which demonstrates the performance of a key technology piece in the space environment. We describe the goals of SSPD-1. The three experiments - Alba, DOLCE and MAPLE are discussed. The launch of SSPD-1 is scheduled for November 6, 2022 on Space X's Transporter 6 mission.

This paper presents a novel flexure design for small scale refelectarrays, specifically CubeSats, that requires high level of deployment precision. Advantages of using a HSC flexure includes additional thermal stability, lesser part count and passive deployment mechanism using the stored strain energy. The proposed design is to use two flat HSC strips of thickness ℎ, length 2𝐿ℎ, width 𝑊ℎ that attach to the core-reflector panel interface. The flexure satisfies two main requirements: ability to fold to a specific compaction 𝑑 without material failure and sufficient stiffness in the deployed configuration. The elastica analysis was implemented to identify the maximum strains in the flexure when folded and provide relations that combine Δ, ℎ, 𝐿ℎ with the maximum allowable strain. A simplified analysis was performed of the vibration of a panel-flexure combination using a two degree of freedom system. The vibration analysis reveals a dimensionless length parameter 𝛽 = (𝐿ℎ/𝐿𝑃) (Δ/ℎ) that determines the transition between bending and shear modes. A non-dimensional relationship was obtained between the natural frequency and 𝛽 that captures the impact of each parameter on the response of the structure. Our current work involves fabricating and testing of the proposed flexures using High Strain Composite materials, with plans to build a prototype for a deployable reflectarray.

A multi-beam Ka-band deployable mesh reflector designed for the NASA’s Earth Ventures Mission named INCUS (Investigation of Convective Updrafts) is presented. It is a 1.6m diameter mesh reflector offset-fed by an array of five small horns. The 4 off-nadir horns are arranged in two rows in order to achieve the desired cross-track overlap at the 3-dB point of the antenna pattern. This solution takes advantage of the platform motion to achieve the desired scan pattern with a simple switched-horn solution. The optimization of the antenna feed location to achieve the 3dB beamwidth overlap is described and the overall reflector antenna performance is presented.