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NASA NTRS2013

Accommodating Thickness in Origami-Based Deployable Arrays

The purpose of this work is to create deployment systems with a large ratio of stowed-to-deployed diameter. Deployment from a compact form to a final flat state can be achieved through origami-inspired folding of panels. There are many models capable of this motion when folded in a material with negligible thickness; however, when the application requires the folding of thick, rigid panels, attention must be paid to the effect of material thickness not only on the final folded state, but also during the folding motion (i.e., the panels must not be required to flex to attain the final folded form). The objective is to develop new methods for deployment from a compact folded form to a large circular array (or other final form). This paper describes a mathematical model for modifying the pattern to accommodate material thickness in the context of the design, modeling, and testing of a deployable system inspired by an origami six-sided flasher model. The model is demonstrated in hardware as a 1/20th scale prototype of a deployable solar array for space applications. The resulting prototype has a ratio of stowed-to-deployed diameter of 9.2 (or 1.25 m deployed outer diameter to 0.136 m stowed outer diameter).

flasher↗

NASA NTRS2019

Electric Sail Tether Deployment System for CubeSats

An Electric Sail (E-Sail) propulsion system consists of long, thin tethers - positively-charged wires extending radially and symmetrically outward from a spacecraft. Tethers must be biased using a high-voltage power supply to ensure that the solar wind produces thrust. While the E-Sail concept shows great promise for flying heliopause missions with higher characteristic acceleration than solar sails, there are significant technical challenges related to deploying and controlling multiple tethers. A typical full-scale design involves a hub and spoke arrangement of 10 to 100 tethers, each 20 km long. In the last 20 years, there have been multiple space mission failures due to tether deployment and control issues, and most configurations involved a single tether. This paper describes an effort to develop and test a simple yet robust single-tether deployment system for a two-6U CubeSat configuration. The project included the following: a) Tether dynamic modeling/simulation b) E-Sail single-tether prototype development and testing c) Space environmental effects testing to identify best materials for further development. These three areas of investigation were needed to provide technical rationale for an E-Sail flight demonstration mission that is expected to be proposed for the 2022 time frame. The project team used an “agile” engineering approach in which E-Sail single-tether prototype designs were iteratively developed and tested to solve problems and identify design improvements. The agile approach was ideal for this low Technology Readiness Level (TRL) project because tether deployer development involved many unknowns in prototype development that could only be discovered through iterative cycles of construction and testing. Extensive modeling and simulation were accomplished for three types of tether deployment: a) Stage 1: propulsive separation with one 6U fixed b) Stage 2: propulsive spin-up with one 6U fixed c) Stage 3: propulsive spin-up with both 6Us free Simulation results were valuable for understanding the propulsive and braking forces needed for controlled tether deployment. This paper describes the evolution, insights, and test/ performance data related to the resultant single-tether two-6U E-Sail test article which has been demonstrated in a test laboratory. The development effort suggests near-term work needed to achieve a useful flight demonstration, and provides ideas for how multiple-tether deployment systems might evolve going forward. A planned next-generation E-Sail prototype will include autonomous propulsive tether deployment while monitoring tether tension, location on the floor, distance between tether ends, acceleration, velocity, and propellant used.

Tinker, Michael↗

NASA NTRSDate not supplied

A Multifunctional Bistable Ultrathin Composite Boom for In-Space Monitoring of Deployment Dynamics

A novel concept of a multifunctional deployable ultrathin composite boom with flexible electronics and circuits integrated in a highly compact configuration, which can monitor its deployment dynamics in space, is presented in this paper. This concept is particularly suitable for CubeSats with extremely limited payload volume. The multifunctional deployable boom will be flight hardware installed in a 3U CubeSat that is scheduled to be launched to the International Space Station (ISS) in 2023 as an in-space technology demonstration. The multifunctional boom consists of a bistable self-deployable ultrathin composite boom, flexible electronics and circuitry for dynamics monitoring, and embedded flexible thin wires for power delivery and data transmission. The design, materials, and fabrication methods for the multifunctional boom are presented. Tests in simulated space environments showed the survivability and stability of the flexible electronics within a temperature range of 7 °C to 50 °C and under a high vacuum level of approximately 1 × 10−6 Torr. Ground-based vibration and deployment tests demonstrated the overall design of the multifunctional boom, data acquisition system, and deployment mechanism. Analyses of the data obtained from the integrated flexible electronics successfully captured the deployment dynamics and identified the natural frequencies of the boom in the range of 0–100 Hz. These results demonstrate that the concept is a promising approach to future multifunctional ultrathin deployable space structures.

Yao Yao↗

NASA NTRSDate not supplied

A Multifunctional Bistable Ultrathin Composite Boom for In-Space Monitoring of Deployment Dynamics

Yao Yao↗

NASA NTRS2022

SWOT and NISAR Boom Ground Deployment Test Challenges & Resolution

NASA’s Jet Propulsion Laboratory is developing two new spacecraft that use radar instruments to characterize temporal changes in the Earth’s surface with unprecedented precision (Figure 1). Both the Surface Water Ocean Topography (SWOT) and the NASA-ISRO Synthetic Aperture Radar (NISAR) spacecraft utilize large, precision flight deployable booms to properly position and support their instrument reflectors. The SWOT spacecraft includes two nearly identical reflector booms, each of which have similar flight deployable hinge designs. The NISAR spacecraft has a single reflector boom, with four unique hinge designs. These booms each undergo a multi-staged flight deployment sequence on orbit to transition from the launch stowed configuration to the science configuration within days of launch (Figure 2). The SWOT and NISAR Projects faced significant challenges relevant to requirement verification as well as hardware safety in their approach to ground testing these large flight deployables. This report summarizes flight deployable system design decisions that contributed to ground testing challenges. The report also summarizes the architecture trade study conducted for ground deployment testing. A summary of key issues encountered during flight deployable ground testing with the chosen common gravity offload system ensues, with discussion of the issues and mitigation measures implemented by both Projects that ultimately enabled successful flight subsystem-level full range of motion ground tests. Recommendations and lessons learned are offered to facilitate ground testability of future analogous large scale flight deployables.

Waters, Kyle C.↗

NASA NTRS2006

MARSIS antenna flight deployment anomaly and resolution

This paper summarizes the resolution of an in flight anomaly that occurred during the deployment of the first of three MARSIS antenna booms. Characteristics of this deployment are described, along with a correlation to finite element models and measured spacecraft inertias, which allowed the intermediate state of the boom to be accurately determined. Based on this information, a spacecraft maneuver was performed that warmed the stalled hinge and led to the first boom successfully locking into its designed geometry. The confirmed partially deployed boom shape was then used to develop a thermal model of the stalled hinge both in its initial solar attitude and during the successful spacecraft maneuver. Results from the hinge thermal model and component level testing were evaluated in order to determine the root cause of the anomaly and the probability of its recurrence on subsequent deployments. These conclusions were then utilized in planning mitigating actions that were implemented during the remaining two boom deployments. Final flight data are presented for both dipole booms indicating a correctly deployed and healthy antenna. The monopole boom deployment was detected but the final state of the boom is unknown.

Mars↗

NASA NTRS1972

Pioneer F/G appendage deployment

An integrated program of testing and analysis is described for developing and flight-qualifying the appendage deployment systems of Pioneer F/G, a spin-stabilized spacecraft which will fly by Jupiter. A combination of deployment system component tests, simplified nonspinning deployment tests, and analyses were used to develop the deployment system hardware. Analytical deployment models, verified by correlation with deployment test performance data and incorporating measured system parameters, were used to project in-flight deployment system performance over the expected range of environmental and hardware extremes.

Hesprich, G. V.↗

NASA NTRS1981

Deployment tests of a 36-element tetrahedral truss module

In the past, models of deployable structures were limited largely to small scale models which could be readily deployed by suspending the model on several soft shock cords. The scale of the deployable truss used in the present investigation precluded the use of this test technique as the gravity forces and moments are of the same order of magnitude as the deployment forces and moments. For these tests, the truss was deployed during free-fall in the LaRC 55' vacuum facility. Appreciably larger trusses could be deployed by lofting the packaged truss upward from the floor of the facility and allowing it to deploy during the upward as well as the downward portion of its trajectory, thus doubling the available test time. It must be realized, of course, that the mechanisms required to loft and decelerate such a large truss would be much more complex than those required for a straight drop.

Herr, R. W.↗

NASA NTRS1984

The STEP/STACBEAM experiment technology development for very large solar array deployers

The Stacking Triangular Articulated Compact Beam (STACBEAM) is discussed with reference to structural testing experiments afforded by ground simulation and the Space Technology Experiments Platform (STEP). The STACBEAM lends itself to a deployment technique which offers a radical improvement in flexible blanket solar array technology. A system for deployment and support of a solar array blanket is described which consists of the blanket, its containment structure, the support structure and its deployer, the blanket stiffening battens, and the deployable boom standoffs. In operation, the blanket is pulled out and supported by the STACBEAM which packages next to the folded blanket. Since the STACBEAM does not rotate during extension, complete control of the blanket is maintained during extension. Deployment of this system occurs one bay at a time in a sequential manner. The deployer provides sufficient rigidity so that beam stiffness is not degraded during the deployment process.

Samuels, R.↗

NASA NTRS2002

Deployment Simulation of Ultra-Lightweight Inflatable Structures

Dynamic deployment analyses of folded inflatable tubes are conducted to investigate modeling issues related to the deployment of solar sail booms. The analyses are necessary because ground tests include gravity effects and may poorly represent deployment in space. A control volume approach, available in the LS-DYNA nonlinear dynamic finite element code, and the ideal gas law are used to simulate the dynamic inflation deployment process. Three deployment issues are investigated for a tube packaged in a Z-fold configuration. The issues are the effect of the rate of inflation, the effect of residual air, and the effect of gravity. The results of the deployment analyses reveal that the time and amount of inflation gas required to achieve a full deployment are related to these issues.

Wang, John T.↗

NASA NTRS2009

Long Cable Deployments During Martian Touchdown: Lessons Learned

The launch of NASA/JPL's next generation Mars rover is planned for the fall of 2011. The landing scheme chosen for this rover represents a step forward in unmanned payload delivery. The rover will be lowered from a rocket powered descent stage and then placed onto the surface while hanging from three bridles. During this touchdown event, the communication between the rover and descent stage is maintained by an electrical umbilical cable which is deployed in parallel with the structural bridles. During the development of the deployment device for the electrical umbilical, many obstacles were identified and overcome. Many of these challenges were due in large part to the helical nature of the packing geometry of the umbilical cable. And although none of these issues resulted in the failure of the design, they increased both development and assembly time. Many of the issues and some of the benefits of a helical deployment were not immediately apparent during the trade studies carried out during the deployment selection process. Tests were conducted upon completion of the device in order to characterize both the deployment and separation characteristics of the cable. Extraction loads were needed for inputs to touchdown models and separation dynamics were required to assess cable-rover recontact risk. Understanding the pros and cons surrounding the deployment of a helically packed cable would most certainly influence the outcome of future trade studies surrounding the selection of cable deployment options.

Mars Science Lab↗

NASA NTRS2015

Deployment Technology of a Heliogyro Solar Sail for Long Duration Propulsion

Interplanetary, multi-mission, station-keeping capabilities will require that a spacecraft employ a highly efficient propulsion-navigation system. The majority of space propulsion systems are fuel-based and require the vehicle to carry and consume fuel as part of the mission. Once the fuel is consumed, the mission is set, thereby limiting the potential capability. Alternatively, a method that derives its acceleration and direction from solar photon pressure using a solar sail would eliminate the requirement of onboard fuel to meet mission objectives. MacNeal theorized that the heliogyro-configured solar sail architecture would be lighter, less complex, cheaper, and less risky to deploy a large sail area versus a masted sail. As sail size increases, the masted sail requires longer booms resulting in increased mass, and chaotic uncontrollable deployment. With a heliogyro, the sail membrane is stowed as a roll of thin film forming a blade when deployed that can extend up to kilometers. Thus, a benefit of using a heliogyro-configured solar sail propulsion technology is the mission scalability as compared to masted versions, which are size constrained. Studies have shown that interplanetary travel is achievable by the heliogyro solar sail concept. Heliogyro solar sail concept also enables multi-mission missions such as sample returns, and supply transportation from Earth to Mars as well as station-keeping missions to provide enhanced warning of solar storm. This paper describes deployment technology being developed at NASA Langley Research Center to deploy and control the center-of-mass/center-of-pressure using a twin bladed heliogyro solar sail 6-unit (6U) CubeSat. The 6U comprises 2x2U blade deployers and 2U for payload. The 2U blade deployers can be mounted to 6U or larger scaled systems to serve as a non-chemical in-space propulsion system. A single solar sail blade length is estimated to be 2.4 km with a total area from two blades of 720 m2; total allowable weight of a 6U CubeSat is approximately 8 kg. This makes the theoretical characteristic acceleration of approximately 0.75 mm/s2 at I AU (astronomical unit), when compared to IKAROS (0.005 mm/s2) and NanoSail-D (0.02 mm/s2).

Peerawan, Wiwattananon↗

NASA NTRSDate not supplied

The Adaptable, Deployable, Entry and Placement Technology (ADEPT) Enabling Advanced Entry, Descent, and Landing Capabilities for SmallSat Missions

The Adaptable, Deployable Entry and Placement Technology (ADEPT) is a mechanically deployable low-ballistic coefficient aeroshell entry system which performs entry, descent, and landing (EDL) functions for a broad range of planetary destinations including Mars, Venus, Titan and Earth. The deployable system allows mission planners to develop an aeroshell design that is stowed like a folded umbrella, and yet prior to the EDL mission segment, transforms into a deployed, near-rigid low ballistic coefficient configuration. The ADEPT architecture is scalable from a sub-1m class to 6m diameter entry vehicles for Venus and and it has been studied at up to 15m – 20m scale for human exploration class missions. There has been growing interest in performing interplanetary missions with small spacecraft as secondary payloads and ADEPT offers small satellite mission designers the ability to consider Aeroassist technologies to expand mission applications. The ADEPT architecture introduces a new approach for entry vehicle design. It deploys a flexible 3D woven carbon fabric as the primary drag surface. It is capable of surviving high aerothermal heating during entry, while maintaining a robust impermeable membrane that provides high drag for vehicle deceleration. The ADEPT project team has advanced this decelerator technology via systems-level testing at the one meter diameter (nano-ADEPT) scale. A subsonic aeroloads test (May 2015) and a pathfinder arc-jet aeroheating test (Sept 2015) and a successful sounding rocket flight (SR-1) launch of a 0.7 meter deployed diameter ADEPT (September 2018). The SR-1 flight experiment demonstrated most of the primary end-to-end mission stages including: launch in a stowed configuration, separation and deployment in zero-g, exo-atmospheric conditions, and passive ballistic re-entry of a 70 degree half-angle cone geometry. The ADEPT SR-1 sounding rocket flight experiment had to address challenges such as stowing within the constrained payload envelope, fabric folding, and packaging concerns.

Paul Wercinski↗

NASA NTRS2021

Characterizing the Aging of Alphasense NO2 Sensors in Long-term Field Deployments

Low-cost NO2 sensors have been widely deployed for atmospheric sampling. While their initial performance has been characterized, few studies have examined their long-term degradation. This study focused on the performance of Alphasense low-cost NO2 sensors (NO2-B42F and NO2-B43F) over 4 years (2016–2020). A total of 29 NO2 sensors from 10 batches were collocated 78 times at two sites with reference instruments. Raw signals from “functional” NO2 sensors correlated linearly with reference NO2 concentrations. After long-term deployment, sensor raw signals started to deviate from reference NO2 concentrations due to sensor aging, an accumulated effect after sensor unpacking. Several sensors eventually became “non-functional” as sensor raw signals showed no correlation with reference NO2 concentrations. Sensor aging and non-functionality may be primarily caused by expiration of the ozone (O3) scrubber built into these sensors so that sensors responded to both ambient NO2 and O3. The influence of O3 on sensor response is quantified through the permutation importance method. Most of the sensors are non-functional after approximately 200–400 days of deployment, and no sensor was functional after 400 days of deployment. This result agrees well with the estimated lifetime of the built-in ozone scrubbers considering the ambient ozone concentration in the Pittsburgh area where these sensors were deployed. To ensure reliable data quality in long-term field deployments, we recommend collocating NO2 sensors with reference instruments regularly after 200–400 days of deployment to identify and replace non-functional sensors in a timely manner.

NO2sensors↗

NASA NTRS2020

A Concept for the Deployment of a Large Lunar Crater Radio Telescope Using Teams of Tethered Robots

Kilometer-scale craters on the far side of the Moon have unique potential as future locations for large radio telescopes, which can observe the universe at wavelengths and frequencies (> 10 m, < 30 MHz) not possible with conventional Earth or orbital-based approaches. Distinct advantages of building a Lunar Crater Radio Telescope (LCRT) on the far side include i) isolation from radio noise due to the Earth’s ionosphere, orbiting satellites, and the Sun, ii) days of uninterrupted dark/cold sky viewing during lunar night, and iii) terrain geometry naturally suited for constructing the largest mesh antenna structure in the Solar System. A key challenge to constructing LCRT on the Moon is related to the complexity of deploying a 1-km diameter antenna and hanging receiver within a lunar crater whose diameter, depth, and slope are 3-5 km, 1 km, and ~30 degrees respectively. In this paper, we first evaluate the trade space for deploying a large, complex structure within a crater, and then provide a more detailed concept evaluation of our favored approach, which employs coordinated teams of tethered rovers to extract and suspend a folded antenna from a lander at the base of a crater. NASA’s Jet Propulsion Laboratory in collaboration with California Institute of Technology (Caltech) have developed a novel robotic system for accessing extremely steep terrains; the Axel rover is a two-wheeled rugged terrain vehicle that is supported by an electro-mechanical tether that provides power, data, and tensile support from a top-side anchor location. Recently, a pair of Axel robots have been used in a DuAxel configuration that allows for four-wheel driving and repeated passive anchoring at different locations. The DuAxel system has unique advantages for deploying an LCRT antenna, including the ability to deploy from a lander near a crater, drive a distance to the crater rim to deploy an Axel, and later, retract the deployed Axel in order to sequentially lift up sections of the antenna. Our proposed concept involves delivering a packaged antenna and receiver to the bottom-center of a crater floor on a lander, then later sending a team of multiple DuAxel rovers to retrieve guide wires from the lander, which are pulled to the top of the crater. We explore this concept in detail and provide some initial quantitative analysis to demonstrate the feasibility of our system with respect to the spatial and mass properties of the antenna as juxtaposed to DuAxel capabilities. Finally, we outline next steps towards validating our concept on the way to a future lunar deployment opportunity.

Hallinan, Greg↗

NASA NTRS2021

A Survey of CubeSat Deployable Structures: The First Decade

In the past decade CubeSats have made their way into the spotlight. They have evolved from small, university educational opportunities, to industry and governments using them make new discoveries and monetize space. However, with the small, constrained CubeSat form factor; there is often a need to expand the CubeSat through deployable mechanisms once the satellite is in space. This paper is a survey of deployable structures and their actuating mechanisms for CubeSats. The goal of this paper is to provide the applications within which deployable structures have been used in the past for CubeSats, the mechanisms with regards to how they deploy, the lessons learned, and limitations of the various types of deployables. The inputs to this paper come from a relational database in development to track launched CubeSat missions with deployable structures. From this database we can find insightful trends. This paper specifically focuses on the first decade of CubeSat deployables, from 2000 to 2010.

Arya, Manan↗

NASA NTRS2022

Detection and Mitigation of Transient Instabilities in Deployable Booms

Deployable space structures offer versatile uses in many space applications, as they can create very large space structures from a single space launch. However, they usually present important dynamical challenges during the critical deployment phase. In this paper, we focus on the dynamical behavior of deployable booms and the important challenge of avoiding collisions between boom’s structural elements or with other parts of the spacecraft, that can potentially have serious consequences. We define this kind of collisions as “transient instabilities” in deployable booms. A new paradigm to tackle these issues is presented in this paper, using techniques from multi-agent/swarm literature like transient instability prevention methods that avoid inter-agent collisions. We have adapted these techniques to deployable booms to guarantee safe deployment of booms and avoid collisions. We also present computational simulations of boom deployment in order to validate our techniques and demonstrate its use for future space missions.

Balas, Mark↗

NASA NTRS2024

International Space Station Satellite Deployment: Jettison Policy and Best Practices for Satellite Payload Developers

The International Space Station (ISS) deploys dozens of small satellites into Low Earth Orbit (LEO) each year. This presentation and associated paper cover the ISS Jettison Policy requirements and review/approval process, as well as best practices for satellite Payload Developers who have satellites manifested for deployment from ISS. Specifically, topics will include ISS Jettison Policy requirements to limit generation of orbital debris, limit risk of collision with ISS, and limit risk of collision with ISS visiting vehicles. The paper will include details on the ISS Program jettison candidate analysis and approval process, timelines for data submittal to ISS Program, resources for small satellite developers, and design & analysis recommendations for small satellite developers to maximize their likelihood of successful deployment from ISS. New Station deploy capabilities and ways the ISS Program addresses and facilitates innovations in small satellite technology, including propulsion systems, deorbit devices, constellation development, and novel tech demos, will also be explored. The 2024 session topic that best fits this abstract is Orbital Debris, SSA & STM. The ISS Jettison Policy intends to quantify and control the risks of deploying and operating small satellites, not only to ensure the safety of the humans flying in space, but also to preserve the orbital environment for world space activities and enable the significant benefits brought by such utilization. The ISS Program is committed to working with smallsat providers to address their challenges and enable safe, accessible, innovative missions. The Policy has grown with the industry, with each deploy yielding hard-earned lessons learned that improve our process – not only for the next deploy campaign, but with applicability and adaptability for future applications in LEO and beyond.

Jettison Policy↗