Conference PaperPDF Available

Advanced Deployable Structural Systems for Small Satellites

  • NASA Langley Research Center, Virginia, United States

Abstract and Figures

One of the key challenges for small satellites is packaging and reliable deployment of structural booms and arrays used for power, communication, and scientific instruments. The lack of reliable and efficient boom and membrane deployment concepts for small satellites is addressed in this work through a collaborative project between NASA and DLR. The paper provides a state of the art overview on existing spacecraft deployable appendages, the special requirements for small satellites, and initial concepts for deployable booms and arrays needed for various small satellite applications. The goal is to enhance deployable boom predictability and ground testability, develop designs that are tolerant of manufacturing imperfections, and incorporate simple and reliable deployment systems.
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Advanced Deployable Structural Systems for Small Satellites
W. Keith Belvin1, Marco Straubel2, W. Keats Wilkie1,
Martin E. Zander3, Juan M. Fernandez1, and Martin F. Hillebrandt2
(1) NASA Langley Research Center, Hampton, VA, USA
(2) DLR Institute of Composite Structures and Adaptive Systems, Braunschweig, GERMANY
(3)Institute of Adaptronics and Function Integration, Technische Universität Braunschweig, GERMANY
One of the key challenges for small satellites is packaging and reliable deployment of structural booms and
arrays used for power, communication, and scientific instruments. The lack of reliable and efficient boom and
membrane deployment concepts for small satellites is addressed in this work through a collaborative project
between NASA and DLR. The paper provides a state of the art overview on existing spacecraft deployable
appendages, the special requirements for small satellites, and initial concepts for deployable booms and
arrays needed for various small satellite applications. The goal is to enhance deployable boom predictability
and ground testability, develop designs that are tolerant of manufacturing imperfections, and incorporate
simple and reliable deployment systems.
Over the past decade, the size of spacecraft components and subsystems has been significantly reduced. As a
result, small satellites of the future may provide comparable or even better performance than much larger
satellites of the past. To increase the functionality of small satellites, novel methods for packaging and
deployment of structural booms and arrays are needed.
State of the art spacecraft boom concepts are mainly designed for large satellites with large appendages.
Scaling existing boom designs for small satellite applications leads to manufacturing issues due to limited
packaging volume. Currently, there is a lack of reliable and efficient boom and membrane deployment
concepts for small satellites. In particular, booms with high structural performance and reliable deployment in
the 5-20 meter length class are not available.
To develop small satellite boom and array concepts, NASA and DLR began a joint project in 2016 to develop
advanced deployable structural systems for small satellites. The project focuses on deployable booms and
deployment mechanisms for small satellite applications such as solar arrays, solar sails, drag sails and
instrument booms.
The paper provides a state of the art overview on existing spacecraft deployable appendages, the special
requirements for small satellites, and initial concepts for deployable booms and arrays needed for various
small satellite applications. The goal is to enhance deployable boom predictability and ground testability,
develop designs that are tolerant of manufacturing imperfections, and incorporate simple and reliable
deployment systems.
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Due to the miniaturization of spacecraft avionics, small satellites can now provide many of the services that
larger satellites performed in the past. Consequently, mission applications previously achievable only by large
satellites became target applications for small satellites. While the avionics and instruments have decreased in
size, some satellite components that require collecting or reflecting photons and electromagnetic (EM)
radiation must necessarily remain large, for example, photovoltaic power arrays, communication antennas,
and solar sails. Thus small satellites need booms and arrays that package in small volumes and miniaturized
deployment subsystems to enable system engineers to design more functional small satellites.
This section describes some applications for deployable booms. A key observation is that the dimensions of
deployable structures are not necessarily scaling with the dimensions of their host satellites! Thus large (5-20
m) lightweight deployable booms are critically needed for small satellites. These booms must not only
provide stiffness and thermal stability, they must also be reliably deployed after being stowed for many years.
2.1 Solar Arrays
Solar Arrays generate the power to run all electric subsystems of the hosting satellite. The power generated
depends on the efficiency of the photovoltaic cells and the total array area. Consequently, the size of the array
depends on the power demand of the hosting satellite, which will vary with different missions and not
necessarily with the satellite size.
Classical solar arrays are based on hinged panels for “accordion type” deployment. Small spacecraft require
lighter arrays with higher packaging efficiency. Thus flexible substrates and deployable booms are being used
to lightweight solar arrays for small spacecraft.
2.2 Communication Antennas
Communication antennas are used to communicate with Earth ground stations or with other satellites.
Communication with high data rates and low power consumption requires large reflectors. High data rate
communication antennas are usually realized by parabolic reflector antennas (see Fig. 1) and in some cases by
flat array antennas (see Fig. 2).
Fig. 1. Photo of the SRAL EM parabolic radar
antenna of the European Sentinel 3 satellite (image
credit: TAS, ESA)
Fig. 2. Canadian RadarSat2 with two symmetric
solar arrays (deployed sideways) and one large
deployed radar array antenna (downward facing),
(image credit: CSA)
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Phased array antennas steer the beam electronically by defining dedicated phase shifts for each array element.
It is thereby able to point the antenna on a target without a physically reorientation of the antenna structure.
The dimension of the antenna and the required surface accuracy are related to the operating frequency band.
For a given communication data rate, higher frequencies can utilize smaller antennas, but high frequency
antennas require higher surface accuracy due to shorter wavelengths.
2.3 Radar Antennas
Radar antennas are in general very similar to communication antennas. The significant difference is the usage
of this antenna. Radar antennas are usually scanning the ground below or adjacent to the ground path.
Therefore, the antenna does not require pointing to one spot (like a ground station) and can be attached
directly to the satellite structure without steering mechanisms. The required size of the antenna depends on the
frequency band (higher frequencies allow smaller antennas) but also on the desired spatial resolution of the
acquired images (finer resolution requires larger antennas).
2.4 Drag Sails
Drag sails (see Fig. 3) are passive deorbiting devices with large deployable surfaces that use the residual
atmosphere in low earth orbit (LEO) to decelerate a satellite after it has finished its mission. Due to the
constant deceleration, the orbit of the satellite is lowered followed by atmospheric re-entry and burn-up.
The drag sail is deployed after the primary mission is completed (~10-15 years), which results in a very long
stowed lifetime of the drag sail. This results in challenges for materials selection, mechanisms design, and
Fig. 3. Deployment sequence of Deorbit Sail CubeSat, launched 2015, deployment motor
failure (project under lead of University of Surrey, GB. Booms and boom deployer by DLR)
2.5 Solar sails
Solar Sails are propellantless propulsion systems as shown in Fig. 4 that generate thrust by reflecting solar
photons. Therefore, they require large but very light reflective surfaces that are usually provided by very thin
metallically coated membranes. Every square meter of sail area increases the solar pressure induced force and
every additional kilogram limits the acceleration that results from this force. Hence, extreme lightweight
design is crucial for solar sails. The loads on solar sails during mission are limited to solar pressure, maneuver
and thermal loads.
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Fig. 4. GOSSAMER-1 Solar Sail developed by DLR
2.6 Instrument booms
Instrument booms are used to deploy sensors from a satellite in order to limit the influence of the inherent
satellite magnetic field on the sensor data (see Fig. 5). Another application is a boom that elevates the position
of a lander camera to oversee its landing site or the operation area of a rover (like in Fig. 6.). Such booms can
also be used to offset a camera from a host satellite to allow the supervision of more complex deployment
Instrument booms support a single tip mass during deployment and in the deployed state. Primary design
drivers are maneuver and thermal loads (in-space) and in the case of surface systems, gravity loads as well.
Fig. 6. Deployable stereo camera systems of
Phoenix mars lander, (image credit: Lockheed
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Deployable mast concepts considered here are scalable to specific masses lower than 0.5 kg/m and have a
bending stiffness EI greater than 1500 Nm². These values are chosen to comply with the needs of lightweight
structures and to narrow the considered booms concepts to a reasonable number. The following overview is
just a brief introduction into deployable space booms. More detailed descriptions can be found in [1], [2] and
3.1 Storable Extendible Tubular Member (STEM) booms
Astro Aerospace’s Storable Extendible Tubular Member (STEM) is a deployable structure (see Fig. 7) that
has flown since the 1960s. The STEM is piece of steel or other material that rolls up flat on a drum and that
returns to its circular shape on deployment via motor command. It is capable of pushing or pulling and is
useful as a deployment structure for other booms. The stowed STEM fits into a small space and can extend
many times its stowed length.
A derivative is called the Bi-STEM [4], which can also be used as a deployable and retractable telescoping
mast. The Bi-STEM has very high accurate positioning characteristics. A telescoping mast can employ a Bi-
STEM (a two-piece STEM) boom as an actuator and stabilizer, which alleviates the need for the deployed
telescoping mast segments to overlap. Due to this feature and because the segments can be fully overlapped
when stowed, the mast enables an unusually lightweight and compact launch configuration.
Fig. 7: Variations of the STEM boom concept [Source: NASA]
3.2 Collapsible Tube Mast (CTM) Booms
Sener developed the collapsible tube mast (CTM) [5] which is a boom originally made of a Beryllium Copper
alloy that can be rolled and deployed from a canister as shown in Fig. 8. It consists of two thin flexible shells
with a biconvex shape bonded at their edges by resistance seam welding. The boom can be flattened and
coiled up on a cylindrical core that enables very long masts to be stowed within a very small volume because
it is almost independent from the mast’s length [6].
Beginning in 1998, DLR developed different versions of carbon fiber reinforced polymer (CFRP) CTMs of
different scales and for various applications (see Fig. 9). In preparation for a 20m X 20M solar sail mission
called Gossamer-2, DLR designed and built 14 m CFRP booms and tested them in simulated space and launch
environments. Composite laminate CTM booms are the leading boom concept for the collaborative
NASA/DLR project.
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Fig. 8: Deployment module of the CTM-Mast
[Source: Sener]
Fig. 9: DLR composite booms of different scale [Source: DLR]
3.3 Truss Booms
Truss booms provide higher strength and stiffness than tubular booms; however, these systems are much
larger in size and typically require a canister for packaging that exceeds the volume allowed on small
satellites. Nevertheless, these truss booms have proven to be highly reliable and efficient for many space
missions over the past few decades.
ATK’s CoilABLE masts, Fig. 10, have continuous carbon fiber longerons that can be elastically coiled for
stowage. For deployment there are two possibilities: a motorized stowage canister which enables deployment
without tip rotation; or the stowed package is located at the tip and will rotate during deployment.
The canister guarantees that at least one fully deployed mast segment is attached to the spacecraft so that force
transmission is given even in the beginning of deployment phase (for the tip rotation deployment a canister is
needed as well but with a shorter length). The stowage length for the CoilABLE mast is between 0.5% and 2%
of the deployed length but the canister size must be taken into account. The CoilABLE masts were developed
for high ratios of bending stiffness to weight and are flight proven [1].
Fig. 10: CoilABLE-Boom during deployment (left) and design sketch (right)
[Source: ATK/ABLE Engineering]
ATK’s Folding Articulated Square Truss (FAST) and the Able Deployable Articulated Mast (ADAM) have
much higher strength and stiffness then the CoilABLE boom. However, the higher stiffness results from larger
diameters and stiffer components that result in a weight above 1 kg/m.
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The ADAM-Mast is rectangular truss developed for applications where a high bending strength and a high
geometrical accuracy is required. It is folded in a similar way to the CoilABLE-Boom but without
deformation of longerons or battens. Instead each joint is equipped with hinges to connect the longerons. Fig.
11 shows the process of deployment. The ADAM-Mast is flight proven and was used for example for the
SRTM-Mission (Shuttle Radar Topography Mission; STS-99) [NASA Shuttle Radar Topography Mission].
Eight FAST-Masts (see Fig. 12) are currently used on the International Space Stations (ISS) to support the
solar arrays. The masts are 32m long, have a diameter of 1.09 m [7] and are deployed out of 2.32 m long
containers. Fig. 13 shows the deployment method. Every second batten ring is flexible. Therefore, the battens
are able to buckle and two cells of the boom can be collapsed. The deployment is driven by the strain energy
of the buckled battens and is controlled by a retaining mechanism in the container.
Fig. 11: Beginning
deployment of
ADAMmast (Source:
Fig. 12: Deployed FAST mast as
used on ISS (Courtesy of Able
Engineering Extracted from [7])
Fig. 13: FAST mast deployment
Principle(Extracted from [7] )
The AstroMast developed by Northrop-Grumman is based on a similar concept as ATK’s CoilABLE booms.
There are only minor differences in truss-architecture, for example the mast is twisted by 120° in the deployed
configuration (see Fig. 14). The Solar Sail AstroMast is an untwisted AstroMast with a small sail at its tip and
is used to counter disturbance torques from unbalanced solar pressure on satellites .
Fig. 14: AstroMast in stowed (left) and deployed configuration (right) [8]
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3.4 Telescopic Mast
Another type of mast from Northrop-Grumman is the Telescopic Mast (see Fig. 15). An internal mounted
STEM-boom drives the deployment and enables retraction [9]. The Telescopic Mast locks in place after
deployment and can be provided with a retractability feature. This mast can be sized for various stiffnesses
and strengths, based on specific needs. The materials suitable for the mast include aluminum, steel, fiberglass,
graphite/epoxy and carbon-carbon.
3.5 Triangular Rollable and Collapsible (TRAC) Booms
The Air Force Research Laboratory developed the TRAC boom [10] as shown in Fig. 16. This boom is made
of two curved C-shaped section joined along one edge. The TRAC boom is attractive to the CubeSat
community due to its extremely efficient packaging scheme. The boom can be constructed from stainless
steel or carbon composite laminates. The deployed structural performance of these booms is described in
Fig. 15: Telescopic Mast [9]
Fig. 16: TRAC Boom [10]
Continued development of CTM booms (see section 3.2) using thin ply CFRP composite laminates has been
undertaken at DLR and NASA.
DLR developed the booms for the CubeSat DeorbitSail launched in the summer of 2015. In this project, led
by Surrey Space Center, U.K., DLR designed, built and qualified new CFRP booms and its deployment
system. For this application the booms and mechanism were designed, as depicted in Fig. 17a), to demonstrate
rapid deorbiting using a 4 m x 4 m squared drag sail suitable for small size satellites [11]. Several models have
been built for testing. Fig. 17b) shows a full size qualification model displayed at the DLR Space Structure
Lab@ Uni in Braunschweig. Typical qualification testing [12] [13] is shown in Fig. 17c). Despite the intense
qualification and successful launch of DeorbitSail, the deployment of the booms and sails could not be
realized on-orbit due to software and electronic issues on the spacecraft leaving the mission uncompleted.
With the continued advancement in high strain, thin-ply composite laminates, NASA investigated the use of
thin shell composite booms for a Near-Earth Asteroid (NEA) Scout mission [14]. The NEA Scout is a 6U
CubeSat being developed for a robotic reconnaissance mission that will be deployed to fly by and return data
from an asteroid representative of NEAs that may one day be human destinations. The booms and deployment
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STO-MP-AVT-257 26 - 9
mechanism passed all thermal-vacuum testing followed by post-test deployment. Based on the common
interests in this class of small satellite boom technology, NASA and DLR will collaborate on developing
scalable deployable boom technology for future space missions.
Fig. 17. DeorbitSail deployable structural subsystems: a) DLR developed small CFRP Booms & Deployment
module, b) full size Qualification model of DeorbitSail at the DLR Space Structure Lab@ Uni, in Braunschweig, c)
testing and qualifying CFRP booms and deployment module in DLR facilities (Source: DLR)
Current research and development at DLR [15-18] involve: 1) studies of robustness and tolerance of complete
structural Gossamer space systems and subsystems; 2) the development and flight qualification of boom, sail
and deployment technology for commercial drag-sail applications used as sub-systems on 600-1500 kg class
satellites in LEO; and 3) studies on the design of ultra-light weight booms and truss technology for large solar
arrays. A large focus at DLR is put on the strategic and important development of new boom technology
exceeding current designs and opening new applications as those of PV arrays or membrane antennas for
small to medium sized satellites, using the synergy of collaborations with other agencies like NASA.
NASA research and development is focused on low cost, small volume, high structural performance booms
that are both predictable and reliably deployed. Specifically, the technical challenges to be addressed by the
joint NASA/DLR project include:
Scalability of boom fabrication methods to tens of meters.
Packaging in small volumes without failure (damage initiation and propagation).
Behavior of materials at very high strain states (beyond ASTM testing standards).
Use of new thin-ply materials with low micromechanical behavior knowledge (braids, spread-tow
fabrics, CNT, hybrid laminates).
Material (radiation tolerance, permeability, moisture absorption, micro-cracking).
Creep/stress relaxation during long-term stowage (boom shape, deployment force and dynamics
Durability of bonded/stitched sections.
Long-term aging in space environment.
Achievement of bi-stability effects in booms other than cylindrical shaped.
Numerical models for large systems that cannot be fully tested on the ground.
Testing and characterization of lightweight structures under a 1g environment.
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The collaborative project will target 5-20 meter deployable booms with the following design requirements for
small satellite applications:
o The stowed structure shall have a first Eigen frequency of 100Hz or higher.
o The deployed structure shall have a first Eigen frequency of 0.05Hz or higher.
o The stowed structure will withstand the static and dynamic loads.
o The deployed structure will withstand all expected thermo mechanical loads.
Thermal Stability
o The structure shall be designed in a way that it fulfills criteria on pointing accuracy and
surface quality under all mission relevant environments.
o The structural concepts shall be scalable or modular in order to react to changing demands
for surface area.
o Processes that are cost efficient, repeatable, and with known tolerances.
The three year collaborative project will conclude with testing of one or more advanced deployable boom
concept(s) (ground and parabolic flights) to achieve a technology readiness level (TRL) of 6. The following
tests are anticipated:
Boom stiffness and strength characterization.
1-g functional deployment tests with optional gravity compensation.
Thermal vacuum and vibration tests of the stowed system .
Partial and full deployment tests.
0-g functional deployment tests during parabolic flight campaign (see Fig. 18).
Fig. 18. Exemplary frame of a DLR parabolic flight campaign from 2009 used to verify different
deployment control principles for DLR composite booms [19]
Advanced Deployable Structural Systems for Small Satellites
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Small satellites of the future may provide comparable or even better performance than much larger satellites
of the past. One of the key challenges for small satellites is packaging and reliable deployment of structural
booms and arrays used for power, communication, and scientific instruments. The lack of reliable and
efficient boom and membrane deployment concepts for small satellites is addressed in this work through a
collaborative project between NASA and DLR.
The state of the art and recent experience in deployable booms for small satellites shows a technology gap for
booms 5-20 meters in length that can be stowed in small volumes (for example, 3U CubeSats). The current
project will advance at least one boom concept to TRL 6 by 2019 to meet the following goals:
Satisfy unique requirements of Small Satellites (volume, mass, and power).
Maximize ground testability.
Permit the use of low-cost manufacturing processes.
Be scalable for use as elements of hierarchical structures (for example, trusses).
Have high deployment reliability.
Have controlled deployment behavior and predictable deployed dynamics.
Tubular thin walled shell composite booms offer the promise of high specific strength and stiffness and very
good thermal stability. Initial designs are underway and hardware prototypes are in development. Static and
dynamic ground tests will be used to verify the boom strength and stiffness. Detailed modeling and simulation
will be used for design and to characterize nonlinear thermal/mechanical response. Zero gravity tests will be
used to verify deployment forces and reliability of the boom and deployment mechanisms.
The end goal of the project is to fill the technology gap for small satellite deployable booms with technology
that is predictable, testable, and reliable.
C. H. Jenkins, Recent Advances in Gossamer Spacecraft, Vol. 212, Progress in Astronautics and
Aeronautics Series, 212, AIAA, 2006, p. 344.
C. Sickinger, Verifikation entfaltbarer Composite-Booms f¨ur Gossamer-Raumfahrtsysteme,
Dissertation, Technische Universit¨at Carolo-Wilhemina zu Braunschweig, Mar 2009, Publisher: Shaker,
ISBN: 978-3-8322-8049-9.
M. Straubel, Design and Sizing Method for Deployable Space Antennas, Dissertation, November 2012,
M. W. Thomson, Deployable and Retractable Telescoping Tubular Structure Development, AIAA, AHS
and ASEE, Aerospace Design Conference, Irvine, CA, 1993.
F. Del Campo and J. I. R. Urien, Collapsible Tube Mast Technology Demonstration Program, Space
Technology-Industrial and Commercial Application, pp. 61-76, Jan. 1993.
ESA/ESTEC, CTM Technology Demonstration Programme Bridging Phase Final Report, 1991.
G. Tibert, Deployable Tensegrity Structures for Space Application, Stockholm, Sweden, 2003. Technical
Report 2002:04 ISSN 0348-467X.
Advanced Deployable Structural Systems for Small Satellites
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N. Grumman, AstroMast Data-Sheets DS307 and DS407, 2010.
N. Grumman, Telescopic Mast Data-Sheet DS401, 2010.
J. Banik, and T. Murphey, Performance Validation of the Triangular Rollable and Collapsible Mast, 24th
AIAA/USU Conference on Small Satellites, Logan, UT, August 2010.
M. Hillebrandt, S. Meyer, M. E. Zander, M. Straubel and C. Hühne, The Boom Design of the DE-ORBIT
Sail Satellite, Braunschweig, Germany, ESA/DLR/CNES, 2014.
M. Hillebrandt, S. Meyer, M. E. Zander, C. Hühne and M. Sinapius, Mechanical Characterization of
Deployable Thin Shell CFRP Booms for CubeSat DE-ORBIT Sail, Braunschweig, Germany,
M. E. Zander, M. Hillebrandt, M. Sinapius and C. Hühne, Mechanical Characterization of Deployable
Thin Shell CFRP Booms for the CubeSat DE-ORBIT SAIL, Jerusalem, Israel, IAF International
Astronautical Federation, 2015.
J. M. Fernandez, Advanced Deployable Shell-based Composite Booms for Small Satellite Applications,
in 14th European Conference on Spacecraft Structures, Materials and Environmental Testing, Toulouse,
France, 2016
M. E. Zander, A. Wilken, M. Sinapius and C. Hühne, Mechnical Testing of Deployable Thin Shell CFRP
Booms in Ideal and Realistic Interfaces for the Solar Sail Demonstrator GOSSAMER-1, in 14th
European Conference on Spacecraft Structures, Materials and Environmental Testing, Toulouse, France,
M. Straubel, P. Seefeldt, P. Spietz and C. Huehne, The Design and Test of the GOSSAMER-1 Boom
Deployment Mechanisms Engineering Model, 2nd AIAA Spacecraft Structures Conference, Kissimmee,
FL, USA, 2015.
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Deployment Mechanisms for GOSSAMER-2, Jerusalem, Israel, IAF International Astronautical
Federation, 2015.
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Deployable Solar Arrays for Electric Propelled Spacecraft, in 14th European Conference on Spacecraft
Structures, Materials and Environmental Testing, Toulouse, France, 2016.
M. Straubel, M. Sinapius and S. Langlois, On-Ground Rigidised, Deployable Masts for Large Gossamer
Space Structures, in European Conference on Spacecraft Structures, Materials & Mechanical Testing,
Toulouse, France, 2009.
... Compared to large space systems the small systems offer fast development time, low cost, the possibility of large swarm missions and more scientific return on investment [23]. Up to date, the National Aeronautics and Space Administration (NASA) has developed lightweight deployable composite booms for small spacecraft missions [24,25]. Adorningly, one of the key challenges for small spacecrafts is the packaging and the reliable deployment of structural components [24]. ...
... Up to date, the National Aeronautics and Space Administration (NASA) has developed lightweight deployable composite booms for small spacecraft missions [24,25]. Adorningly, one of the key challenges for small spacecrafts is the packaging and the reliable deployment of structural components [24]. In the future, 4D printed smart structures will facilitate custom built, small scale, deployable space applications. ...
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... The first CTM ever used in space was jointly developed by NASA and ESA for the ULYSSES mission [2] . The development of CTLT is also considered as a key technique of German Aerospace Center DLR's solar sailing technology [3][4][5][6] , and in 2009 an agreement was made between ESA and DLR by which they started a three-step project aiming to develop, prove, and demonstrate that CTLT can serve as a safe and reliable component for long-lasting and deep space missions [7,8] . NASA has also recently expressed an interest in small CTLTs as a candidate solar sail boom for low-cost deep space exploration and science missions [9][10][11][12][13][14] . ...
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This paper describes challenges and lessons learned throughout the concept, design, assembling, integrating, and testing for hardware and software of Virginia Tech (VT) ThickSat, a testbed for lightweight deployable space structures designed by engineering students at Virginia Tech. The project started in 2017 as part of VT's senior design undergraduate team in collaboration with the Virginia Space and Near Space Launch Systems. The project's mission is to prove passive deployment of a spring table boom in low earth orbit, obtain deployment confirmation and transmit a picture back to Earth. To develop this project, over 25 different undergraduate and graduate students participated. In this process, they reached many breaking points and tough technical decisions. Throughout its development, the mission faced significant design reviews. A maximum allowed 100mA power draw from the bus and a top 150 kiloByte packet size transmission for its entire 28-hour mission. The resulting design can be replicated and easily scalable for much more significant roles under the exact requirements. This paper builds the challenges and lessons learned from redesigning, assembling, integrating, and testing hardware and software. Furthermore, it describes the restrictive design characteristics of the ThinSat program in detail, which led the students to come with intelligent solutions for its many different design interactions. The ThickSat solution includes a custom low-power PCB for an STM32 microcontroller, a servo actuated release mechanism, and a versatile chassis for the ThinSat Program. This study comprises an analytical point of view from the senior monitoring group and other engineers from the Center for Space Science and Engineering Research, known as (Space@VT), summarizing the experience from a student-led ThinSat project. The outcome of this paper is to share an experience that leads to bolster future SmallSat missions at Virginia Tech and other institutions.
... Telescoping booms also have a comparable form of a collapsible truss boom which incorporates elements from the flexible link boom to achieve higher stiffness and strength such as the deployable AstroMast seen in Figure 16. The truss booms are enclosed in a canister and are currently quite large when packaged, this makes it difficult to be implemented for small spacecraft [11]. ...
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This paper overviews the design and assembly process for a bench-test prototype model of a solar sail-based debris capture LEO orbiter. The purpose of this paper is to prove the successful operation of the CubeSat’s various subsystems in a 1G environment and to provide the requisite knowledge for future progression of this project. The core systems being designed and tested for this prototype are the boom deployment mechanism, debris capture mechanism, and the vision and electronics systems. Due to time and resource constraints, the deployment of the sail will be excluded from this iteration of the prototype. The end goal of this project is to provide a suitable foundation for future funding opportunities to further the TRL of the design through additional investigative research.
... Deployable truss structures have been used in space, mainly as masts, booms or beams, where the total deployed length is much greater than individual truss bays. For example, the Folding Articulated Square Truss Mast (FASTMast) (Fig. 5a and 5b) is used on the International Space Station (ISS) to support the solar arrays [11,12] and the Shuttle Radar Topography Mission (SRTM) used a 60-m long deployed FASTMast to support scientific instruments mapping the Earth's topography [13]. The FASTMast and SRTM both used coilable longeron booms. ...
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A trade study was conducted that evaluated viable concepts of operation for the packaging and deployment (P&D) of novel deployable modular truss modules, called TriTruss modules, that can be assembled to form a large aperture In-Space Assembled Telescopes (iSAT). In this first phase of an ongoing more comprehensive trade study, concepts were proposed and then evaluated based on initial metrics representing features of an ideal P&D concept. The ideal TriTruss P&D concept is defined as one that: allows for efficient packaging, has sufficient geometric versatility to be launch vehicle independent, provides a stiff and lightweight structure, has low mechanical complexity, and has component modularity. The P&D concept should allow for prelaunch subsystem or utility integration if required. The concept should be kinematically simple and be robotically deployed using a minimum number of specialized tools. The P&D concepts evaluated are categorized as: core collapse, face collapse, and erectable structures. Sub-scale models were constructed to help understand the kinematics and mechanical complexity required to enable P&D. Based on a weighting scale, the most promising candidate P&D concepts have been selected and will undergo more rigorous structural design, analysis, and testing in the study's next phase. The ultimate goal of the comprehensive trade study will be to recommend a single TriTruss design and associated P&D concept that will be built and evaluated at
... Since CFRP is rigid, lightweight and withstands the harsh conditions in space, such structures are broadly used in aerospace despite the higher material and development costs: Examples for CFRPs in astronautics are honeycomb CNT sandwich primary structures for satellites [1], thin-shell self-deploying satellite booms [2,3], antenna systems with extremely low thermal distortion [4], rigid light baffles for telescopes [5] or parts of thermal regulation systems [6]. Additionally, winded CFRP components are state-of-the-art to produce pressurized tanks, also for launch vehicles [7]. ...
Conference Paper
Fiber-reinforced materials offer a large improvement in structural performance if specific load cases can be determined. In aerospace, lightweight structures are crucial because of launcher limitations. For academic purpose CubeSats are a powerful concept to participate in space research on a low-cost-level. Reducing the structural mass, while keeping the mechanical performance, provides a bigger payload mass budget. Additionally, there are several types of payloads that do not fit in common structural components. Inasmuch as CubeSats are mainly used in research, such systems change substantially, so that an easily adaptive method would be beneficial to be no longer restricted by prefabricated structural components. The developed rapid prototyping technology tackles these issues by having an automated 3D deposi- tion method which can produce extremely lightweight as well as geometrical- and load-adaptive primary structures with minimum space requirements. A fiber deposition head for a six-axis robot has been devel- oped to impregnate and wind a single carbon roving on a frame to produce 3D integral components. The geometry of the frame can be adjusted to the required application by introducing holes or attachment points at nearly any position. Its modular layout varies, so that it only can be fabricated economically by fused deposition modeling and removed after resin curing easily. Furthermore, by blending additives into the resin it is possible to create material gradient components. Hence adaptiveness can be generated e.g. in terms of solar energy absorption. Compared with 1U aluminum wall structures available on market a carbon fiber-reinforced plastics (CFRP) winded structure results in a mass saving of 45% for a solid and 76% for a skeletonized wall segment, premised on calculations for two layers of CFRP made of 24K rovings with a fineness of 1600 tex at a fiber-volume-fraction of only 50%. This concludes that maximum potential can only arise with optimized fiber path generation. Costs can be saved in terms of material (no fiber blend), molding (3D printed frame), design of fiber path (sys- tematic guidelines), assembly (integral design) and manufacture (robotic production). The advantages will be demonstrated with a generic non-in-situ-sensory 1U CubeSat because it is easy to compare it to other systems due to the strict design specifications. The paper will include detailed information on the design of the robotic fiber deposition head, the modular and adjustable frame, the winding pattern generation and the mechanical testing.
... A complete overview of the state of the art in spacecraft deployable appendages can be found in Belvin, et al. (2016), where the author discuss about the challenge and limitations of the new deployable structures concepts and designs, thus justifying the recent studies and advances in the area. Similarly, Fetchko, et al. (2004) present a brief review of tdeployable booms, arguing that the technological advances in high performance materials are limited by the old-fashion structural boom designs. ...
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Abstract. In this work, we perform a simplified nonlinear dynamic analysis of an aerospace structure - a truss boom - to study the impact of our geometrically and material nonlinear model in its dynamic behavior. For this, we have implement a numerical integration step-by-step algorithm for the resolution of non-linear dynamic systems based on the Newmark-beta algorithm, where the resolution of dynamic equilibrium iterations in each time step are performed by Newton-Raphson method. The boom structure, similar with the ADAMmast©, is modelled as a 2D truss structure under the influence of a dissipative harmonic load to simulate the thruster effect in a hypothetical attitude control situation. Finally, we present a brief discussion of the program in the scope of bi-dimensional problems. Keywords: Nonlinear dynamics, Planar truss, Aerospace Structures, Deployable Structures.
This study proposes a concept of a single-layer deployable truss structure driven by elastic components which is applicable to small satellites. The structure is self-deployable from its stowed state to a planar regular hexagon configuration, and the concept is compared with other three competing concepts in terms of some geometric metrics. The deployability and serviceability of the model are verified through both numerical analyses and experiments. The deployment process of the structure is investigated and demonstrated by a deployment test. Robustness and stability analyses for the deployment are also conducted by considering failure of some elastic components. Flatness of the deployed structure is analyzed and measured. Modal analysis and frequency identification tests reveal that the fundamental frequency of the deployed structure is around 2.3 Hz. It is concluded that the proposed single-layer deployable truss structure is valid and has high potential application to small satellites.
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The long list of advantages that small satellites can offer in terms of fast development time, low cost, possibility of large swarm missions, scientific and/or educational return on investment are, in general, in conflict with the challenging mass and volume constraints imposed on the spacecraft subsystems. Many systems such as solar panels, antennas, sensors, telescopes, solar sails, drag sails, sun shades, radiators, and payloads are in need of reliable deployment of structural booms and arrays for power generation, communications, propulsion, deorbiting, thermal control and scientific instruments. The paper presents ongoing research and development of thin-shell rollable composite booms designed under the particularly stringent and challenging requirements of CubeSat-based solar sails. Several new boom concepts are proposed and other existing ones are improved upon using thin-ply composite materials to yield unprecedentedly compact deployable structures for a wide range of small satellite applications. Many laminates were investigated on short sections of each boom design, and the most promising ones were selected for the fabrication of longer booms. Inducing bi-stability on some of the booms was studied experimentally by choosing a particularly different laminate construction for each of the two thin-shell walls producing favorable results. For every boom introduced, the scalable fabrication process developed to keep the overall boom system cost down is shown. Finally, the initial test results carried out with purposely designed boom structural characterization test methods with gravity off-loading are presented to compare the structural performance of the booms under expected operational and general load cases.
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Gossamer space structures like solar sails, drag sails or solar arrays expand to very large dimensions while at the same time aiming for low masses, leading to a low areal density. In many cases these applications are designed using long lightweight booms, spanning out sails or other membranes. For launch and transfer however the booms need to be stowable and compact, while they are demanded to withstand the sail loads during deployment and in operation phase. In DLR´s Solar Sail demonstrator Gossamer-1, a 5 m x 5 m squared solar sail, four sail quadrants are spanned out by four thin shell CFRP booms. Facilitating a tip deployment, the booms are flattened and coiled each on a cylinder of a deployment unit for stowage, while the boom roots are fixed in an interface to the spacecraft bus in a cross like arrangement. Having its purpose in transferring occurring boom loads into the main structure, the interface needs to be flexible at the same time, mimicking the changing cross section of the boom, once it deploys and transforms from flat to its full cross sectional dimensions. Another interface connects the sails to the booms, while holding the boom´s cross section in a semi-deployed state. These found conditions are imperfect compared to using the full cross sectional shape of the boom with its full second moment of area and an ideally clamped root fixation. Simulating the boom and its interfaces of the Gossamer-1 demonstrator, several booms are tested in an advanced vertical boom test stand at the DLR Space Structures Lab @ Uni of DLR Braunschweig. Mechanical characteristics and properties of full scale booms with ideal boundary conditions and realistic interfaces are being investigated, enabling a direct comparison. Applying certain realistic load cases, like lateral bending, axial compression and combinations of both, thresholds and robustness as well as load carrying capabilities after buckling several times are quantified in practical tests. Thus characterizing the Gossamer-1 boom under ideal and realistic boundary conditions is realized, specifying the boom performance in reality and providing the base for a robust Gossamer Spacecraft design that facilitates thin shell CFRP booms. This paper gives a detailed insight on the test stand and setup, used interfaces and boundary conditions, testing procedures, and discusses the test results of both configurations on characteristic load-displacement curves and acquired values for buckling failure design.
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As the technology for solar electric propulsion has been further matured in the past years, the number of current and planned electric propelled spacecraft is rising. This technology can be applied to different mission scenarios such as: Deep space probes, communication satellites (transferred from a GTO or LEO to a GEO) or even reusable Moon or Mars supply modules. Although electric propulsion requires only a relatively small portion of fuel it does demand a lot of power. For deep space probes or Moon and Mars supply modules constant power supply with some hundred kW and up to 1MW are realistic. State-of-the art solar array concepts could be adapted to such dimensions but the question needs to be asked if these architectures are the most suited for very huge arrays. Moreover, the availability and constant maturing of flexible photovoltaic cells raises the question if membrane based, gossamer structures could be a more weight efficient alternative. Based on current DLR technologies for deployable space structures two concepts for large solar arrays are examined. The analysis is done for a power spectrum of 2:5kW to 500kW and two types of solar cell technologies: thin-film solar cells (TF) applied to a thin membrane substrate and state-of-the-art high performance multi-junction cells (MJ). The first concept is based on the DLR GOSSAMER solar sail architecture while the second is a flexible blanket design similar to the ISS. To evaluate conceptual design changes several sub-configurations are examined as well by use of parametric finite element models. The configurations are analysed regarding their performance parameters specific power (power per array mass), stowage volume efficiency (power per stowed volume) and structural mass ratio (structural, mechanism and residual mass divided by photovoltaic blanket mass). For all configurations a partially dramatic decrease in performance with increasing power and size is observed. While in the small to medium power region (2:5.. 100kW) good performance is achieved for both configurations, in the high power region (100.. 500kW) especially stowage volume efficiency and stowed dimensions become a critical factor. The impact of array size can be observed when comparing both cell technologies. To compensate for the loss in cell efficiency, TF arrays are significantly larger. While in the small power regions performance values to MJ arrays show small differences this diverges strongly with increasing array power and thereby size. In case of structural mass ratio a difference up to a factor of 3 and in stowage volume up to a factor of 5 is observed. Comparing the concepts among each other the flexible blanket design outperforms the solar sail structure. Typical values for a 100kW-array of the best performing sub-configurations using MJ cells are 184W=kg specific mass, structural mass ratio of 0:5 and stowage volume efficiency of 35kW=m3 in case of the solar sail configuration. For the flexible blanket design 273W=kg in specific mass, a structural mass ratio of 0:2 and a stowage volume efficiency of 110kW=m3 is achieved. Furthermore critical components, development needs and recommendations for further enhancement of array performance are identified and a break-even cell efficiency for the thin-film photovoltaic technology is calculated.
Conference Paper
DLR is currently developing several mechanisms for boom and membrane deploying spacecrafts. Two of the related projects are dedicated to solar sailing technology to investigate this promising alternative space propulsion technology. One supports the Gossamer-1 mission in which a 5 m by 5 m solar sail technology demonstrator is developed; the second is the planned, up scaled successor Gossamer-2 with a size of 20 m x 20 m. One goal of the development process in these missions is to achieve a deployment mechanism, that can be scaled up and used in future missions with larger sail areas (Gossamer-2: 20 m x 20 m, Gossamer-3: 50 m x 50 m) spanned out by larger booms. Within the EU funded DEPLOYTECH project the well matured Gossamer-1 boom deployment mechanism has been adapted for the Gossamer-2 size. Therefore, DLR scaled up and redesigned the deployment mechanism to work with the larger 14 m CFRP boom dedicated for the Gossamer-2 size spacecraft. In this paper the concept and design of the up-scaled boom deployment mechanism are described, followed by the main part that addresses the environmental testing performed on the prototype of the up-scaled boom deployment unit. The objectives, the test procedure and the results from thermal-vacuum, vibration and functional deployment tests are presented in detail. Furthermore, FE-Analyses for simulating the vibration tests are performed. In this regard the used FE-model is described and the obtained results are shown in detail. Consequently, the simulated and measured results are compared and discussed.
Conference Paper
DE-ORBIT SAIL is a cubesat based drag sail for the de-orbiting of satellites in a low earth orbit. It is scheduled for launch in late 2014 and will deploy a 25m² sail supported by deployable carbon fiber booms designed and manufactured by DLR. This boom possesses a closed cross-section formed by two omega-shaped half-shells. Due to this cross-sectional design the boom features a high torsional stiffness. Thereby a high bending strength is achieved compared to other boom concepts for similar applications as the boom is less sensitive to flexural torsional buckling. The boom concept selection is based on a detailed analysis of three types of deployable booms which differ in their cross-sectional design. From this analysis the double-omega boom was determined as most suited for DE-ORBIT SAIL. For the manufacturing of the booms a novel method is used where the booms are manufactured in an integral way in one piece.
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The risk for spacecraft in Low Earth orbit (LEO) to be hit or damaged is increasing due to a growing space debris population. Reduction and limitation of this risk is addressed by de-orbiting end-of-life maneuvers. To realize such a maneuver a drag augmentation device that increases the drag efficient surface of a satellite, acting with the residual atmosphere and accelerating the decay of orbit altitude until re-entry into Earth’s atmosphere, is necessary. Currently such a device is being flown with the CubeSat based drag sail satellite DeorbitSail by the University of Surrey, DLR, and other partners. The goal is to demonstrate the in-orbit deployment of a 4 m x 4 m drag sail, suitable for small and medium size satellites, as an end-of-life de-orbiting device. The device has a squared sail design utilizing 4 triangular membrane sail segments that are deployed and spanned out by thin shell CFRP (carbon fiber reinforced plastics) booms developed by DLR. The focus of this paper is on one of the main structures, the deployable thin shell CFRP-booms that are susceptible to buckling. Properties and characteristics of full scale booms regarding structural load capacities in all applicable load directions are being determined in a newly introduced vertical test stand. Thresholds and robustness for certain load cases like axial compression and lateral bending, and load carrying capabilities after buckling several times, are determined and quantified in practical testing, thus specifying the boom performance in reality. This is of high importance in order to be able to predict and design a robust drag sail structure, not failing due to overloading in space. This paper gives information of the applicable loading on the booms derived by the space application, the used test stand and equipment as well as the testing itself. Finally the acquired test results are analyzed and discussed.
Conference Paper
The main focus of this paper is the detailed introduction of the oom deployment concept implemented for the German eployable membrane technology demonstrator Gossamer-1. The technology aims for solar sailing, thin-film photovoltaic and drag augmentation as possible use cases. Therefore, the main functional and geometrical requirements for the mechanisms are derived from the mission design and the global sail deployment concept which are both introduced in detail. The regarding mechanism design is explained on concept level and illustrated with evaluating tests of the engineering model. Finally, an outlook on the future of the project, including the current design status of the engineering qualification model, is given.
A specific design of the Collapsible Tube Mast (CTM) to be flown aboard the NASA STS is being prepared by SENER. The work is done under ESA contract, as part of the Technology Demonstration Program (TDP). The CTM is intended to deploy and retract a 15 kg In Flight Contamination Experiment (IFCE) at distance up to 15 m from the Orbiter cargo bay. Objectives of the mission are to measure the contamination in the vicinity of the Orbiter, as well as to demonstrate the CTM performances in-orbit (this is the first flight of a retractable CTM).
The dimensions of space borne instrument-, antenna- and solar array structures do often exceed the available envelope of common launch vehicles. To realize such systems anyhow the structures are provided with folding mechanisms that allow a space saving orbit transfer. A side effect from this deployment concept is the fact that a structure which provides a sufficient deployed stiffness for space environment can also fulfil the stiffness requirements for launcher payload in stowed configuration. Consequently, the sizing of such structures demands the consideration of different configurations that are composed out of the same essential structural parts. The thesis, thereby, illustrates how the roles of the single components change for the different configurations. For instance, a component that carries a major part of the load in a first configuration can be nearly offloaded in a second one. Moreover, this component can load another part with its own mass, whereas, this other component is passive in the first configuration and needs to be carried. Thus, the core of this thesis contains the introduction into an exemplary antenna structure design as well as a method for an adequate sizing of all relevant structural parts. The developed method is then integrated in a Finite Element Analysis (FEA) aided, closed loop sizing chain. The used automated close loop sizing helps to rapidly react on changed requirements and generates the possibility of performing fast parameter studies to improve the understanding of such a sophisticated structure. The validity of these approaches is finally proven by presenting a re-sized antenna configuration after the parameter study and its evaluation. This final configuration more than meets the previously defined requirements for the exemplary antenna structure.