Inflatable Solar Arrays: Revolutionary Technology? - Team-Logic

1999-01-2551
Inflatable Solar Arrays: Revolutionary Technology
Mark S. Grahne, David P. Cadogan
ILC Dover, Inc.
Copyright © 1998 Society of Automotive Engineers, Inc
ABSTRACT
Recent technological advancements in space inflatable
structures, in the areas of material rigidization and
controlled deployment, have presented a new possibility to
the space community with a low cost, lightweight
alternative to mechanically deployed space structures.
Space inflatable structures have many benefits and
advantages over current mechanical systems. They are
low in mass and can be packaged into small volumes,
which can potentially reduce the overall program cost by
reducing the launch vehicle size. Reduction in total
system mass and deployment complexity can also
increase system reliability.
This new technology is fast becoming a reality, especially
in the field of inflatable solar arrays and other applications
for spacecraft components. Many solar array applications
such as the Mars rover inflatable solar array, the JPL
Deep Space Four (ST4) inflatable blanket solar array, and
the Teledesic blanket solar array have been developed
and prototypes have been built and tested. Several flight
experiments are underway and will be flown in the very
near future.
INTRODUCTION
Recently, ILC Dover has developed several methods of
deploying and rigidizing inflatable structures in orbit or on
planetary surfaces. These technologies enable a flexible
composite laminate structure to be densely packaged for
launch, deployed via inflation gas in space, and finally
cured, or rigidized, in orbit (in situ) without further need of
inflation pressure to maintain structural integrity.
In conjunction with the recent enhancements made in the
space inflatable structures technology, the advancements
in the technology of flexible thin film solar cells have
made the inflatable solar array very attractive to the next
generation of spacecraft. The combination of these two
technologies enable the further development of
lightweight, low cost, and small packing volume solar
arrays. In other words, spacecraft could potentially have
larger and more powerful solar arrays without the
undesirable mass and storage penalty of rigid array
systems if these technologies were employed.
This paper describes several types of inflatable solar
arrays, and related applications for the technology that
may benefit solar array technology in the future. A
number of specific missions, which are under
development, are also discussed. Recent developments
in rigidization methods and controlled deployment are also
presented to support the flight readiness of this
technology.
With the advent of new space inflatable structure
technologies, many new applications have been
envisioned. Applications of inflatable systems including
Mars rover solar arrays, satellite blanket solar arrays,
reflectarray radar antennas, communications antennas,
synthetic aperture radar arrays, solar concentrators,
structural members such as booms and trusses, impact
attenuation devices, habitats, and sun shields, have been
developed and enhanced through many pre-flight
programs. Out of these applications, inflatable solar
arrays are one of the most promising candidates thus far
for further development to advance the state-of-the-art of
satellite structures, especially in the areas of specific
power and reducing system cost.
1
SPACE INFLATABLE TECHNOLOGY
ENHANCEMENTS
Two areas of enhancement that contributed directly to the
advancement of the application of inflatable solar array
technology are the development of structural materials
with various rigidization methods and the development of
controlled deployment mechanisms. In general, space
inflatable structures are fabricated from flexible materials
(thin films or coated fabrics) that are made structural by
internal pressure. Environmental threats such as
micrometeoroids and debris impact increase the
probability of inflatable structure leakage over time. This
given, in orbit rigidization of the structure is needed to
extend system life. Pure inflatable structures are used
only for missions with short operational life and where the

supply of make up gas does not pose a problem. Most of
the applications discussed below require in situ rigidization
to provide long term structural integrity.
RIGIDIZABLE STRUCTURES
Space rigidized structures are fabricated from flexible
composite laminates that are rigidized in situ via some
external influence. They can be fabricated into many
different shapes such as toroids, spheres, dish structures,
tubes, etc., which can be designed into various types of
structures. This class of structures can be deployed in
various orbits and in gravitational surface environments
such as the Moon or Mars. The rigidized components are
designed for typical operational lifetimes of seven to
fifteen years without environmental concerns.
Over the last forty years many rigidization methods have
been investigated sporadically. However, with more
intensive research and development in the recent years by
the space inflatable industry, many reliable rigidized
structural components have been produced through the
use of advanced materials and design. Some of the most
promising rigidization methods include:
• Heat Cured Thermoset Composite Laminates
(Thermal Heating)
• Thin-walled Aluminum/polyimide Laminates
• Thermoplastic Composite Laminates (Passive
Cooling)
• UV Curable Composite Laminates
• Foam Inflation
• Inflation Gas Reaction Laminates
Out of these rigidization methods ‘Thermal Heating’ and
‘Thin-walled Aluminum’ are the most promising for space
and planetary surface applications respectively. The
Thermal Heating method is currently under further
development at ILC Dover for flight experiment in the year
2000. The Thin-walled Aluminum System and the
Thermal Heating System will be discussed in greater
detail later. For further information on other methods of
rigidization the reader is encouraged to consult the
reference documents listed in the paper.
Figure 1 - Composite Laminate Cross-Section
The MLI heating blanket is designed and fabricated from
various layers of vapor-deposited aluminum (VDA)
Kapton, VDA Mylar, and spacers. The purpose of the MLI
blanket is to keep out the harsh space environment and at
the same time maintain the required curing temperature
inside. The design of the MLI blanket is tailored to the
specific mission environment and requirements. The
support tube laminate typically consists of four separate
layers: (1) The restraint layer that maintains the shape of
the inflatable structure; (2) The heater assembly layer
which provides the proper temperature for deployment
and curing (in some designs the restraint layer and the
heater layer are combined to become one assembly); (3)
The rigidizable composite laminate layer, fabricated
from prepreg materials such as epoxy/graphite, is the
support structure when cured; And (4) The bladder layer,
manufactured from black Kapton, keeps and maintains
inflation pressure during deployment. The thermoplastic
composite laminates, which use passive cooling as the
curing method, has a similar construction.
Thin-walled Aluminum Method
In this approach of rigidization a laminate is fabricated
from Kapton film and ductile aluminum, where the Kapton
film is positioned on both sides of the aluminum as shown
in Figure 2.
Thermal Heating Method
The composite laminate system, which consists of a
thermoset matrix resin and a fiber reinforcement such as
graphite, is cured or rigidized by heating. The thermoset
resin hardens after being heated to a specified
temperature and cure time. This rigidization method can
be designed to cure from solar energy, or from the
spacecraft power, or from a combination of both. The
properties of the composite material are consistent with
those used in today’s spacecraft design. A typical
composite laminate cross-section, as shown in Figure 1,
consists of two multiple-layered components: (1) the MLI
blanket, and (2) the support tube laminate.
Figure 2 - Aluminum Laminate Cross-Section
2
To rigidize the cross-section, the structure is inflated to
eliminate the wrinkles in the laminate and to a point of just
yielding the aluminum. After yielding the aluminum the
Kapton/aluminum laminate is ‘rigidized’ and will maintain
structural integrity. The inflation gas is then vented to
space. A typical aluminum laminate cross-section

consists of two multiple-layered components: (1) the MLI
blanket, and (2) the support tube laminate. The MLI
blanket is similar to that used for thermal cured laminates,
and it is tailored to mission environment and
requirements. The support tube laminate is fabricated
from layers of Kapton-adhesive-aluminum-adhesive-
Kapton laminate. The thickness of each layer and the
number of layers used can be tailored and designed to
mission specific requirements.
Controlled Deployment System
In conjunction with the development of rigidization
methods, ILC Dover has also developed a number of
deployment techniques to ensure a proper and controlled
deployment of inflatable structures. The purpose for the
controlled deployment system is to (1) keep the deploying
system within a known envelope, (2) improve structural
reliability during deployment by avoiding entanglement
with itself or other components, and (3) minimize shock or
impulse induced to the spacecraft during deployment.
Some of the most promising controlled deployment
devices include:
• Columnation Devices
• Roll-up/Reverse Roll-up Devices
• Internal Compartmentalization
• Break Cords/Peel Flaps
• Becket Loops
Out of these controlled deployment devices the Roll-up
design and the Columnation design are the most
promising for the solar array application. The Roll-up
design is currently under further development at ILC
Dover for flight experiment in the near future. The Roll-up
and the Columnation Devices will be discussed next in
greater detail. The reader is encouraged to consult the
reference documents for more detail.
Roll-up Device
One method of deployment control that has seen wide
application in tubes and struts is the roll-up. This
approach utilizes a rolled inflatable tube with an
embedded mechanism to control its rate of unrolling when
inflation gas is introduced. This is similar to a common
party favor that unfurls when you blow into it. There are
two classes of deployment control used in roll-up devices;
1)mechanisms embedded in the tube itself, and 2)
mechanisms mounted at the end of the tube.
Figure 3 - Roll-up Device (with Membrane)
The second class of deployment control method involves
utilizing a torque mechanism at the end of the tube. An
ILC Dover proprietary ‘Torque Mechanism’ is located in
the end cap to provide resistance to unrolling so that
interim beam stiffness during deployment is achieved.
This approach is simple, reliable, compact, and the roll-up
is a benign packing procedure for rigidizable laminates.
Columnation Devices
The columnation device, Figure 4, provides a deployment
method that when inflated extends axially in a straight
telescopic motion with some degree of beam stiffness.
The inflatable tube is axially collapsed on a short mandrel
in its packed state. The top of the mandrel has a tube
compression feature (seal) that applies some resistance to
axial motion of the tube. Inflation gas is introduced
through the mandrel into the forward end of the tube. As
plug load builds in the inflatable, it overcomes the
frictional resistance and advances the tube. This method
places axial load in the tube wall during deployment,
allowing it to behave as an inflated beam and exhibit
structural stiffness. Several variations of this device,
including inflatable and collapsible mandrels that allow the
packed beam to fit into very small volumes, have been
developed and tested with good results.
Embedded mechanisms include Velcro strips mounted
longitudinally on the tube’s exterior, or constant force
springs mounted to the tube's interior in the same fashion.
Introduction of gas into the tube would separate the Velcro
due to shape change during inflation, or unroll the springs.
Each system provides a calculable value of resistance,
which dictates the internal pressure in the tube. This in
turn dictates the interim beam stiffness of the tube during
deployment, prior to rigidization.
Figure 4 - Columnation Device
3

INFLATABLE SOLAR ARRAYS
The advantages of using inflatable systems technology in
designing a large solar array are reduced stowage volume
and mass, increased specific power (greater than 100
W/kg), and reduced cost over current mechanically
deployed solar arrays. The inflatable solar array is
particularly attractive for missions that demand high power
output with launch vehicle size restrictions. ILC is working
or has developed a number of different solar array
systems for planetary or space use. A few of these
systems are described below.
ST4 INFLATABLE SOLAR ARRAY
The JPL ST4 spacecraft requires high power output arrays
(2 wings at 6 kW per wing) to accomplish the mission of
rendezvous with the Temple 1 comet and landing to
perform scientific measurements. The ST4 inflatable
solar array is a 3-meter by 15-meter solar array capable of
producing 6 kW of power. The array is currently under
development for a shuttle flight experiment in late 2000.
ILC Dover is the prime contractor to JPL for this
experiment, with AEC-Able developing the array blanket
and L’Garde developing the instrumentation. The basic
configuration of the ST4 inflatable solar array, Figure 5,
consists of four major subsystems:
• Solar Array Blanket
• Structural Support Components
• Controlled Deployment System
• Inflation System
Top Panel
(Full length)
Inflatable/Rigidizable Tube
Solar
Panels
9.8 ft [3.00 m]
Solar Array Blanket
48 ft [14.65 m]
Bottom Panel
(Split)
Figure 5 - ST4 Inflatable Solar Array
Inflation
System
The configuration of the ST4 solar array is a modular split
blanket style with the deployment tube located on the
array centerline. When stowed, the solar array modules
are accordion-folded. Currently, Sharp’s rigid highefficiency
silicon photovoltaic assemblies are the basic
building blocks for the modular blanket solar array.
Flexible thin film solar cells using amorphous silicon,
copper indium gallium diselenide or other materials hold
great promise to provide even lower cost and lighter
Gimbal
weight photovoltaic modules in similar arrays in the future.
listed in this paper.
Structural Support Components
This subsystem includes the inflatable beam, stowage
panels, plume-offset panels, launch ties, and launch tie
release mechanism. The inflatable beam can be
fabricated from any of the aforementioned rigidization
systems. However, the thermal heating method will be
utilized for the flight experiment.
The purpose of the inflatable beam is to provide a
deployment mechanism and support structure for the solar
array. It is similar in function to the mechanical
deployment truss masts currently utilized on spacecraft.
The beam is located on center of the split blanket and is
stowed by rolling the tube, see Figure 6.
22 Solar Panels
Roll-up Tube
118.11” [3000.00 mm]
52.70” [1338.58 mm] TYP.
Launch Ties
Top Panel
22.00” [558.80 mm]
Inflation System
14.00” [355.60 mm]
4.50” [114.30 mm]
Figure 6 - ST4 Inflatable Solar Array Stowed Configuration
Controlled Deployment System
The ILC roll-up device with the embedded torque
mechanism is the current baseline design for the ST4
inflatable solar array deployment system. When gas is
introduced in the base of the tube, the tube inflates and
begins to unfurl. The torque mechanism provides
resistance to the unrolling action and builds pressure in
the tube at a uniform and constant level. The pressure in
the tube yields a tensile stress in the tube wall, which
provides interim beam stiffness during deployment. The
torque mechanism is attached to the tip spreader plate on
the array, which deploys the array blanket. Guide wires
are also attached to both spreader plates and the blanket
for control of the blanket during deployment.
Inflation System
Inflation systems for the space inflatable structures have
included compressed gas, gas generators, and
sublimation of powders and liquids into gases. Each of
these approaches has its advantages and disadvantages
depending on the specific application. In recent years,
compressed gas systems have become very low mass
and highly reliable, making them the most practical option
for most space inflatable applications. A compressed gas
system is currently baselined for the ST4 inflatable solar
array inflation system.
4

MARS ROVER INFLATABLE SOLAR ARRAY
TELEDESIC INFLATABLE SOLAR ARRAY
The Teledesic program envisions a constellation of 288
satellites in low earth orbit to provide high data rate
communication from anywhere on earth. Original
concepts for the satellite included a pair of 6 kW solar
array, see Figure 7, with a cost target of $100/W in
production. Inflatable technology was considered to be a
leading candidate in meeting this goal and was therefore
investigated by the Teledesic team.
JPL is developing several advanced rover vehicles for the
exploration of the Martian surface. Several concepts call
for large deployable solar arrays to meet the power needs
of the rover. One such concept, see Figure 9, utilizes 1.5
meter deployable wheels and an inflatable solar array to
cover vast surface areas in rapid times.
Figure 9 - Inflatable Mars Rover Prototype
Figure 7 - Teledesic Satellite
ILC Dover, under contract to Boeing, designed an
inflatable structure to support a 3-meter by 10-meter
rectangular satellite solar array. A full-scale prototype
demonstration unit was also fabricated and used in
deployment trials, see Figure 8.
The Mars Rover Solar Array prototype, Figure 10, is a
working full-scale inflatable solar array for rover
application. This prototype is a system that would be
packaged in a small volume for launch and deployed in
situ to collect solar energy. The array is parasol shaped
and consists of four main components: (1) the canopy,
which is the membrane that carries the solar modules; (2)
the inflatable torus; (3) the inflatable column; and (4)
sixteen solar modules. The sixteen (16) Kapton gores (the
canopy) of the solar array are tensioned by a 8.0 cm
(tube) diameter by 150 cm (major) diameter 16-sided
inflatable torus, and supported by a 10.0 cm diameter
inflatable column.
Due to the requirement for multiple deployments, the torus
of the prototype is constructed from a lightweight
aluminized Mylar assembled by pressure sensitive
adhesive tape of the same material. The prototype
inflatable column is fabricated from a laminate of Kapton
and aluminized Mylar. The inflatable torus and column of
the future flight unit will be constructed from thin-walled
aluminum/Kapton laminates or UV cured laminates.
Figure 8 - Teledesic Inflatable Solar Array
Due to the limited packing volume of the array, the mast
system is designed with a deployment tube, whose only
function is to deploy the array, and two composite
rigidizable tubes to provide structural support after
deployment. This design was also constrained by a
requirement for limiting frontal area in any orientation to
reduce drag during orbital ascent. This requirement
forced the use of a three-tube system rather than a single
multifunctional tube structure. The mast structure is
located on the backside of the array. The deployment
tube uses an inflatable columnation device for controlled
deployment. The two structural tubes are z-folded into a
small compact stowage volume and pulled out during
deployment.
Solar cells selected for the prototype array are
manufactured by Iowa Thin Film Technologies Inc. These
modules consist of tandem amorphous silicon (a-Si) solar
cell material deposited on 2 mil thick polyimide. Each
solar module is made of three, 12 volt, 50 mA, solar
panels connected in parallel with a blocking diode. All
three panels and the blocking diode are enclosed inside 5
mil polyester encapsulant. Each module is capable of
generating 12 volt and 150 mA at 1000-watts/m 2 light
intensity with 5% cell efficiency. The Mars solar array
prototype is therefore capable of generating approximately
20 watts of power in a terrestrial environment or 12 watts
in a Martian environment. Further optimization in cell
population pattern, cell efficiency and mass will further
increase specific power to weight ratio.
5

SYNTHETIC APERTURE RADAR ARRAY (SAR)
Figure 10 - Mars Rover Inflatable Solar Array
POWER SPHERE
ILC Dover together with the Aerospace Corporation
developed concepts for a Power Sphere solar array, which
would enable the use of micro and nano satellites. The
design of the power sphere allows for constant power
generation independent of the sphere’s orientation. The
sphere could be packaged in very small stowage volume
and would deploy and inflate using sublimating powder or
compressed gas. The power sphere would generate in the
neighborhood of 5watss of power with a system mass of
600grams. ILC Dover fabricated a proof of concept unit
for demonstration purposes, Figure 11.
The SAR program was a development effort performed by
ILC Dover for NASA’s JPL to demonstrate the feasibility of
a low mass inflatable Synthetic Aperture Radar Array
Antenna. The prototype system, see Figure 12, was a flat
multiple-layered metalized thin film microstrip array (a
three-membrane assembly) tensioned and supported by
an inflated frame that can be rigidized after deployment.
The size of the full-scale array was a 10-m x 3.3-m
structure. The system was stowed in a rolled-up
configuration on the spacecraft bus, and was deployed in
a controlled manner via inflation gas. The objective of
this program was to develop a functional subscale system
that was less than 2 kg/m 2 of radiating area when
projected to full scale. The subscale prototype met the
requirements and demonstrated the feasibility of the low
mass SAR concept.
Figure 12 - Synthetic Aperture Radar Array
The deployment method and supporting frame concept
used in the SAR could also be used for a flexible blanket
type solar array in future developments. That is, the
membrane assembly of the SAR could be replaced with a
flexible thin film solar cell blanket and the system would
then act as an inflatable solar array.
JPL 3-METER KA-BAND INFLATABLE
REFLECTARRAY
Figure 11 – Power Sphere
RELATED APPLICATIONS
Developments in related applications are producing more
packaging and deployment concepts that are applicable
for deploying future inflatable solar arrays with flexible
thin-film array blankets. Investigations in these projects
are pushing the inflatable technologies forward both in
rigidizable structures and controlled deployment systems.
In conjunction with improvements made in the inflatable
technologies, further advancements in the flexible thinfilm
photovoltaic cells will make these concepts prime
candidates for future inflatable solar arrays. The concepts
discussed below also employed many of the latest
advancements in the inflatable technology.
The 3-Meter Ka-Band Reflectarray program is another
development effort performed by ILC Dover for JPL to
demonstrate the feasibility of super low mass
telecommunication and spaceborne SAR concepts. This
3-Meter prototype, see Figure 13, consists of four major
subassemblies: (1) the membrane assembly; (2) the rigid
frame assembly; (3) the inflatable frame assembly; and
(4) the suspension system.
6

Figure 13 - JPL 3-Meter Ka-Band Inflatable Microstrip
Reflectarray
The reflectarray membrane, supplied by JPL and
assembled by ILC, is suspended by a horseshoe shaped
structure assembled from a straight rigid frame assembly
and a U-shaped inflatable frame assembly. The rigid
frame assembly is made from graphite epoxy with
aluminum end caps. The inflatable frame assembly is
fabricated from urethane coated Kevlar to simulate a
rigidizable material in the final application. The
membrane support frame is 25 cm in diameter and the
feed horn support torus is 7.5 cm in diameter. The feed
horn torus is supported by three vertical struts grouped
within 90 degrees of each other. The struts are tapered in
diameter, from the larger diameter of the membrane
support tube to the smaller diameter of the feed horn
torus, to minimize material in the RF path as well as
system mass. One of the critical requirements for this
program is the flatness of the membrane. For this reason
many packaging and deployment concepts were
considered, with the current configuration being the best
for this particular application. Again, similar to the
rectangular frame used for the SAR array, the horseshoe
frame concept could also be used for solar array
deployment.
NGST (ISIS) SUNSHIELD
ILC Dover is working with JPL and GSFC on the NGST
(ISIS) Sunshield program. The goal of the program is to
demonstrate a 1/3-scale model of the Next Generation
Space Telescope (NGST) sunshield. A ½ scale model
has already been demonstrated in 1G and the current
program calls for the 1/3 scale model to be demonstrated
on the shuttle in 2000. The sunshield is a diamond
shaped membrane structure that measures approximately
5 x 11 x 1 meters when deployed. The system consists of
4 layers of membranes that are deployed and supported
by 4 inflatable beams. The membranes, when stowed, are
accordion-folded or “Z” folded into a small packing volume
around the bus structure. The roll-up method for
controlled deployment and the heat cured thermoset
composite laminates are baselined for this demonstration.
The system shown in Figure 14 is a half scale model
deployed at ILC with a gravity negation system attached to
the sunshield.
Figure 14 - NGST Half Scale Sunshield
CONCLUSIONS
Inflatable solar array technology continues to mature and
expand the possibilities for large-scale satellite solar
arrays as well as planetary surface arrays for rover use.
Technology advancements in inflatable rigidizable
structures and controlled deployment systems, in
conjunction with advancements in solar cell technology
have lead to increases in specific power, reductions in
system cost, and minimization of package volume for
solar arrays. Lab and thermal vacuum chamber
demonstrations of this technology have shown the viability
of the approach. The flight demonstration of the ST4 solar
array will provide the necessary flight heritage to allow the
use of this technology for future deep space and
commercial missions.
CONTACT
Mark S. Grahne
ILC Dover, Inc.
One Moonwalker Road
Frederica, DE 19946
Grahnm@ilcdover.com
302.335.3911x462
302.335.1320 fax
www.ilcdover.com
REFERENCES
1. D.P. Cadogan, M. Grahne, and M. Mikulas, “Inflatable
Space Structures: A New Paradigm For Space
Structure Design,” IAF-98-I.1.02, 49th International
Astronautical Congress, Melbourne, Australia,
September 28 - October 2, 1998.
2. C. M. Satter and R. E. Freeland, “Inflatable Structures
Technology Applications and Requirements,” AIAA 95-
3737, AIAA 1995 Space Programs and Technologies
Conference, Huntsville, AL, September 26-28, 1995.
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3. D.P. Cadogan, R.W. Lingo, and C.R. Sandy, “Proof-
Of-Concept Demonstration Model of the Inflatable
Roll-Up Synthetic Aperture Radar Array Antenna,” ILC
Dover Final Report for JPL Contract No. 96836,
September 1997.
4. M.C. Lou, V.A. Feria, and J. Huang, “Development of
An Inflatable Space Synthetic Aperture Radar,” AIAA
98-2103, 1998.
5. V.A. Feria, M.C. Lou, J. Huang, and S.E. Speer,
“Lightweight Deployable Space Radar Arrays,” AIAA
98-1933, 1998
6. J. K. Lin and D. P. Cadogan, “Concept Development,
Design, and Fabrication of an Inflatable Solar Array
for a Mars Rover Application,” ILC Dover Final Report
for JPL, June 1998.
7. E.S. Fairbanks and M.T. Gates, “Adaptation of Thin-
Film Photovoltaic Technology For Use in Space,” 26th
IEEE PVSC, Anaheim, California, 1997.
8. S. Guha, et. al., “Recent Progress in Amorphous
Silicon Alloy Leading to 13% Stable Cell Efficiency,”
26th IEEE PVSC, Anaheim, California, 1997.
9. F. Jeffrey, et. al., “Lightweight, Flexible, Monolithic
Thin-Film Amorphous Silicon Modules on Continuous
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