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Abstract
Large, space-based antennas are needed for a variety of different
applications. Since there is no meaningful orbital assembly capability
planned at this time, any large space structures will have to
be self deployable. Current concepts for large, conventional,
mechanical, self-deployable space structures tend to be very expensive
and mechanically complicated. Current antenna-user requirements
are so stringent (with respect to the need for very low-cost,
high-deployment reliability, low weight, and packaged-volume and
usable aperture precision) that new and innovative approaches
to accommodate large space structures are needed. Fortunately,
a newly developed class of space structures, called inflatable-deployable
structures, has great potential for satisfying these stringent
user requirements. A concept under development at L'Garde, Inc.,
for a large, inflatable-deployable antenna represents an excellent
example of this new type of structure.
The NASA Office of Space Access and Technology initiated the In-Space
Technology Experiments Program (IN-STEP) specifically to accommodate
the verification and/or validation of unique, innovative, and
high-payoff technologies in the space environment. The potential
of the L'Garde, Inc., concept has been recognized and resulted
in its selection for an IN-STEP experiment. The objectives of
the experiment are to verify low cost and light weight by building
a 14-meter-diameter flight-quality reflector antenna structure,
demonstrate deployable reliability in a realistic environment,
and measure the reflector surface precision in a realistic gravity
and thermal environment. The approach will utilize the Space Transportation
System (STS)-launched, recoverable Spartan spacecraft as the experiment
carrier.
The flight-system functional performance requirements originate
from the experiments technical objectives. The design requirements
for the flight hardware system are based on a combination of system
functional-performance requirements; basic inflatable structures
capability; the L'Garde, Inc., technical data base resulting from
the development and launch of a large number of inflatable "decoy
type" structures and the space environment effects on a large,
thin-film structure in low earth orbit. The large-space structure/environmental
interactions include the effects of atmospheric drag on the attitude
stability of the structure, the effects of the orbital thermal
environment and atomic oxygen on the thin-film materials, and
low orbit radiation effects on electronic components. Both requirements
and environmental effects are specified for each subsystem; e.g.,
(a) the basic support structure that houses the inflatable structure
and the other subsystems and interfaces with the experiment carrier,
the Spartan, (b) the inflatable-deployable antenna structure,
(c) the inflation system, (d) the surface-measurement system,
and (e) the electronic system. For each of the subsystems, this
paper will identify and describe the key
and unique design drivers, the impact of the environmental interactions,
the type of analysis used for simulating subsystem performance,
the type of developmental testing used for design refinement and
validation, the specialized processing, manufacturing and assembly
techniques, and description of the final design.
The experiment is being managed by the Jet Propulsion Laboratory.
The flight hardware development at L'Garde, Inc., Tustin, California,
is currently in Phase C/D, and the experiment is manifested to
fly on STS 77 in April 1996 as primary/sharing payload.
Introduction
Space-deployable antennas are needed for a variety of applications
that include space-based, very long-baseline interferometry (VLBI),
mobile communications, active microwave sensing, earth observation
radiometry, synthetic aperture radar, spacecraft communications,
and DOD space-based radar.1 Recent constraints on the availability
of resources for these types of applications within NASA, ne science
community, the commercial sector, and even the DOD, have resulted
in stringent user application requirements. Therefore, the real
key to accommodating these missions is the development of new
concepts for low-cost and mechanically reliable antenna structures.
Other important features include low weight, high mechanical-packaging
efficiency, usable aperture precision, and long-term dimensional
stability. Realistically, however, meaningful demonstrations of
innovative concept capabilities will have to be accomplished to
attract any kind of serious user interest.
A relatively new and unique concept for an inflatable-deployable
space antenna structure that has tremendous potential for accommodating
such stringent user requirements is under development by L'Garde,
Inc., Tustin, California. 1,3 In fact, serious user interest has
resulted in the selection of this concept for a NASA In-Space
Technology Experiments Program (IN-STEP) space-based experiment.
This class of experiments is based on demonstrating and evaluating
the performance of promising concepts with low-cost flight hardware.
The experiment objectives are selected specifically to validate
antenna-user criteria and to demonstrate the development of large,
flight-quality hardware for a low cost, high mechanical-packaging
efficiency, low weight, requirements high deployment reliability,
usable reflector-surface requirements,
precision, an(l thermal stability in a realistic environment
The experiment is currently in the final stage of flight-hardware
assembly and qualification testing. It is manifested to be flown
on STS 77 in late April 1996. This paper describes the design
of the experiment flight hardware and identifies the key issues
for each of the subsystems that comprise the experiment system.
The information contained in this paper and in References 1 and
2 is intended to provide a complete summary of the experiment
justification, technical approach and flight hardware.
Experiment-System Performance Requirements
The experiment-system functional requirements are based on
the experiment objectives and the inflatable structures concept
capability, constrained by the NASA experiment resources available
and the capability of the experiment carrier, the Spartan (Figure
1). The antenna structural configuration is based on the L'Garde,
Inc., basic inflatable-antenna concept. The 14-meter-diameter
reflector size is based on an extrapolation of the 9-meter baseline
structures data base and the current size limit for manufacturing
capability at L'Garde, Inc. Moreover, this structure can be accommodated
by Spartan, and it is large enough to be used for real applications,
such as VLBI and commercial mobile communications. The surface-precision
goal of 1 mm rms on orbit is based on the current analytical performance
projections, manufacturing, assembly, and alignment capability
at L'Garde, Inc. Validation and characterization of the deployment
sequence will be done on orbit, which provides a realistic operational
environment. High mechanical-packaging efficiency will be demonstrated
by stowing the inflatable structure in a small canister. The inflight
single-orbit measurement of surface precision and its thermal
stability will provide a measurement of the concept value for
different potential applications.
Qualification of the experiments hardware is being accomplished
by both analysis and test for compliance with functional performance
and STS safety requirements.
Subsystem Functional Requirements
The experiment-subsystem functional requirements are driven by the system functional requirements with design parameters bounded by the L'Garde, Inc., flight data base for inflatable structures and the environmental-interaction effects on the experiment hardware (Figure l). The subsystems needed to accommodate the experiment include (a) the inflatable structure, (b) the canister or bus structure, (c) the inflation system, (d) the surface-measurement system, and (e) the electronic system (Figure 2). The combination of these subsystems represents the simplest approach for satisfying the system functional requirements. The design and performance of the actual flight hardware will be based on how well the subsystem functional requirements were satisfied. The key design drivers, the types of analysis used, the functional developmental testing, and pictures of the final designs will be discussed in the appropriate subsystem sections of this paper.
Inflatable Structure
The subsystem sunctional requirements for the inflatable structure are given by Table 1. The antenna configuration is an off-axis parabolic reflector structure consisting of (a) a 14-meter-diameter, multiple-gore reflector structure and a transparent canopy (which is a mirror shape of the reflector) to maintain gas pressure on orbit, (b) a torus structure that supports the reflector/canopy circumferentially, and (c) three 28 meter-long struts that interface the torus structure with the canister which is located at the center of curvature of the reflector to accommodate operation of the surface measurement system.
Reflector
The major challenge is to design and fabricate a 14-meter-diameter,
multiple-gore reflector structure with an orbital surface precision
on the order of 1 mm rms and with enough reflectivity to accommodate
orbital operation of the Surface Accuracy Measurement System (SAMS).
The system reflector-gore geometry is determined by the L'Garde,
Inc., FLATE code, which uses the desired orbital-membrane configuration,
its operating stress levels, and materials properties of the membrane
and bonded seams to define the number of gores, their zero stress
shape, and the operating gas pressure.4 The membrane-materials
properties (which include thickness and non-linear modulus) are
experimentally characterized for use in the detail design. The
techniques for handling, laying out, marking, cutting, and the
butt-joint bonding of the one-quarter-mil mylar membrane were
developed on previous programs at L'Garde, Inc., and demonstrated
on a 9-meter diameter reflector structure. Mylar was selected
for this experiment because of its availability, low cost, and
its extensive use in previous flight applications of inflatable
structures.
Fabrication of the 14-meter-diameter reflector was based on using
62 individual one-quarter-mil aluminized mylar gores (Figure 3).
The gores were assembled on full-scale, specialized tooling that
was designed to account for the difference in curvature between
the zero strain-assembly condition of the membrane and the orbitally
loaded configuration. The seams are butt-joined, utilizing a doubler
of the same material on one side only. The adhesive was a standard,
space-qualified, flexible material used on a number of previous
programs. The ground handling of the membrane required the development
of special folding techniques so that the material could be stowed
in a small compact package for ease of handling and deployment.
The surface precision of the as-manufactured reflector structure
is determined by mounting the membrane on a fixture that simulates
its interface with the canopy stmcture and torus. A pressure differential
across the structure (equivalent to that on orbit) produces a
surf;3ce that represents the "upper bound" of reflector
precision that would be achieved if the assembly were perfect.
The flight-reflector structure has a measured surface precision
on the order of 1 to 2 mm rms.
Table 1. Inflatable Structure Subsystem Requirements
|
| · Based on L'Garde, Inc., inflatable-deployable antenna concept · I4-meter off-axis parabolic reflector · Surface accuracy goal of 1 mm rms · Optically reflective surface on reflector · Clear canopy · Torus and struts provide basic support structure · Packagable in a container compatible with carrier vehicle · t/D = 1/2 (Parent) · Deployment time compatible with single-orbit experiment · Structural Stability · Dimensional Stability |
Canopy
The primary design requirement for the canopy is that it should be a "mirror" image of the reflector structure. That is, the design should be based on using the same materials and number of gores, but would not require the same surface precision or vapor-deposited aluminum as that used on the mylar. This approach will result in a force balance between the two doubly curved surfaces at their interface where they are assembled to form a lenticular structure. Since this canopy structure used for the reflector, no special tests were done to verify the canopy's configuration. Assembly of the canopy with the reflector is accomplished by bonding both structures at their outer perimeter to a flexible ring structure. This interface structure provides sufficient stiffness to transfer the loads from the inflated lenticular structure to the torus, yet it is flexible enough to be stowed with the membrane structure. Mechanical packaging tests of this structure are accomplished after assembly of the "lenticular" structure.
Torus
The design of the torus structure is driven by the size of
the reflector and the circumferential tension loads from the lenticular
structure. These loads produce compression and bending in the
torus structure. The material selected was neoprene-coated kevlar
because (a) L'Garde, Inc., has extensive experience with the handling
and bonding of this material, (b) it is commercially available
and inexpensive, (c) it stows efficiently, and (d) it has adequate
strength and stiffness for accommodating the experiment. The detail
design of the torus is based on a L'Garde, Inc., specialized code,
which uses as input the external loading on the torus, its geometry,
operating stress level, and the material properties to determine
the required diameter and operating pressure.5 The material properties
for the analysis are experimentally characterized. A full-scale
engineering model torus is shown in Figure 4.
The specialized tests for the torus include mechanical packaging
efficiency and neutral buoyancy flotation to determine the attachment
plane for the torus/lenticular reflector assembly under zero g
loading conditions. The mechanical packaging consists of repeated
stowing of the torus in the canister structure, using different
"folding" techniques. Success for this test is determined
by the best packaging efficiency, coupled to a final folding configuration
that lends itself to "deployment by inflation", as established
by previous flight-hardware experience. The neutral buoyancy state
for the torus is achieved by (a) completely filling the structure
with water, (b) locating it in a trough filled with water, (c)
pressurizing the torus to a differential pressure of 0.02 MPa,
which simulates orbital loading, and (d) applying a small amount
of distributed flotation to offset the negative buoyant forces
resulting from the fluid displaced by the volume of the fabric
used for the structure. This technique worked so well that the
torus could be manually displaced in the trough with essentially
no measurable restoring forces observed. In this unloaded state,
the mounting plane for the lenticular structure was located with
a rotating laser beam and marked for subsequent attachment of
the interface hardware.
Assembly of the torus/lenticular structure is based on using 62
discrete mechanisms located along the mounting plane that can
be adjusted individually to impart a uniformly distributed load
to the torus. The adjustments are interactively made to simulate
the circumferential geometry of the lenticular structure when
properly aligned on its mounting fixturing. After the integration
of the two structures, additional mechanical packaging tests are
done, starting with the folding configurations successfully developed
with the torus structure.
Struts
The design drivers for the struts include (a) a required length
of 28 meters, (b) a minimum structural frequency of 0.25 Hz, (c)
manufacturing tolerances of 0.5 cm for bending and torsion distortions,
and (d) an operating pressure the same as that for the torus.
The detail design is based on using the same neoprene-coated kevlar
material as used on the torus. The diameter and resulting bending
stiffness are based on a requirement for a minimum natural frequency
to accommodate the orbital stability needed for the experiment.
The minimum diameter required was 35.6 cm. The materials-processing
and fabrication techniques used are near identical to the ones
used on the torus. A full-scale engineering model is shown in
Figure 5.
Specialized testing included manufacturing evaluation for bending,
twist, and mechanical packaging efficiency. The quality of the
manufactured struts was established by "floating" them
in a full-length water trough. Calibration marks on the ends of
the tubes indicated the degree of relative rotation. Measurement
of lateral translation along the length of the strut, as it is
rotated in the water trough, is a direct measure of the bending
as a result of manufacturing. These tests showed that the torsional
permanent set in the structure was about 15°, but did not
affect the functional performance, and the bending was on the
order of 50 mm maximum deflection.
The end fittings for the struts are essentially flat plates machined
from aluminum and bonded to the interior area of the end of the
tubes. Such fittings are simple, inexpensive, and easy to interface
with the canister panel-support fittings and the torus-to-strut
interface fittings. The geometry of these fittings, because of
their size and rigidity, has a significant impact on the mechanical
packaging techniques. Folding patterns were developed that kept
the thin membrane material from the vicinity of the end fittings.
The results of the mechanical packaging tests of the canopy, torus,
and struts individually were used for arriving at the packaging
techniques for the complete full-size inflatable structure. Validation
of this approach for stowing the inflatable structure was done
by adding MLI blankets and electrical wiring to the inflatable
model and successfully packaging it in the canister.
The design drivers for the canister include (a) providing the
load-carrying structure for all elements of the experiment, except
the equipment panel that remains with the Spartan, (b) interface
structure with the Spartan, (c) deployable panels to accommodate
ejection of the stowed inflatable antenna structure, (d) smooth
surface compartment to house the stowed inflatable structure,
(e) interface with the struts, and (f) high structural-design
margins to minimize the need for expensive structural qualification
verification testing (Table 2).
The design approach for the canister (or bus structure) employs
aluminum honeycomb panels for the basic load-carrying elements
of this structure. The justification for this selection is that
(a) it is very stiff for its weight, (b) it is available and relatively
inexpensive, c) the smooth face sheets provide the appropriate
surfaces for the stowed inflatable structure, and (d) L'Garde,
Inc., has extensive experience using this type of structure.
Table 2. Canister Subsystem Requirements
| · House and support all elements
of experiment · Basic load-carrying structure · Structural design loads from Spartan · Deployable · Provide ejector for inflatable structure · Provide interface for struts · Interface with Spartan · High design margins |
The configuration development of the canister starts with (a)
the volume required for the stowed inflatable antenna structure,
(b) the volume and mounting surfaces required for the inflation
system, electronics, and functional elements of the SAMS, (c)
the geometry of the interface with the Spartan and the limitations
imposed by the dynamic envelope of the STS, (d) mechanization
to accommodate release of the stowed inflatable structure in a
controlled manner, (e) providing the mounting points for the three
strut structures, and (f) taking advantage of the weight available
for the experiment to develop large, structural design margins
and a structure with a first natural mode above 35 Hz to preclude
a requirement for a modal survey test.
These requirements resulted in a canister structure that is 2.0
meters long, 1.1 meters wide, and 0.46 meters high. Four spring-loaded
doors open to allow deployment of the inflatable antenna structure.
Part of the SAMS light panel is attached to the inside surface
of the top cover, and the three side doors contain "pods"
for storage of the inflatable struts. All doors are held in place
by pin-puller latches. A large spring loaded plate is located
in the floor of the canister for purposes of ejecting the inflatable
structure at the beginning of the deployment sequence. This plate
also supports the rest of the light panels.
The structural design of the canister was based on quasi-static
acceleration loads defined at the Spartan center of mass. A standard
finite-element code for the determination of panel stresses and
loads was used with hand analyses for fittings and mechanisms
stress, and margins. Because of (a) the complexity of the honeycomb
construction for modeling, (b) the geometry of the canister, (c)
the mechanisms tie points needed to accommodate articulation of
the panels, (d) and an ejection plate for pushing the stowed inflatable
structure away from the canister, the final model was 23,000 d.o.f.
This model, after experimental verification of the fundamental
structural modes below 50 Hz, was used with the Spartan analytical
model for the coupled-loads analysis for the combined loads of
the Spartan/IAE on the shuttle.
Special functional tests for the canister included (a) panel deployment,
(b) ejection-panel spring calibration, (c) pyro/pin-puller release,
and (d) structural natural frequency identification. The full-scale
engineering model and hardware used for the tests are shown in
Figure 6. The panel tests consisted of repeated articulations
with adjustments of spring stiffness and damping of the actuator
to ensure timely and low shock deployment. The ejection-panel
tests were based on repeated deployments of a mass simulation
of the inflatable structure to accommodate evaluation and adjustment
of the spring cluster to achieve the proper ejection velocity.
Pyro pin-puller release tests were done to demonstrate a "clean"
release and functional performance of the "initial motion"
kick-off springs. Forced vibration tests of the full-scale engineering
test unit, which simulate the full-up canister system, were conducted
at the Goddard Space Flight Center (GSFC) to determine all the
significant structural modes below 50 Hz which turned out to be
35 Hz. The results were then used for test/analysis correlation
for validation of the structural model that was used for the Spartan/IAE/STS-coupled
loads analysis. A highly non linear dynamic response of the stowed
inflatable structure resulted in the need to analytically account
for the change in the response frequency of two significant modes
in the linear analytical model. This was necessary to account
for the actual frequency shift of several Hz in the structure
when under high-level dynamic loading, since simulation of the
non-linear characteristics is not practical.
Surface Accuracy Measurement System
The design drivers for the surface measurement subsystem include (a) remote measurement of the reflector surface on orbit and in the presence of near direct sunlight with a resolution of +0.1 mm rms, (b) coverage of at least 90% of the surface, (c) a measurement cycle of no more than 40 seconds, and (d) a development and flight hardware cost of under $1M (Table 3). A number of systems were identified for possible application to IAE. However, only one system as even close to being affordable for the IAE. That system is based on a Digital Imaging Radiometer (DIR) developed by McDonnell Douglas for measurement of slope errors on ground-based solar concentrators. 6, 7, 8 The concept is based on using a number of discrete light sources, located near the center of curvature of a surface, and then photographing the reflected rays. Surface deviations from a perfect surface will result in shading patterns as seen by the camera. The magnitude and distribution of such shading patterns are used to determine the slope error distribution of the antenna aperture. The components needed to implement this approach include (a) a number of discrete light sources mounted on panels, (b) high resolution video cameras, (c) video records, and (d) electronic circuits to sequence and control the triggering of a large number of short interval light bursts.
Table 3. SAMS Subsystem Requirements
|
· Remote measurement of reflector surface · Concept based on MDAC DIR · Measurement data recorded on VCR · Sample data transmitted to STS · Measurement accuracy +0.1 mm rms · Surface Coverage > 90 percent · Surface measurement cycle < seconds · Development and flight hardware under $1M |
The initial step of the design was to analytically characterize
the system performance parametrically, with the resulting information
used to determine (a) the number of light sources. (b) the light
panel size an i shape, (c) the characteristics of the light sources,
such as wavelength and luminous intensity, (d) the camera characteristics
that include pixel resolution, flux sensitivity, band pass filter
and its center, and (e) the electrical and software requirements
for operating the cameras and timing for the light panels.
The next phase of system development involved identification of
flight-qualified components for the system. The biggest challenge,
of course, was the video cameras. Fortunately, a Videospection
CCD camera that had been previously flown on the STS was identified
and found to meet the functional performance requirements (Figure
7). The next challenge was to identify light sources (such as
high-output LEDs) that were available in large numbers at low
cost, since the final design utilizes 512 clusters of 36 LEDs
(Figure 7). A relatively new product by Rohm CORP. was located,
but was not flight qualified. However, since the flight system
will operate with up to 20% LED failure, the usual flight certification
was not required. The mounting of the individual LEDs on the supporting
panels utilized standard techniques for mounting electronic components
on printed wiring boards.
The special tests required for development of the flight hardware
included (a) reflectivity of the aluminized mylar to LED illumination,
(b) camera aperture evaluation and calibration, (c) light-source
intensity evaluation, (d) system characterization, using a scale-size
calibration mirror, and (e) full-scale calibration, using a 3-meter-inflatable
section of the full-size reflector structure. The reflectivity
of the aluminized mylar was obtained by Mounting the actual aluminized
mylar membrane on flat plates, illuminating the reflection surface
with the flight-type LEDs, and measuring the return signal. This
reflectivity data was used in the design of a 0.30 meter-diameter
glass mirror to be used for subsequent system calibration. The
camera-aperture calibration was accomplished by tests in direct
sunlight and in the dark with variable LED intensity. The light
source intensity was determined by direct measurement as a function
of applied voltage. The complete system was evaluated for the
first time by using the 0.30 meter diameter mirror, which had
the same reflectivity as that of the membrane and a 28-meter radius
of curvature. This test established the required flux density
of the return signal, the required camera-aperture opening, and
the adequacy of the LED's output. The full-scale test, using a
3-meter section of the 14-meter reflector offered the first opportunity
to evaluate the system in a realistic manner. By rotating the
camera so that its field or view moves across the surface of the
inflatable reflector, a simulation of an on-orbit measurement
has been achieved. The results of this test demonstrated that
(a) the measurement resolution of the system was on the order
of 0.1 to 0.2 mm rms, (b) the final camera aperture settings were
established, (c) the time required for multiple measurements was
verified, and (d) the final configuration for the light-panel
performance was determined.
The demonstrated performance of the surface measurement system
effectively satisfied the subsystem design requirements. In fact,
a number of performance results exceeded expectation. The total
development and hardware cost was under $1M; measurements were
successfully made in near direct sunlight; and the flight qualified
cameras were more than adequate for the system.
Inflation Subsystem
The key design drivers for the inflation subsystem included
(a) high-pressure nitrogen gas storage for the inflatable structure,
(b) sensors, valves, and regulators for implementing the control
of inflation, (c) using a functional concept based on previous
successful L'Garde, Inc., designs, and (d) maximizing the use
of Spartan cold-gas attitude control-system components (Table
4).
Table 4. Inflation Subsystem Requirements
|
· Active pressure control system · Provide pressure vessels, regulators, sensors, and valves to supply 52930 cc N2 at 20.68 MPa · Control inflation pressure to the required level - Deployment, design flow rate ±3 % · Maximize use of Spartan components · Maximize use of previous successful L'Garde, Inc., designs |
The functional design of the subsystem is nearly identical
in concept to the ones successfully flown by L'Garde, Inc., for
much smaller inflatable structures. Analysis of mass flow was
used to establish component requirements. Component selection
was based on previously qualified hardware used for the Spartan
attitude-control, cold-gas system and on previous L'Garde, Inc.,
flight systems. The supporting structure used for mounting the
tanks, plumbing, and components is an aluminum honeycomb panel
similar to that used for the canister. This panel also contains
the electronics, control boxes, and SAMS video cameras and is
permanently attached to the Spartan for return to earth after
completion of the experiment. The two large structural composite
gas tanks utilize the same mounting configuration as that for
Spartan to minimize requalification costs. The component mounting
and tubing are similar to previously used designs by L'Garde,
Inc. (Figure 8). Design factors of 2.5 over operating pressure
were used. No attempt was made to develop a light-weight, highly
compact inflation system for this experiment because of cost limitations.
Specialized testing included leak, proof pressure, and functional
performance validation. Simulation of orbital inflation was done
by using two large tanks which were evacuated and then filled
with gas using the sensing and gas-flow control techniques proposed
for the experiment. Requirements of the design were based on these
test results.
The inflation subsystem development was relatively straightforward,
and its functional performance easily satisfied the subsystem
requirements.
Electronic Subsystem
The design driver for the electronic subsystem is the initiation,
sequencing, and control of all IAE functions that include (a)
pyrotechnic release devices, (b) pyrotechnic valves, (c) synchronization/control
of the video cameras, VCRs, and light panels, (d) multiplexing
of engineering data, (e) logic and control of the inflatable pressures,
and (f) interface with the Spartan (Table 5). The electronics
are designed around the Intel 87C196K-MOS processor.
The subsystem design was based on conventional electrical-circuit
analysis and hardware packaging techniques successfully used on
previous L'Garde, Inc.,
flight hardware. The electrical components were selected from
JPL parts lists that identify sources for high-quality and reliable
hardware. The components are mounted on four-layer circuit boards
using standard approaches. The circuit boards are integrated by
insertion into the motherboard (Figure 9). The completed motherboard/circuit
board assembly is housed in an aluminum enclosure which is mounted
on the equipment assembly structure. Cabling used in the electronic
subsystem utilizes standard nickel-plated-type connectors.
Functional testing were based on component evaluation, circuit
characterization, assembly performance, and overall subsystem
capability.
Table 5. Electronic Subsystem Requirements
| · Experiment sequence and timing · Deployment control and inflation control · Instrumentation control · SAMS control · Health and status monitoring · MIL components to maximum extent possible |
Environmental Interactions
The experiment structure/environmental interactions include
the effects of (a) atmospheric drag on the stability of the structure,
(b) orbital thermal environment and atomic oxygen on the thin-film
materials, (c) UV radiation on the thin-film materials, (d) low-orbit
radiation on electronic components and, (e) impact of space debris
on the structure.
The atmospheric drag on the l4-meter solid surface antenna reflector
results in a significant relative separation of the Spartan/IAE
from the Orbiter. This separation (without STS station keeping)
is 8 km, 20-km, and 50 km on the first, second, and third orbit,
respectively. Consequently, the experiment was designed to be
completed during just one orbit to minimize STS propellant. Additionally,
analysis has shown that these same drag forces tend to stabilize
the deployed structure with its longitudinal axis parallel to
the ram direction. This means that minimal control authority is
required.
Consideration of the orbital thermal environment is included in
the design of the inflatable support structure by the addition
of MLI blankets to the torus and struts to maintain dimensional
stability. However, the lenticular structure, which consists of
the transparent canopy and the aluminized mylar reflector, was
not treated in any way for the control of temperature during the
one-orbit experiment. Consequently, the internal pressure will
decrease significantly when the structure falls in the earth's
shadow; but the inflation system will compensate for this. However,
for a long-term application of this concept, different and improved
flexible materials and thermal control coatings will be employed
t() maintain dimensional stability of the reflector.
Atomic oxygen is not considered a problem for this experiment
because of its one-orbit duration. The estimated degradation of
the reflector membrane for one orbit is 7.62 x 10-5 mm. However,
for real applications, special materials and/or surface coating
will have to be utilized.
The effects of UV radiation on the thin-film materials and the
South Atlantic Anomaly on the IAE electronics were addressed.
There are a number of candidate membrane materials with real potential
for being radiation resistant, as compared to those used for the
experiments. For long-term applications, such materials would
have to be used. Cost constraints for the experiment precluded
the use radiation-hardened electronic components. Therefore, to
significantly lower the probability of a single-event upset in
the 296 km and 39° inclination orbit, STS operations specify
that the Spartan/IAE will not be put into orbit near the South
Atlantic Anomaly.
The impact of the IAE colliding with sizeable ;pace debris is
not considered a problem for the experiment because of adequate
make-up gas. Even in a long-term application, this is still not
a major problem, since the torus and struts would be rigidized,
and the operating pressure in the lenticular structure would be
two orders of magnitude below that of the IAE, requiring only
a small amount of make-up gas.
Conclusions
Significant accomplishments at this time include (a) the fabrication
of a large, flight-quality-deployable space structure for under
$1?vI, (b) demonstration of mechanical packaging of a 14- by 28-meter
space structure in a container the size of a large office desk,
(c) manufacture of a 14-meter-diameter reflector membrane that
has a surface precision on the order of I to 2 mm rms, (d) development
of a space-qualified, surface-accuracy measurement system that
operates in the presence of near direct sunlight for well under
$1M, and (e) the experimental determination of the torus/canopy
interface on a large, inflatable torus by simulating 0-g in a
full-scale, neutral-buoyancy trough.
The remaining experiment objectives to be accomplished on orbit
include (a) validation and characterization of the deployment
sequence, (b) determination of the reflector-surface precision
and its thermal stability in a realistic operational environment.
Acknowledgments
The research described in this paper was carried out at the Jet Propulsion Laboratory, California Institute of Technology, under a contract with the National Aeronautics and Space Administration. Development of the flight experiment is being done by L'Garde, Inc., under contract to JPL.
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