The traditional approach to satellite design is a customized and highly optimized satellite bus. The primary design
driver is to minimize mass but often at the expense of schedule and non-recurring engineering costs. The result after
years of development is a high performance system with minimal flexibility. Consequently, there is a need for
responsive, small satellites that are able to accommodate different missions, changing threats, and emerging
technologies for which the traditional development approach is unable to satisfy. Instead, systems must be modular
and/or robust. One of the subsystems that will be challenging for the development of modular and/or robust
architectures is the thermal control subsystem (TCS). To design a traditional TCS, virtually every aspect of the
mission, the satellite, and the components must be known before an intense design program can be completed.
However, the mission, payload, components, and requirements are largely unknown before mission initiation. To
provide a baseline for the TCS design and to help bound the problem for the development of robust thermal systems,
the range of external and internal heat loads for small satellites were evaluated. From this analysis, the realistic worst
design cases were identified along with other requirements for robust thermal control systems. Finally, the paper will
discuss the merits of various thermal architectures and the challenges associated with achieving the requirements for
robust thermal control for responsive satellite buses.
The performance of mission-critical components and systems within spacecraft and satellites requires the ability to
control the local thermal environment. Under conditions of relatively constant component and system loading, this
would involve radiative dissipation of both internally and externally generated heat loads and altering thermal
balances to provide heating where necessary. As the local thermal load changes with component use, the need arises
to alter the heat transfer rates and dissipation within the spacecraft. It is also desirable to be able to evaluate,
reconfigure or repair space-based thermal control systems using only ground station commands. These needs can be
met using a Plug-and-Play variable-emittance control system where operational analysis and reconfiguration is
accomplished via an improved Universal Serial Bus (USB) or space-wire controlled architecture.
This paper presents a modular, USB/space-wire-driven thermal control system using a solid-state thin-film infrared
variable-emittance device (EclipseVED<sup>TM</sup>) from Eclipse Energy Systems, Inc. The paper discusses critical issues
including connectivity, device-control scale-up for the advancement of an integrated variable-emittance system,
comparison of device weight to other variable emittance systems, the capacity to replace or repair devices in-flight,
the survivability of the system in space and the importance of individual device control.
The Air Force Research Laboratory/Space Vehicles Directorate (AFRL/RV) is developing a satellite structural
architecture in support of the Department of Defense's Operationally Responsive Space (ORS) initiative. Such a
structural architecture must enable rapid Assembly, Integration, and Test (AI&T) of the satellite, accommodate multiple
configurations (to include structural configurations, components, and payloads), and incorporate structurally integrated
thermal management and electronics, while providing sufficient strength, stiffness, and alignment accuracy. The chosen
approach will allow a wide range of satellite structures to be assembled from a relatively small set of structural
components. This paper details the efforts of AFRL, and its contractors, to develop the technology necessary to realize
The Eclipse infrared electro-chromic device (IR-ECD) is an all-solid-state monolithic vacuum deposited
thin film system functioning as an electrically controlled dimmable mirror in the IR region. The maximum
reflectance corresponding to the bleached condition of the system is around 90% (low-e condition, e=0.1).
The minimum reflectance reaches nearly zero in the colored condition of the system (high emmittance,
e=1). It is a variable emittance electro-chromic device (VE-ECD). The average emissivity modulation of
the Eclipse VE-ECD is 0.7 in the 8-12 micron region, and at 9.7 micron (room temperature) it reaches a
value of 0.9. Half and full emissivity modulations occur within 2 and 10 minutes, respectively. Because of
its low mass (5 g/m<sup>2</sup>), low voltage requirement (±1 V), extremely good emissivity control properties (from
0 to 0.9 at 300 K), and highly repeatable deposition process, the VE-ECD technology is very attractive for
satellite thermal control applications. The Eclipse VE-ECD has been under evaluation in a real space
environment since March 8, 2007. This paper presents recent developments on Eclipse's VE-ECD
including space test results.
The Department of Defense is actively pursuing a Responsive Space capability that will dramatically reduce the cost and
time associated with getting a payload into space. In order to enable that capability, our space systems must be modular
and flexible to cover a wide range of missions, configurations, duty cycles, and orbits. This places requirements on the
entire satellite infrastructure: payloads, avionics, electrical harnessing, structure, thermal management system, etc. The
Integrated Structural Systems Team at the Air Force Research Laboratory, Space Vehicles Directorate, has been tasked
with developing structural and thermal solutions that will enable a Responsive Space capability. This paper details a
"symbiotic" solution where thermal management functionality is embedded within the structure of the satellite. This
approach is based on the flight proven and structurally efficient isogrid architecture. In our rendition, the ribs serve as
fluidic passages for thermal management, and passively activated valves are used to control flow to the individual
components. As the paper will explain, our analysis has shown this design to be structurally efficient and thermally
responsive to a wide range of potential satellite missions, payloads, configurations, and orbits.
Launch vehicles produce high levels of acoustic energy and vibration loads that can severely damage satellites during
launch. Because of these high loads, the satellite structure is much more robust than it needs to be for on-orbit
operations. Traditionally, acoustic foam is used for acoustic mitigation; however, it is ineffective at frequencies below
500 Hz. For this reason we investigated three different modified acoustic foam concepts consisting of a thin metal foil, a
semi-rigid spacer, and a melamine foam substrate to improve the low frequency acoustic performance of the melamine
foam. The goal of the Hybrid Acoustically Layered Foil (HALF) Foam concept was to excite bending waves within the
plane of the foil to cause inter-particle interactions thus increasing the transmission loss of the foam. To determine the
performance of the system, a transmission loss tube was constructed, and the normal incidence transmission loss for each
sample was measured. The tests confirm the excitation of bending waves at the target frequency of 500 Hz and a
significant increase, on the order of 8 dB, in the transmission loss.
There is a critical need, not just in the Department of Defense (DOD) but the entire space industry, to reduce the development time and overall cost of satellite missions. To that end, the DOD is actively pursuing the capability to reduce the deployment time of a new system from years to weeks or even days. The goal is to provide the advantages space affords not just to the strategic planner but also to the battlefield commanders. One of the most challenging aspects of this problem is the satellite's thermal control system (TCS). Traditionally the TCS must be vigorously designed, analyzed, tested, and optimized from the ground up for every satellite mission. This "reinvention of the wheel" is costly and time intensive. The next generation satellite TCS must be modular and scalable in order to cover a wide range of applications, orbits, and mission requirements. To meet these requirements a robust thermal control system utilizing forced convection thermal switches was investigated. The problem was investigated in two separate stages. The first focused on the overall design of the bus. The second stage focused on the overarching bus architecture and the design impacts of employing a thermal switch based TCS design. For the hot case, the fan provided additional cooling to increase the heat transfer rate of the subsystem. During the cold case, the result was a significant reduction in survival heater power.
Extreme noise and vibration levels at lift-off and during ascent can damage sensitive payload components. Recently, the Air Force Research Laboratory, Space Vehicles Directorate has investigated a composite structure fabrication approach, called chamber-core, for building payload fairings. Chamber-core offers a strong, lightweight structure with inherent noise attenuation characteristics. It uses one-inch square axial tubes that are sandwiched between inner and outer face-sheets to form a cylindrical fairing structure. These hollow tubes can be used as acoustic dampers to attenuate the amplitude response of low frequency acoustic resonances within the fairing’s volume. A cylindrical, graphite-epoxy chamber-core structure was built to study noise transmission characteristics and to quantify the achievable performance improvement. The cylinder was tested in a semi-reverberant acoustics laboratory using bandlimited random noise at sound pressure levels up to 110 dB. The performance was measured using external and internal microphones. The noise reduction was computed as the ratio of the spatially averaged external response to the spatially averaged interior response. The noise reduction provided by the chamber-core cylinder was measured over three bandwidths, 20 Hz to 500 Hz, 20 Hz to 2000 Hz, and 20 Hz to 5000 Hz. For the bare cylinder with no acoustic resonators, the structure provided approximately 13 dB of attenuation over the 20 Hz to 500 Hz bandwidth. With the axial tubes acting as acoustic resonators at various frequencies over the bandwidth, the noise reduction provided by the cylinder increased to 18.2 dB, an overall increase of 4.8 dB over the bandwidth. Narrow-band reductions greater than 10 dB were observed at specific low frequency acoustic resonances. This was accomplished with virtually no added mass to the composite cylinder.