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The Wide Field Camera 3 (WFC3) instrument has undergone two complete thermal vacuum tests (TV1 and TV2), during which valuable lessons were learned regarding test configuration, test execution, model capabilities, and modeling practices. The very complex thermal design of WFC3 produced a number of challenging aspects to ground testing with numerous ThermoElectric Coolers and heat pipes, not all of which were functional. Lessons learned during TV1 resulted in significant upgrades to the model capabilities and a change in the test environment approach for TV2. These upgrades proved invaluable during TV2 when pre-test modeling assumptions proved to be false. Each of the lessons learned relate to one of two following broad statements: 1. Ensure the design can be tested and that the effect of non-flight like conditions is well understood, particularly with respect to non passive devices (TECs, Heat Pipes, etc) 2.

The following is the first in a series of white papers which will be issued as a result of a task to define and qualify space grade fiber optic cable assemblies. Though to qualify and use a fiber optic cable in space requires treatment of the cable assembly as a system, it is very important to understand the design and behavior of its parts. These papers will address that need, providing information and “lessons learned” that are being collected in the process of procuring, testing and specifying the final assemblies. This installment covers information on optical fiber, coatings, cable components, design guidelines and limitations, radiation and reliability.

The NASA-wide CEV Smart Buyer Team (SBT) was assembled in January 2006 and was tasked with the development of a NASA in-house design for the CEV Crew Module (CM), Service Module (SM), and Launch Abort System (LAS). This effort drew upon over 250 engineers from all of the 10 NASA Centers. In 6 weeks, this in-house design was developed. The Thermal Systems Team was responsible for the definition of the active and passive design architecture. The SBT effort for Thermal Systems can be best characterized as a design architecting activity. Proof-of-concepts were assessed through system-level trade studies and analyses using simplified modeling. This nimble design approach permitted definition of a point design and assessing its design robustness in a timely fashion. This paper will describe the architecting process and present trade studies and proposed thermal designs

Spacecraft radiators reject heat to their surroundings. Radiators can be deployable or mounted on the body of the spacecraft. NASA's Crew Exploration Vehicle is to use body mounted radiators. Coatings play an important role in heat rejection. The coatings provide the radiator surface with the desired optical properties of low solar absorptance and high infrared emittance. These specialized surfaces are applied to the radiator panel in a number of ways, including conventional spraying, plasma spraying, or as an appliqué. Not specifically designed for a weathering environment, little is known about the durability of conventional paints, coatings, and appliqués upon exposure to weathering and subsequent exposure to solar wind and ultraviolet radiation exposure. In addition to maintaining their desired optical properties, the coatings must also continue to adhere to the underlying radiator panel.

On Day 97, 2005, a temperature excursion of the Burst Alert Telescope (BAT) loop heat pipe (LHP) #1 compensation chamber (CC) caused this LHP shut down. It had no impact on the Gamma Ray Burst (GRB) detection because LHP #0 was nominal. After LHP #1 was started up and its primary heat controller was disabled on Day 98, both LHPs have been nominal. On Day 337, 2004, the X-Ray Telescope (XRT) thermo-electric cooler (TEC) power supply (PS) suffered a single point failure. The charge-coupled device (CCD) has been cooled by the radiator passively to -50°C or colder most of the time. The CCD temperature meets the main objective of pinpointing GRB afterglow positions. With these anomalies overcome, the Instrument Module (IM) thermal control system (TCS) is nominal during the first 2.5 years in flight.

The system of diffusion equations generated by SINDA is solved by an eigenmethod. This avoids the current numerical integration and its instability problems. The nonlinear terms such as radiative coupling are treated as forcing functions, and their rates of change determine the time steps, which are automatically set to yield any desired accuracy. The method can be significantly faster than the current methods. An example problem is solved in sufficient detail to allow a programmer to install the method in SINDA, so that its use would be just as convenient as calling for the Crank-Nicholson method. Existing SINDA models could be run without change, and the modelling advantages of SINDA are retained.

New Space Support Equipment (SSE) components developed for the Hubble Space Telescope Second Servicing Mission are described, with particular emphasis on how flight experience from the 1993 First Servicing Mission was utilized in the design and testing process. The new components include the Second Axial Carrier (SAC) Axial Scientific Instrument Protective Enclosure (ASIPE), the magnetic-damped SAC ASIPE Load Isolation System, the Enhanced Power Distribution and Switching Unit (EPDSU), and the Multi-Mission Orbital Replacement Unit Protective Enclosure (MOPE). Analytical modeling predictions are compared with on-orbit data from the Hubble Space Telescope (HST) Second Servicing Mission. Those involved in thermal designs of hardware for use on the Shuttle or Space Station, particularly with astronaut interaction, may find interest in this paper.

The success of CAPL 2 flight experiment has stirred many interests in using capillary pumped loop (CPL) devices for spacecraft thermal control. With only one evaporator in the loop, CAPL 2 was considered a point design for the Earth Observing System (EOS-AM). To realize the full benefits of CPLs, a reliable system with multiple evaporators must be developed and successfully demonstrated in space. The Capillary Pumped Loop (CAPL 3) Flight Experiment was designed to flight demonstrate a multiple evaporator CPL in a space environment. New hardware and concepts were developed for CAPL 3 to enable reliable start-up, constant conductance operation, and heat load sharing. A rigorous ground test program was developed and extensive characterization tests were conducted. All performance requirements were met, and the loop demonstrated very reliable operation.

The operating temperature of a loop heat pipe (LHP) with a single evaporator is governed by the compensation chamber (CC) temperature, which in turn is a function of the evaporator power, condenser sink temperature, and ambient temperature. As the operating condition changes, the CC temperature will also change during the transient but eventually reach a new steady temperature. Under certain conditions, however, the LHP never really reaches a true steady state, but instead displays an oscillatory behavior. This paper presents a study on the oscillation of the loop operating temperature with amplitudes on the order of 1 Kelvin and periods on the order seconds to minutes. The source of the high frequency temperature oscillation is the fast movement of the vapor front in the condenser section, which usually occurs when the vapor front is near the condenser inlet or the condenser outlet.

The Geoscience Laser Altimeter System (GLAS) instrument is NASA Goddard Space Flight Center's first application of Loop Heat Pipe technology that provides selectable/stable temperature levels for the lasers and other electronics over a widely varying mission environment. GLAS was successfully launched as the sole science instrument aboard the Ice, Clouds, and Land Elevation Satellite (ICESat) from Vandenberg AFB at 4:45pm PST on January 12, 2003. After SC commissioning, the LHPs started easily and have provided selectable and stable temperatures for the lasers and other electronics. This paper discusses the thermal development background and testing, along with details of early flight thermal performance data.

NASA’s Hubble Space Telescope was designed for periodic servicing by Space Shuttle astronauts performing extravehicular activities (EVAs), to service, maintain, repair, and upgrade the telescope. Through three successful servicing missions to date, EVA processes have been developed by applying a series of important lessons learned. These lessons learned are also applicable to many other future human spaceflight and robotic missions, such as International Space Station, satellite retrieval and servicing, and long-duration spaceflight. HST has become NASA’s pathfinder for observatories, EVA development, and EVA mission execution.

This paper presents test results of an experimental study of low power operation in a loop heat pipe. The main objective was to demonstrate how changes in the vapor void fraction inside the evaporator core would affect the loop behavior. The fluid inventory and the relative tilt between the evaporator and the compensation chamber were varied so as to create different vapor void fractions in the evaporator core. The effect on the loop start-up, operating temperature, and capillary limit was investigated. Test results indicate that the vapor void fraction inside the evaporator core is the single most important factor in determining the loop operation at low powers.

The Hubble Space Telescope (HST) was launched in 1990 and has undergone several Servicing Missions that have replaced and repaired various scientific and support hardware. As preparations begin for Servicing Mission Four (SM4) in 2008 and the life extension activities that follow, the Telescope Thermal Math Model (TMM) has been improved using the latest thermal analysis software and techniques. Several efforts have been made to improve the HST system-level TMM since launch. A brief history of the major model updates, as well as the motivation behind the changes has been provided. The current improvements have provided the HST systems-level TMM a greater level of detail, while making model control more user-friendly and the results easier to verify. Several modeling techniques useful for spacecraft thermal design and operations support are discussed.

Following the satellite-level thermal vacuum test for the Cloud-Aerosol Lidar and Infrared Pathfinder Satellite Observation spacecraft, project thermal engineers determined that the radiator used to cool the Integrated Lidar Transmitter subsystem during its operation was oversized. In addition, the thermal team also determined that the operational heaters were undersized, thus creating two related problems. Without the benefit of an additional thermal vacuum test, it was necessary to develop and prove by analysis a laser temperature control scheme using the available resources within the spacecraft along with proper resizing of the radiator. A resizing methodology and new laser temperature control scheme were devised that allowed, with a minimum of 20% heater power margin, the operating laser to maintain temperature at the preferable set point. This control scheme provided a solution to a critical project problem.

A new, general-purpose computer program (BIFAC) has been developed for computing thermal radiant interchange among opaque surfaces that need not be perfectly diffuse or perfectly specular. The method uses the full bi-directional reflectance distribution function (BRDF) to determine directional radiosities, and thence heat fluxes, between surfaces. The method gives more accurate average interchange factors for diffuse surfaces, because it better represents interaction in corners. The maximum error in a stringent test using a specular surface was 8.9%, in great part because the exact specular solution does not include the real specular cone that is used in BIFAC.

This paper explores the possibilities of cooling a permanently inhabited lunar base with a reverse Brayton cycle Thermal Control System (TCS). Based on an initial stage outpost, the cooling needs are defined. A thermodynamic performance model for the Brayton cycle is derived using ideal gas analysis. This model includes inefficiencies and irreversibilities of the components. The free parameters in the thermodynamic model are successively removed using limiting values for efficiencies and determining operating parameters by suboptimizations. In essence a model for cooling efficiency as a function of rejection temperature alone is obtained. For every component of the system a mass model is applied and the overall mass is determined. The last remaining degree of freedom, the rejection temperature, is eliminated by an optimization for lowest overall mass. The result for minimal TCS mass is compared to a reference TCS using a Rankine cycle.

The Capillary Pumped Loop Flight Experiment (CAPL) is a prototype of the Earth Observing System (EOS) instrument thermal control systems, which are based on two-phase heat transfer technology. The CAPL experiment has been functionally tested in a thermal vacuum chamber at NASA's Goddard Space Flight Center (GSFC). The tests performed included start-up tests, simulated EOS instrument power profiles, low and high power profiles, a variety of uneven coldplate heating tests, subcooling requirement tests, an induced deprime test, reprimes, saturation temperature changes, and a hybrid (mechanical pump-assist) test. There were a few unexpected evaporator deprimes, but overall the testing was successful. The results of all of the tests are discussed, with emphasis on the deprimes and suspected causes.

The Capillary Pumped Loop Flight Experiment (CAPL) employs a passive two-phase thermal control system that uses the latent heat of vaporization of ammonia to transfer heat over long distances. CAPL was designed as a prototype of the Earth Observing System (EOS) instrument thermal control systems. The purpose of the mission was to provide validation of the system performance in micro-gravity, prior to implementation on EOS. CAPL was flown on STS-60 in February, 1994, with some unexpected results related to gravitational effects on two-phase systems. Flight test results and post flight investigations will be addressed, along with a brief description of the experiment design.

An IBM Personal Computer (PC) version of the Groove Analysis Program (GAP) was developed to predict the steady state heat transport capability of axially grooved heat pipes for a specified groove geometry and working fluid. In the model, the heat transport capability of an axially grooved heat pipe, usually governed by the capillary limit, is determined by the numerical solution of the governing equation for momentum conservation with the appropriate boundary conditions. This paper discusses the theory behind the development of the GAP model. It also presents many useful capabilities of the model. Furthermore, correlations of flight test performance data using GAP are presented and discussed.

The Hubble Space Telescope (HST) was designed to be serviced from the shuttle by astronauts performing extravehicular activities (EVA). During the first HST Servicing Mission (STS-61) two types of power tools were flown, the Power Ratchet Tool (PRT) and the HST Power Tool. Each tool had both benefits and drawbacks. An objective for the second HST servicing mission was to combine the reliability, accuracy, and programmability of the PRT with the pistol grip ergonomics and compactness of the HST Power Tool into a new tool called the EVA Pistol Grip Tool (PGT). The PGT is a self-contained, microprocessor controlled, battery powered, 3/8-inch drive hand-held tool. The PGT may also be used as a non-powered ratchet wrench. Numerous torque, speed, and turn or angle limits can be programmed into the PGT for use during various servicing missions. Batteries Modules are replaceable during ground, Intravehicular Activities (IVA), and EVA operations.