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Using Loop Heat Pipes (LHPs) for controlling the temperature of the source of heat has been considered for many applications. However, traditional LHPs can require significant amounts of power for source temperature control. A number of techniques have been identified and implemented to reduce control power requirements. One of the very first design approaches was to thermally couple the liquid line bringing subcooled liquid from the condenser to the vapor line entering the condenser with a number of “coupling blocks”. In another application, a variable conductance heat pipe (VCHP) was used to couple the liquid line to the LHP evaporator. A third generation approach has been developed that offers even further reductions in control power. The paper discusses earlier generations of control power reduction approaches with their advantages and disadvantages. It also describes the third generation of the approach, which is currently in manufacturing.

Future spacesuit systems for the exploration of Mars will need to be much lighter than current designs, while at the same time reducing the consumption of water for crew cooling. One of the technology paths NASA has identified to achieve these objectives is the replacement of current high pressure oxygen storage technology in extravehicular activity (EVA) systems with cryogenic technology that can simultaneously reduce the mass of tankage required for oxygen storage and enable the use of the stored oxygen as a means of cooling the EVA astronaut. During the past year NASA has funded production of a prototype system demonstrating this capability in a design that will allow the cryogenic oxygen to be used in any attitude and gravity environment. This paper describes the design and manufacture of the prototype system. The potential significance and application of the system is also discussed.

New goals set for the U.S. space program focus on reestablishing the human presence on the Moon and sending the first manned mission to Mars in the beginning of the 21st century. The necessary first step in the support of these goals is identifying requirements that drive the development of new technologies. Since extravehicular activity (EVA) will be an integral part of the establishment of both a lunar base and the exploration of the martian surface, this is an area where specific requirements for the EVA systems need to be determined. EVA on the lunar and martian surfaces presents unique conditions in which an extravehicular mobility unit (EMU) must operate. These conditions include environmental factors such as partial gravity, dust, thermal gradients, atmospheric conditions, lighting, and radiation.

This paper introduces the concepts utilized for the integration of a cryogenic capillary pumped loop into a flight experiment. The Cryogenic Capillary Pumped Loop (CCPL) version V, which was recently manufactured (9/97), is to be integrated into the Cryogenic Thermal Storage Unit (CRYOTSU) flight experiment as a secondary experiment. CRYOTSU, a Get-Away-Special (GAS) Can experiment, is currently manifested on STS-95 with an anticipated launch date of October 1998. The CCPL uses nitrogen as the working fluid with a 70-120 K operating temperature. The primary benefit of the CCPL is as a heat transport device in cryogenic bus systems. The primary issue of structurally supporting the CCPL while reducing parasitic heat loads will be detailed.

The ADRAD deployable radiator is in development at Swales Aerospace to provide additional heat rejection area for spacecraft without envelope impact. The ADRAD design incorporates ALPHA loop heat pipes, an aluminum honeycomb radiator with embedded condenser, OSR optical coating, spherical bearing hinges, pyrotechnic release devices and snubbers. This paper describes the design of ADRAD to a set of “generic” GEO requirements, including a nominal heat rejection capacity of 1250 W. Thermal, structural and mechanism considerations are described along with the comprehensive systems approach necessary to produce an integrated subsystem.

The Advanced Lightweight Radiator Development Unit (ALRDU) is a composite radiator panel built for the Advanced Thermal Exploration Technology Development Project. As a point of reference, the ALRDU is sized to meet the requirements captured during an early design iteration of the Crew Exploration Vehicle (CEV). The composite radiator panel development included an evaluation of various tubing and composite panel attachment methods previously performed at NASA Glenn Research Center (GRC). The resulting lessons learned combined with previous composite radiator experiences were used to support the ALRDU design. The aforementioned experience was used to determine key CEV radiator design requirements used by Ball Aerospace and XCA for the ALRDU development. Testing of the ALRDU was performed at NASA Johnson Space Center (JSC) in an integrated thermal vacuum system test.

The Earth Observing-1 spacecraft, built by Swales Aerospace for NASA's Goddard Space Flight Center (GSFC), was successfully launched on a Boeing Delta-II ELV on November 21, 2000. The EO-1 spacecraft thermal design is a cold bias design using passive radiators, regulated conductive paths, thermal coatings, louvers, thermostatically controlled heaters and multi-layer insulating (MLI) blankets. Five of the six passive radiators were aluminum honeycomb panels. The sixth panel was a technology demonstration referred to as the Carbon Carbon Radiator (CCR) panel. Carbon-Carbon (C-C) is a special class of composite materials in which both the reinforcing fibers and matrix materials are made of pure carbon. The use of high conductivity fibers in C-C fabrication yields composite materials that have high stiffness and high thermal conductivity.

A small SINDA/FLUINT logic routine was developed to improve upon standard spacecraft-to-instrument thermal model interface methodology for steady state analysis. Rather than the standard approach of providing backloads and/or conductive limits with uniform spacecraft temperatures, this methodology enables the instrument thermal engineer to make more informed design decisions by providing more information regarding the source and magnitude of the sink temperatures and backloads. The instrument thermal engineer can use the model information provided from the spacecraft thermal engineer to make more informed design decisions in subsequent analysis, and can be less dependent on the spacecraft thermal engineer.

The Spacecraft Radiative Thermal Model Exchange System is a technology developed for the bi-directional exchange of spacecraft radiative thermal models via the TMG thermal software package. It provides a means for quickly and accurately transferring models between TMG and theree of the major thermal radiation codes used in the spacecraft industry, particularly the ESARAD and Thermica packages, which are widely used by contractors to the European Space Agency, and the TSS code which is prevalent in the United States space industry. In order to reconcile element-based and primitives-based modeling approaches, this system includes an interactive primitives-based modeling system, enabling users to construct, import, and manipulate primitives-based radiation models in TMG.

A Small Loop Heat Pipe (SLHP) featuring a wick of only 1.27 cm (0.5 inches) in diameter has been designed for use in spacecraft thermal control. It has several features to accommodate a wide range of environmental conditions in both operating and non-operating states. These include flexible transport lines to facilitate hardware integration, a radiator capable of sustaining over 100 freeze-thaw cycles using ammonia as a working fluid and a structural integrity to sustain acceleration loads up to 30 g. The small LHP has a maximum heat transport capacity of 120 Watts with thermal conductance ranging from 17 to 21 W/°C. The design incorporates heaters on the compensation chamber to modulate the heat transport from full-on to full-stop conditions. A set of start up heaters are attached to the evaporator body using a specially designed fin to assist the LHP in starting up when it is connected to a large thermal mass.

This paper describes the development and testing status of several novel components and integration tools for space-based cryogenic applications. These advanced devices offer functionality in the areas of cryogenic thermal switching, cryogenic thermal transport, cryogenic thermal storage, and cryogenic integration. As such, they help solve problems associated with cryocooler redundancy, across-gimbal thermal transport, large focal plane array cooling, fluid-based cryogenic transport, and low vibration thermal links. The devices discussed in the paper include a differential thermal expansion cryogenic thermal switch, an across-gimbal thermal transport system, a cryogenic loop heat pipe, a cryogenic capillary pumped loop, a beryllium cryogenic thermal storage unit, a high performance flexible conductive link, a kevlar cable structural support system, and a high conductance make-break cryogenic thermal interface.

Like a Loop Heat Pipe (LHP), a Cryogenic Loop Heat Pipe (CLHP) is a passive two-phase heat transport system that utilizes the capillary pressure developed in a fine pore evaporator wick to circulate the system's working fluid. To demonstrate startup from a supercritical temperature and an operation below ambient temperature for passive bench cooling applications, a CLHP was developed and tested in a thermal vacuum chamber. The system requires startup from a maximum outgassing temperature of 335K over an operating temperature range of 215 to 218K, and an orbital average heat transport capability of 39W. Ethane was selected as the working fluid because it has heat transport properties that are suitable for the operating temperature of 218K. This paper provides a description of the CLHP concept, the development of the design including proof of concept development and testing of a CLHP designed to provide passive cooling of optical instruments.

NASA's current vision for exploration dictates that radiators for a Crew Exploration Vehicle (CEV), a Lunar Surface Access Module (LSAM), and a lunar base will need to be developed. These applications present new challenges when compared to previous radiators on the Space Shuttle and International Space Station (ISS). In addition, many technological advances have been made that could positively impact future radiator design. This paper outlines new requirements for future radiators and documents a trade study performed to select some promising technologies for further evaluation. These technologies include carbon composites substrates as well as Optical Solar Reflectors (OSRs), a lithium based white paint, and electrochromic thin films for optical coatings.

Advanced life support and habitat functions for exploration missions will require the autonomous control of many interdependent subsystems. A controls evaluation was envisioned to help define the questions necessary to develop an architecture capable of this degree of autonomy. To conduct this evaluation, a biological test bed that consists of a pair of interdependent subsystems was developed. The biological test bed represents an analog of a life support subsystem necessary for long-duration, human-rated exploration missions. The test bed consists of a packed bed anoxic bioreactor for the removal of organic carbon, coupled with an aerobic nitrifier that converts ammonia to nitrogen.

From 1997 to 2001, the third author worked with a team of engineers at JSC to develop the requirements and basic design for the Bioregenerative Planetary Life Support Systems Test Complex, or BIO-Plex. Under the Advanced Integration Matrix (AIM) Project, this earlier effort is extended to an investigation of methods and approaches for Advanced Systems Integration and Control. The intent is to understand and validate the use of software as an integrating function for complex heterogeneous systems, particularly for Advanced Life Support (ALS), in the context of support of mission operations. Preliminary investigations undertaken in the summer of 2004 indicate that integration of controls for the type of dynamic, non-linear, closed-loop biological systems under investigation for ALS systems require a different systems engineering approach than that required for traditional avionics systems.

Having been in operation for over 15 years, the Hubble Space Telescope (HST) had experienced significant changes in both hardware upgrades and operational modes. The changes were necessary to improve performance of some equipment and to replace failed electronics in others. Hardware replacements were done in several servicing missions. To accommodate the change in physical condition of HST, alterations in the way the telescope is operated were also required. The final opportunity to make any hardware changes on HST is during Servicing Mission 4 (SM-4) which is currently scheduled for December of 2007. It is important to make the most appropriate changes in order to ensure that HST will be in good operating condition until its planned termination. In order to provide manifest input to the HST project for the final servicing mission, the HST thermal team must conduct careful evaluation of every single piece of hardware on HST.

Loop Heat Pipes (LHPs) are passive two-phase heat transport devices that have been baselined for many spacecraft thermal management applications. The design life of a spacecraft can extend to 15 years or longer, thus requiring a robust thermal management system. Based on conventional aluminum/ammonia heat pipe experience, there exists a potential for the generation of noncondensible gas in LHPs over the spacecraft lifetime. In addition, some applications would have the LHP evaporator attached directly to spacecraft equipment having large thermal mass. To address the potential issues associated with LHP operation with noncondensible gas and large thermal mass attached to the evaporator, a test program was implemented to examine the effect of mass and gas on ammonia LHP performance. Many laboratory test programs for LHPs have heat delivered to the evaporator through light-weight aluminum heater blocks.

Loop Heat Pipe (LHP) technology has advanced to the point that LHPs are baselined for thermal control systems in many spacecraft applications. These applications typically utilize a loop heat pipe with a single evaporator. However, many emerging applications involve heat sources with large thermal footprints, or multiple heat sources that would be better served by LHPs with multiple evaporators. Dual evaporator LHPs with separate reservoirs for each evaporator have been successfully developed, but the volume and weight of such systems become impractical as the number of the evaporators increase to more than three or four. Other investigators have proposed systems containing several evaporators that are coupled to a common reservoir with a conduit to contain a capillary link (secondary wick). This approach places several restrictions on the relative location of the evaporators due to the limitation of the capillary link.

The thermal design and analysis of the Earth Observing-1 (EO-1) spacecraft, built by Swales Aerospace for NASA's Goddard Space Flight Center (GSFC), consisted of a Thermal Synthesis System1 (TSS) geometric math model (GMM) and a SINDA/FLUINT2 thermal math model (TMM). These models took advantage of the submodel capability of TSS and SINDA/FLUINT providing a simplified approach for merging spacecraft and instrument models. In addition to the spacecraft thermal model, there is the Advanced Land Imager (ALI) instrument model by MIT/LL, the Hyperion instrument by TRW, the Atmospheric Corrector (AC) instrument by GSFC, and the New Millenium Program (NMP) experiments. Separate thermal models were developed for each NMP experiment which included, the Pulse Plasma Thruster (PPT) by Primex, Lightweight Flexible Solar Array (LFSA) by Lockheed, X-Band Phased Array by Boeing and the Carbon-Carbon Radiator that was developed as a joint effort between NASA and industry.

The Earth Observing-1 (EO-1) spacecraft is the first earth orbiting spacecraft in NASA's New Millennium Program. The New Millennium Program is part of the agency's Mission to Planet Earth enterprise, a series of space missions designed to enhance our knowledge of the Earth and its environmental systems. The EO-1's mission is to employ advanced remote-sensing technologies, including hyperspectral and multispectral imaging techniques, to develop highly accurate terrestrial images. In order to accomplish this mission, the spacecraft contains three primary instruments: Advanced Land Imager (ALI), Atmospheric Corrector, and Hyperion. The bus supporting these sensors is part of a 3-axis stabilized, nadir pointing spacecraft that employs an articulating solar array to provide a constant voltage, regulated power bus. EO-1 also contains several new technologies such as a carbon-carbon radiator and a pulsed plasma thruster that will be verified as part of the secondary mission objectives.