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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.

Loop Heat Pipe (LHP) technology has advanced to the point that LHPs are baselined for thermal control systems in spacecraft applications. Many of the applications also require redundant systems to address reliability concerns. In the redundant design, two LHPs are plumbed in parallel to the same heat source and sink. The LHPs are totally separate, and each is designed to fully accommodate the total heat load at the source if the other LHP should fail. Due to the self-regulating nature of an LHP, questions have been raised regarding the expected behavior of two LHPs operating in parallel between the same source and sink, particularly their ability to self-start and equally share the heat load. To demonstrate the application of LHPs in a redundant system, two totally independent LHPs, each with the same condenser plate and heat source, were fabricated and tested.

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.

This paper describes the flight test results of the fifth generation cryogenic capillary pumped loop (CCPL-5) which flew on the Space Shuttle STS-95 in October of 1998 as part of the CRYOTSU Flight Experiment. This flight was the first in-space demonstration of the CCPL, a lightweight heat transport and thermal switching device for future integrated cryogenic bus systems. The CCPL-5 utilized nitrogen as the working fluid and operated between 75K and 110K. Flight results indicated excellent performance of the CCPL-5 in a micro-gravity environment. The CCPL could start from a supercritical condition in all tests, and the reservoir set point temperature controlled the loop operating temperature regardless of changes in the heat load and/or the sink temperature. In addition, the loop demonstrated successful operation with heat loads ranging from 0.5W to 3W, as well as with parasitic heat loads alone.

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.

In typical loop heat pipe (LHP) applications, the LHP design calls for a dedicated evaporator and a dedicated condenser. Applications exist for reversible loop heat pipes (LHPs), which can transport heat in either direction. In the reversible LHP design, two evaporator pumps are plumbed together, one which acts as an evaporator while the other acts as a condenser. The two pumps can reverse roles, simply by reversing the temperature gradient across the loop. Thus, either pump can be used as an evaporator or a condenser, depending upon the environment. Reversible LHPs can be used to share heat between components, or to cross-strap opposing spacecraft radiators. A reversible LHP was built and tested to demonstrate feasibility and to characterize its performance capabilities and attributes. The device was tested by either alternately heating each evaporator electrically or by inducing a temperature difference between the two ends of the device.

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. Other investigators have proposed systems containing several evaporators that are coupled to a common reservoir; however, this approach places severe restrictions on the relative locations of the evaporators. This paper describes a multiple evaporator loop heat pipe, with a separate reservoir for each evaporator, which can accommodate payloads with large or spatially separated heat sources.

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.

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.

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.

This paper describes the development, operation and testing of an across-gimbal ambient thermal transport system (GATTS) for carrying cryocooler waste heat across a 2-axis gimbal. The principal application for the system is space-based remote sensing spacecraft with gimbaled cryogenics optics and/or infrared sensors. GATTS uses loop heat pipe (LHP) technology with ammonia as the working fluid and small diameter stainless steel tubing to transport 100–275 W across a two-axis gimbal. The tubing is coiled around each gimbal axis to provide flexibility (less than 0.68 N-m [6 lbf-in] of tubing-induced torque per axis) and fatigue life. Stepper motors are implemented to conduct life cycling and to assess the impact of motion on thermal performance. An LHP conductance of approximately 7.5 W/C was demonstrated at 200 W, with and without gimbal motion. At the time this paper was written, the gimbal had successfully completed over 500,000 cycles of operation with no performance degradation.

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.