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The operating temperature of a loop heat pipe (LHP) is governed by the saturation temperature of its compensation chamber (CC); the latter is in turn determined by the balance among the heat leak from the evaporator to the CC, the amount of subcooling carried by the liquid returning to the CC, and the heat exchanged between the CC and ambient. Thus, the operating temperature of an LHP is a function of the evaporator heat input and the condenser sink temperature. The LHP operating temperature can be controlled at a desired set point by actively controlling the CC temperature. Several methods have been developed to control the CC temperature, including direct heating of the CC, coupling block, heat exchanger and separate subcooler, variable conductance heat pipe, vapor by-pass valve, secondary evaporator, and thermoelectric converter. The paper discusses the operating principles, advantages and disadvantages of each method.

A Loop Heat Pipe (LHP) is a passive two-phase heat transfer device developed and successfully employed to cool spacecraft (satellite) electronics. The intrinsic benefits of this technology (lightweight, small volume, high thermal conductance) make it an attractive potential solution to many problems in ground vehicle thermal management. As most published LHP research has focused on cooling orbiting spacecraft components, there is little knowledge of how LHPs perform under the temperature extremes (−40°C to 40°C) and diurnal/seasonal fluctuations anticipated with terrestrial applications. Ambient temperature extremes mandate consideration of transport line heat exchange with the surroundings (parasitic losses/gains). This paper presents results from an experimental investigation of liquid line return temperature impact on system performance for sink temperatures from −30°C to 40°C and evaporator loads up to 700 Watts.

This paper describes a study on the capillary limit of a loop heat pipe (LHP) with two evaporators and two condensers. Both theoretical analysis and experimental investigation are performed. Experimental tests conducted include heat load to one evaporator only, even heat loads to both evaporators, and uneven heat loads to both evaporators. Test results show that after the capillary limit is exceeded, vapor will penetrate through the wick of the weaker evaporator, and the compensation chamber (CC) of that evaporator will control the loop operating temperature regardless of which CC has been in control prior to the event. Because the evaporator can tolerate vapor bubbles, the loop can continue to work after vapor penetration. As the loop operating temperature increases, the system pressure drop actually decreases due to a decrease in liquid and vapor viscosities. Thus, the loop may reach a new steady state at a higher operating temperature after vapor penetration.

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.

The loop heat pipe (LHP) was invented in Russia in the early 1980’s. It is a two-phase heat transfer device that utilizes the evaporation and condensation of a working fluid to transfer heat, and the capillary forces developed in fine porous wicks to circulate the fluid. The LHP is known for its high pumping capability and robust operation because it uses fine-pored metal wicks and the integral evaporator/hydro-accumulator design. It has gained rapid acceptance in recent years as a thermal control device in space applications. This paper presents an overview of the LHP operation. The physical processes and the thermal-hydraulic behaviors of the LHP are first described. Operating characteristics as functions of various parameters including the heat load, sink temperature, ambient temperature, and elevation are presented. Peculiar behaviors in LHP operation such as temperature hysteresis and temperature overshoot during start-up are explained.

In this paper, results of the NRL LHP experimental studies, conducted by Naval Research Laboratory (NRL) and NASA Goddard Space Flight Center, will be presented. Emphasis in this test program is to examine the “turnkey” startup of the NRL LHP and its operational characteristics. Series of tests were performed, including startup tests, power cycling tests, low power tests, and high power tests. The NRL LHP has demonstrated very robust operations throughout the tests. In addition, hysteresis was found at low power operations. Importance of the two-phase dynamics in the evaporator core is realized, which has shown significant effects on loop operations.

The CAPL 2 Flight Experiment, flown on Space Shuttle STS-69 in 1995, was a flight demonstration of a full-scale prototype of a thermal control system planned for the Earth Observing System (EOS-AM) instruments Flight tests successfully demonstrated various CPL operations with simulated EOS-AM power profiles, including baseline and backup start-up procedures. In general, there were no significant differences in CPL performance between one-G and zero-G. However, some unusual behaviors were observed in several start-ups during the flight test. This paper describes CAPL 2 start-ups in detail, and offers explanations for the notably different zero-G behaviors.

This paper describes flight test results of the CAPL 2 Flight Experiment, which is a full scale prototype of a capillary pumped loop (CPL) heat transport system to be used for thermal control of the Earth Observing System (EOS-AM) instruments. One unique feature of CAPL 2 is its capillary starter pump cold plate design, which consists of a single capillary starter pump and two heat pipes. The starter pump enhances start-up success due to its self-priming capability, and provides the necessary capillary pumping force for the entire loop. The heat pipes provide the required isothermalization of the cold plate. Flight tests included those pertinent to specific EOS applications and those intended for verifying generic CPL operating characteristics and performance limits. Experimental results confirmed that the starter pump was indeed self-priming and the loop could be successfully started every time.

The the past, the design of a Capillary Pumped Loop involved mainly on the thermodynamics and heat transfer aspects of the system. The fluid flow dynamics of the working fluid were deemed benign to the system performance. Recently theoretical and experimental studies have revealed several mechanisms that led to the deprime of the capillary pumps. These mechanisms were all related to the dynamics of the fluid movement inside the loop.

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 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) is a prototype of the Earth Observing System (EOS) instrument thermal control systems. Four CAPL flight hardware components were tested in the Instrument Thermal Test Bed at NASA's Goddard Space Flight Center. The components tested were the capillary cold plates, capillary starter pump, heat pipe heat exchangers (HPHXs), and reservoir. The testing verified that all components meet or exceed their individual performance specifications. Consequently, the components have been integrated into the CAPL experiment which will be flown on the Space Shuttle in late 1993.

A Capillary Pumped Loop (CPL) is an advanced two-phase heat transport device which utilizes capillary forces developed within porous wicks to move a working fluid. The advantage this system has over conventional thermal management systems is its ability to transfer large heat loads over long distances at a controlled temperature. Extensive ground testing and two flight experiments have been performed over the past decade which have demonstrated the potential of the CPL as a reliable and versatile thermal control system for space applications. While the performance of CPL's as “black boxes” is now well understood, the internal thermo-fluid dynamics in a CPL are poorly known due to the difficulty of taking internal measurements. In order to visualize transient thermohydraulic processes occurring inside an evaporator, a see-through capillary evaporator was built and tested at NASA's Goddard Space Flight Center.

The High-Power Spacecraft Thermal Management (HPSTM) system was modified and upgraded to facilitate improved performance testing. Modifications to the system included augmenting the heat dissipation capacity of the condenser sink for steady-state high power operation, adding more pressure transducers to monitor pressure drops in various components of the system, installing pressure contact thermocouples on the evaporators to measure the heating surface temperature, providing a coolant loop to one of the evaporator plates for heat load sharing operation, installing a load cell on the reservoir to monitor transient fluid flows, and re-orienting the reservoir to reduce the effects of compressed vapor during transient operations. The system demonstrated a steady, continuous operation at a power input of 20 kW for 10 hours in the capillary mode. Test results also showed about 33% less variation of the reservoir set point temperature during power transients.