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Search indexed NASA NTRS and DOE OSTI research on propulsion, heat transfer, battery materials and energy systems. Follow report and document links to the original sources.

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At least 217 records · Page 12
All sourcesNASA NTRSDOE OSTI
NASA NTRS2012

From Concept-to-Flight: An Active Active Fluid Loop Based Thermal Control System for Mars Science Laboratory Rover

The Mars Science Laboratory (MSL) rover, Curiosity, which was launched on November 26, 2011, incorporates a novel active thermal control system to keep the sensitive electronics and science instruments at safe operating and survival temperatures. While the diurnal temperature variations on the Mars surface range from -120 C to +30 C, the sensitive equipment are kept within -40 C to +50 C. The active thermal control system is based on a single-phase mechanically pumped fluid loop (MPFL) system which removes or recovers excess waste heat and manages it to maintain the sensitive equipment inside the rover at safe temperatures. This paper will describe the entire process of developing this active thermal control system for the MSL rover from concept to flight implementation. The development of the rover thermal control system during its architecture, design, fabrication, integration, testing, and launch is described.

heat rejection systems↗
NASA NTRS2010

Design and Preliminary Thermal Performance of the Mars Science Laboratory Rover Heat Exchangers

The challenging range of proposed landing sites for the Mars Science Laboratory Rover requires a rover thermal management system that is capable of keeping temperatures controlled across a wide variety of environmental conditions. On the Martian surface where temperatures can be as cold as -123 degrees Centigrade and as warm as 38 degrees Centigrade, the Rover relies upon a Mechanically Pumped Fluid Loop (MPFL) and external radiators to maintain the temperature of sensitive electronics and science instruments within a -40 degrees Centigrade to 50 degrees Centigrade range. The MPFL also manages significant waste heat generated from the Rover power source, known as the Multi Mission Radioisotope Thermoelectric Generator (MMRTG). The MMRTG produces 110 Watts of electrical power while generating waste heat equivalent to approximately 2000 Watts. Two similar Heat Exchanger (HX) assemblies were designed to both acquire the heat from the MMRTG and radiate waste heat from the onboard electronics to the surrounding Martian environment. Heat acquisition is accomplished on the interior surface of each HX while heat rejection is accomplished on the exterior surface of each HX. Since these two surfaces need to be at very different temperatures in order for the MPFL to perform efficiently, they need to be thermally isolated from one another. The HXs were therefore designed for high in-plane thermal conductivity and extremely low through-thickness thermal conductivity by using aerogel as an insulator inside composite honeycomb sandwich panels. A complex assembly of hand welded and uniquely bent aluminum tubes are bonded onto the HX panels and were specifically designed to be easily mated and demated to the rest of the Rover Heat Recovery and Rejection System (RHRS) in order to ease the integration effort. During the cruise phase to Mars, the HX assemblies serve the additional function of transferring heat from the Rover MPFL to the separate Cruise Stage MPFL so that heat generated deep inside the Rover can be dissipated via the Cruise Stage radiators. Significant fabrication challenges had to be overcome in order to make the HX design a reality. The cruise phase thermal performance of the Rover HXs was verified in the cruise phase system level thermal vacuum test that was performed at JPL in January of 2009. The Rover HXs were modeled in I-DEAS TMG and predictions are compared to actual data from the test.

Heat Rejection System↗
NASA NTRS2001

Lightweight Radiators Being Developed or Advanced Stirling Radioisotope Power Systems

The thermodynamic heat-to-electric power conversion efficiency of Stirling systems is 3 to 5 times higher than that of thermoelectric converters. Hence for unmanned deep space probes, Stirling advanced radioisotope power systems (ARPS) could deliver up to 5 times as much power as radioisotope thermoelectric generators for the same amount of radioisotope, or they could require one-third to one-fifth as much isotope inventory for the same power output. However, Stirling power systems reject unconverted heat at much lower temperatures than radioisotope thermoelectric generators. Normally, this requires larger and heavier heat-rejection subsystems because of the greater radiator areas, which are proportional to the first power of the heat rejected and the fourth power of the absolute heat-rejection temperature, as specified by the Stefan-Boltzmann radiation heat transfer law. The development of directly coupled disk radiators using very high conductivity encapsulated thermopyrolitic graphite materials represents a significant advance in Stirling ARPS space heat-rejection subsystem technology. A conceptual Stirling ARPS with two engines coupled to a radioisotope general-purpose heat source (GPHS) is shown in the illustration.

Juhasz, Albert J.↗
NASA NTRS2012

Multi-Mission Radioisotope Thermoelectric Generator Heat Exchangers for the Mars Science Laboratory Rover

The addition of the Multi-Mission Radioisotope Thermoelectric Generator (MMRTG) to the Mars Science Laboratory (MSL) Rover requires an advanced thermal control system that is able to both recover and reject the waste heat from the MMRTG as needed in order to maintain the onboard electronics at benign temperatures despite the extreme and widely varying environmental conditions experienced both on the way to Mars and on the Martian surface. Based on the previously successful Mars landed mission thermal control schemes, a mechanically pumped fluid loop (MPFL) architecture was selected as the most robust and efficient means for meeting the MSL thermal requirements. The MSL heat recovery and rejection system (HRS) is comprised of two Freon (CFC-11) MPFLs that interact closely with one another to provide comprehensive thermal management throughout all mission phases. The first loop, called the Rover HRS (RHRS), consists of a set of pumps, thermal control valves, and heat exchangers (HXs) that enables the transport of heat from the MMRTG to the rover electronics during cold conditions or from the electronics straight to the environment for immediate heat rejection during warm conditions. The second loop, called the Cruise HRS (CHRS), is thermally coupled to the RHRS during the cruise to Mars, and provides a means for dissipating the waste heat more directly from the MMRTG as well as from both the cruise stage and rover avionics by promoting circulation to the cruise stage radiators. A multifunctional structure was developed that is capable of both collecting waste heat from the MMRTG and rejecting the waste heat to the surrounding environment. It consists of a pair of honeycomb core sandwich panels with HRS tubes bonded to both sides. Two similar HX assemblies were designed to surround the MMRTG on the aft end of the rover. Heat acquisition is accomplished on the interior (MMRTG facing) surface of each HX while heat rejection is accomplished on the exterior surface of each HX. Since these two surfaces need to be at very different temperatures in order for the fluid loops to perform efficiently, they need to be thermally isolated from one another. The HXs were therefore designed for high in-plane thermal conductivity and extremely low through-thickness thermal conductivity by using aluminum facesheets and aerogel as insulation inside a composite honeycomb core. Complex assemblies of hand-welded and uniquely bent aluminum tubes are bonded onto each side of the HX panels, and are specifically designed to be easily mated and demated to the rest of the RHRS in order to ease the integration effort.

Mastropietro, A. J.↗
NASA NTRS1985

Thermionic reactors for space nuclear power

Thermionic reactor designs for a variety of space power applications spanning the range from 5 kWe to 3 MWe are described. In all of these reactors, nuclear heat is converted directly to electrical energy in thermionic fuel elements (TFEs). A circulating reactor coolant carries heat from the core of TFEs directly to a heat rejection radiator system. The recent design of a thermionic reactor to meet the SP-100 requirements is emphasized. Design studies of reactors at other power levels show that the same TFE can be used over a broad range in power, and that design modifications can extend the range to many megawatts. The design of the SP-100 TFE is similar to that of TFEs operated successfully in test reactors, but with design improvements to extend the operating lifetime to seven years.

Homeyer, W. G.↗
NASA NTRS2010

Thermal Performance of Orion Active Thermal Control System With Seven-Panel Reduced-Curvature Radiator

The active thermal control system (ATCS) of the crew exploration vehicle (Orion) uses radiator panels with fluid loops as the primary system to reject heat from spacecraft. The Lockheed Martin (LM) baseline Orion ATCS uses eight-panel radiator coated with silver Teflon coating (STC) for International Space Station (ISS) missions, and uses seven-panel radiator coated with AZ 93 white paint for lunar missions. As an option to increase the radiator area with minimal impact on other component locations and interfaces, the reduced-curvature (RC) radiator concept was introduced and investigated here for the thermal perspective. Each RC radiator panel has 15 percent more area than each Lockheed Martin (LM) baseline radiator panel. The objective was to determine if the RC seven-panel radiator concept could be used in the ATCS for both ISS and lunar missions. Three radiator configurations the LM baseline, an RC seven-panel radiator with STC, and an RC seven-panel radiator with AZ 93 coating were considered in the ATCS for ISS missions. Two radiator configurations the LM baseline and an RC seven-panel radiator with AZ 93 coating were considered in the ATCS for lunar missions. A Simulink/MATLAB model of the ATCS was used to compute the ATCS performance. Some major hot phases on the thermal timeline were selected because of concern about the large amount of water sublimated for thermal topping. It was concluded that an ATCS with an RC seven-panel radiator could be used for both ISS and lunar missions, but with two different coatings STC for ISS missions and AZ 93 for lunar missions to provide performance similar to or better than that of the LM baseline ATCS.

Wang, Xiao-Yen J.↗
NASA NTRS2011

Spacesuit Evaporator-Absorber-Radiator (SEAR)

For decades advanced spacesuit developers have pursued a regenerable, robust non-venting system for heat rejection. Toward this end, this paper investigates linking together two previously developed technologies, namely NASA's Spacesuit Water Membrane Evaporator (SWME), and Creare's lithium chloride Heat Pump Radiator (HPR). Heat from a liquid cooled garment is transported to SWME that provides cooling through evaporation. The SEAR is evacuated at the onset of operations and thereafter, the water vapor absorption rate of the HPR maintains a low pressure environment for the SWME to evaporate effectively. This water vapor captured by solid LiCl in the HPR with a high enthalpy of absorption, results in sufficient temperature lift to reject most of the heat to space by radiation. After the sortie, the HPR would be heated up in a regenerator to drive off and recover the absorbed evaporant. A one-fourth scale prototype was built and tested in vacuum conditions at a sink temperature of 250 K. The HPR was able to stably reject 60 W over a 7-hour period. A conceptual design of a full-scale radiator is proposed. Excess heat rejection above 240 W would be accomplished through venting of the evaporant. Loop closure rates were predicted for various exploration environment scenarios.

Bue, Grant C.↗
NASA NTRS2012

Spacesuit Evaporator-Absorber-Radiator (SEAR)

For decades advanced spacesuit developers have pursued a regenerable, robust nonventing system for heat rejection. Toward this end, this paper investigates linking together two previously developed technologies, namely NASA s Spacesuit Water Membrane Evaporator (SWME), and Creare s Lithium Chloride Absorber Radiator (LCAR). Heat from a liquid cooled garment is transported to SWME that provides cooling through evaporation. This water vapor is then captured by solid LiCl in the LCAR with a high enthalpy of absorption, resulting in sufficient temperature lift to reject heat to space by radiation. After the sortie, the LCAR would be heated up and dried in a regenerator to drive off and recover the absorbed evaporant. A engineering development prototype was built and tested in vacuum conditions at a sink temperature of 250 K. The LCAR was able to stably reject 75 W over a 7-hour period. A conceptual design of a full-scale radiator is proposed. Excess heat rejection above 240 W would be accomplished through venting of the evaporant. Loop closure rates were predicted for various exploration environment scenarios.

Hodgson, Ed↗
NASA NTRS2014

A Curious Year on Mars - Long-Term Thermal Trends for Mars Science Laboratory Rover's First Martian Year

By the time of this writing, Curiosity, the Mars Science Laboratory (MSL) Rover, has weathered four seasons in Gale Crater, just south of and approaching the foothills of the 5-km high Aeolis Mons, known as "Mount Sharp," at 4.59 deg south latitude. The mission design included a much broader latitude range of 30 deg north to 30 deg south constraining some of the Rover environmental requirements and operations. To date, Curiosity has relayed over 150 MB of thermal telemetry. Curiosity has relayed over 150 MB of thermal telemetry through four seasons. The trends and idiosyncrasies revealed through four seasons of telemetry from Mars are discussed. The better-characterized thermal environment allows for less conservatism in operational models and increases the amount of science data collection. Examples include: the elimination of overheating concerns for some cameras and the use of the previous sol's temperature telemetry along with the conservative soak temperature curve from the winter thermal model, to produce a custom heating prescription for the upcoming weeks thus increasing operation time and reducing heating times. This paper discusses the lessons learned for Rover operation as well as general idiosyncrasies discovered about the local environment-such as the effect of orientation on subsystem temperature variation, a regular morning and afternoon wind, ground and air microclimates with distinct temperature differences from other terrain, and how the Rover affects the local environment. This paper further documents and explains some of the interesting highs and lows of the temperature telemetry data as well as offers explanations for sudden temperature changes on board the Rover.

HRS↗
NASA NTRS2024

xEMU Thermal Vacuum Testing Overview

The Exploration Extravehicular Mobility Unit (xEMU) project was the culimation of over a decade of spacesuit development that was performed in-house at the NASA Johnson Space Center. This project reached a level where almost fully completed development fidelity spacesuits had been designed, assembled and tested in an integrated configuration. The xEMU Development-Verification-Test (DVT) hardware was assembled into two different xEMU test articles and underwent a series of thermal-vacuum tests at the Johnson Space Center’s Chamber B. These tests not only gathered data on thermal performance, but also exercised the life support system, communication system, suit information systems, and suit avionics. This complex test has and will continue to produce many meaningful reports ranging from component level test results (for example on spacesuit boots), test design of heater cages to simulate thermal environments for a spacesuit test article, and higher level thermal performance of subsystems (such as the Portable Life Support System) or the entire assembly. This paper provides an overview of the test configuration and also top level results from this highly successful integrated test of the xEMU.

PLSS↗
NASA NTRSDate not supplied

xEMU Thermal Vacuum Testing Overview

The Exploration Extravehicular Mobility Unit (xEMU) project was the culimation of over a decade of spacesuit development that was performed in-house at the NASA Johnson Space Center. This project reached a level where almost fully completed development fidelity spacesuits had been designed, assembled and tested in an integrated configuration. The xEMU Development-Verification-Test (DVT) hardware was assembled into two different xEMU test articles and underwent a series of thermal-vacuum tests at the Johnson Space Center’s Chamber B. These tests not only gathered data on thermal performance, but also exercised the life support system, communication system, suit information systems, and suit avionics. This complex test has and will continue to produce many meaningful reports ranging from component level test results (for example on spacesuit boots), test design of heater cages to simulate thermal environments for a spacesuit test article, and higher level thermal performance of subsystems (such as the Portable Life Support System) or the entire assembly. This paper provides an overview of the test configuration and also top level results from this highly successful integrated test of the xEMU.

PLSS↗
NASA NTRS1985

Lightweight two-stroke cycle aircraft diesel engine technology enablement program, volume 1

An experimental Single Cylinder Test Engine Program is conducted to confirm the analytically projected performance of a two-stroke cycle diesel engine for aircraft applications. The test engine delivered 78kW indicated power from 1007cc displacement, operating at 3500 RPM on Schnuerle loop scavenged two-stroke cycle. Testing confirms the ability of a proposed 4-cylinder version of such an engine to reach the target power at altitude, in a highly turbocharged configuration. The experimental program defines all necessary parameters to permit design of a multicylinder engine for eventual flight applications; including injection system requirement, turbocharging, heat rejection, breathing, scavenging, and structural requirements. The multicylinder engine concept is configured to operate with an augmented turbocharger, but with no primary scavenge blower. The test program is oriented to provide a balanced turbocharger compressor to turbine power balance without an auxiliary scavenging system. Engine cylinder heat rejection to the ambient air has been significantly reduced and the minimum overall turbocharger efficiency required is within the range of commercially available turbochargers. Analytical studies and finite element modeling is made of insulated configurations of the engines - including both ceramic and metallic versions. A second generation test engine is designed based on current test results.

Freen, P. D.↗
NASA NTRS1991

EMU evolution

Evolution of Extravehicular Mobility Unit (EMU) technology is necessary to support the Extravehicular Activity (EVA) requirements of the Space Station Freedom Program and those of the Space Exploration Initiative (SEI). Key qualities supporting long-duration missions include technologies that are highly reliable, durable, minimize logistics requirements, and are in-flight maintainable and serviceable. While these qualities are common to SSF and SEI EVA, development paths will differ where specific mission requirements impose different constraints. Development of reusable, regenerative technologies is necessary to minimize the logistics penalties. Increased battery discharge/recharge cycle life and usable wet life, compact high current density fuel cells, reusable CO2 absorbing media, and thermal radiation coupled with venting heat rejection technologies are just some methods of reducing consumables. Development must strive for durable, reliable systems that are in-flight serviceable and maintainable, which are vital for missions where logistics capabilities are extremely constrained. Key areas include suit components (e.g., gloves, boots, and cooling garments), and life support hardware such as fans, pumps, instrumentation, and emergency O2 systems. Higher pressure suits will reduce EVA prebreathe requirements and pre-EVA operations overall. Many challenges of higher pressure suits have been addressed by on-going development. Emphasis on glove development is necessary to provide low fatigue, dexterous glove mobility at higher suit pressures. Minimum impact hooks and scars which support an advanced SSF EMU have been identified. These accommodations permit upgrades that support servicing of low volume, high pressure oxygen systems, and hydrogen technologies such as fuel cell, and venting hydrogen heat rejection systems.

Rouen, M.↗
NASA NTRS1981

A pumped, two-phase flow heat transport system for orbiting instrument payloads

A pumped two-phase (heat absorption/heat rejection) thermal transport system for orbiting instrument payloads is investigated. The thermofluid characteristics necessary for the system design are discussed. A preliminary design with a series arrangement of four instrument heat stations and six radiators in a single loop is described in detail, and the total mass is estimated to be 134 kg, with the radiators, instrument heat stations, and fluid reservoir accounting for approximately 86, 24, and 12 kg, respectively. The evaluation of preliminary test results shows that the system has potential advantages; however, further research is necessary in the areas of one-g and zero-g heat transfer coefficients/fluid regimes, fluid by-pass temperature control, and reliability of small pumps.

Fowle, A. A.↗
NASA NTRS1983

Thermal control - Heat buses will operate like a public utility

Active thermal control for the NASA space station concept requires long life heat rejection, highly versatile thermal transport, and efficient system integration. By a significant margin, the heat radiator will be the largest and most exposed portion of the space station thermal system. Transport requirements encompass the collection and movement of thermal energy from the space station's heat sources to the radiator heat sink at required temperature levels. In a decentralized thermal system, each space station module would collect and reject all of the waste heat generated, thereby requiring no module interconnections. This scheme does not, however, allow waste heat from one module to be used by another. In a centralized system, heat must be transported across module boundaries. A high capacity monogroove heat pipe has been developed to simplify space radiators design and operation.

Ellis, W. E.↗