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Heat rejection system requirements of specific mission types (space station, planetary exploration, commercial, very high power, and military missions) are discussed. Heat pipe radiators, weight and volume reduction, stable coatings, and working fluids are addressed.
Ceramic integration technology has been recognized as an enabling technology for the implementation of advanced ceramic systems in a number of high-temperature applications in aerospace, power generation, nuclear, chemical, and electronic industries. Various ceramic integration technologies (joining, brazing, attachments, repair, etc.) play a role in fabrication and manufacturing of large and complex shaped parts of various functionalities. However, the development of robust and reliable integrated systems with optimum performance requires the understanding of many thermochemical and thermomechanical factors, particularly for high temperature applications. In this presentation, various challenges and opportunities in design, fabrication, and testing of integrated similar (ceramic-ceramic) and dissimilar (ceramic-metal) material systems will be discussed. Experimental results for bonding and integration of SiC based LDI fuel injector, high conductivity C/C composite based heat rejection system, solid oxide fuel cells system, ultra high temperature ceramics for leading edges, and ceramic composites for thermostructural applications will be presented. Potential opportunities and need for the development of innovative design philosophies, approaches, and integrated system testing under simulated application conditions will also be discussed.
The X-ray telescope (XRT) on board the Swift Gamma Ray Burst Explorer has successfully operated since the spacecraft launch on 20 November 2004, automatically locating GRB afterglows, measuring their spectra and lightcurves and performing observations of high-energy sources. In this work we investigate the properties of the instrumental background, focusing on its dynamic behavior on both long and short timescales. The operational temperature of the CCD is the main factor that influences the XRT background level. After the failure of the Swift active on-board temperature control system, the XRT detector now operates at a temperature range between -75C and -45C thanks to a passive cooling Heat Rejection System. We report on the long-term effects on the background caused by radiation, consisting mainly of proton irradiation in Swift's low Earth orbit and on the short-term effects of transits through the South Atlantic Anomaly (SAA), which expose the detector to periods of intense proton flux. We have determined the fraction of the detector background that is due to the internal, instrumental background and the part that is due to unresolved astrophysical sources (the cosmic X-ray background) by investigating the degree of vignetting of the measured background and comparing it to the expected value from calibration data.
Two heat rejection radiator systems for the International Space Station (ISS) have undergone thermal vacuum qualification testing at the NASA Glenn Research Center (GRC), Plum Brook Station, Sandusky, Ohio. The testing was performed in the Space Power Facility (SPF), the largest thermal vacuum chamber in the world. The heat rejection system radiator was tested first; it removes heat from the ISS crew living quarters. The second system tested was the photovoltaic radiator (PVR), which rejects heat from the ISS photovoltaic arrays and the electrical power-conditioning equipment. The testing included thermal cycling, hot- and cold-soaked deployments, thermal gradient deployments, verification of the onboard heater controls, and for the PVR, thermal performance tests with ammonia flow. Both radiator systems are orbital replacement units for ease of replacement on the ISS. One key to the success of these tests was the performance of the infrared heater system. It was used in conjunction with a gaseous-nitrogen-cooled cryoshroud in the SPF vacuum chamber to achieve the required thermal vacuum conditions for the qualification tests. The heater, which was designed specifically for these tests, was highly successful and easily met the test requirements. This report discusses the heating requirements, the heater design features, the design approach, and the mathematical basis of the design.
Aerospace and terrestrial applications of thermionic direct energy conversion and advances in direct energy conversion (DEC) technology are surveyed. Electrode materials, the cesium plasma drop (the difference between the barrier index and the collector work function), DEC voltage/current characteristics, conversion efficiency, and operating temperatures are discussed. Attention is centered on nuclear reactor system thermionic DEC devices, for in-core or out-of-core operation. Thermionic fuel elements, the radiation shield, power conditions, and a waste heat rejection system are considered among the thermionic DEC system components. Terrestrial applications include topping power systems in fossil fuel and solar power generation.
A computer program was developed to facilitate parametric performance evaluation of heat pipes in lightweight heat rejection systems. A description of the code, user's manual, and sample inputs are provided. The emphasis is placed on the analysis and design of homogeneous wick heat pipes. The analysis of the annular heat pipe is included as part of the heat pipe radiator subroutine.
A compilation of several lunar surface thermal management and power system studies completed under contract and IR&D is presented. The work includes analysis and preliminary design of all major components of an integrated thermal management system, including loads determination, active internal acquisition and transport equipment, external transport systems (active and passive), passive insulation, solar shielding, and a range of lunar surface radiator concepts. Several computer codes were utilized in support of this study, including RADSIM to calculate radiation exchange factors and view factors, RADIATOR (developed in-house) for heat rejection system sizing and performance analysis over a lunar day, SURPWER for power system sizing, and CRYSTORE for cryogenic system performance predictions. Although much of the work was performed in support of lunar rover studies, any or all of the results can be applied to a range of surface applications. Output data include thermal loads summaries, subsystem performance data, mass, and volume estimates (where applicable), integrated and worst-case lunar day radiator size/mass and effective sink temperatures for several concepts (shielded and unshielded), and external transport system performance estimates for both single and two-phase (heat pumped) transport loops. Several advanced radiator concepts are presented, along with brief assessments of possible system benefits and potential drawbacks. System point designs are presented for several cases, executed in support of the contract and IR&D studies, although the parametric nature of the analysis is stressed to illustrate applicability of the analysis procedure to a wide variety of lunar surface systems. The reference configuration(s) derived from the various studies will be presented along with supporting criteria. A preliminary design will also be presented for the reference basing scenario, including qualitative data regarding TPS concerns and issues.
It was determined that space tug will require the capability to perform its mission within a broad range of thermal environments with currently planned mission durations of up to seven days, so an investigation was conducted to define a thermal design for the forward and intertank compartments and fuel cell heat rejection system that satisfies tug requirements for low inclination geosynchronous deploy and retrieve missions. Passive concepts were demonstrated analytically for both the forward and intertank compartments, and a worst case external heating environment was determined for use during the study. The thermal control system specifications and designs which resulted from the research are shown.
In-cylinder components and tribological system concepts were designed, fabricated and tested at conditions anticipated for a 55% thermal efficiency heavy duty diesel engine for the year 2000 and beyond. A Cummins L10 single cylinder research engine was used to evaluate a spherical joint piston and connecting rod with 19.3 MPa (2800 psi) peak cylinder pressure capability, a thermal fatigue resistant insulated cylinder head, radial combustion seal cylinder liners, a highly compliant steel top compression ring, a variable geometry turbocharger, and a microwave heated particulate trap. Components successfully demonstrated in the final test included spherical joint connecting rod with a fiber reinforced piston, high conformability steel top rings with wear resistant coatings, ceramic exhaust ports with strategic oil cooling and radial combustion seal cylinder liner with cooling jacket transfer fins. A Cummins 6B diesel was used to develop the analytical methods, materials, manufacturing technology and engine components for lighter weight diesel engines without sacrificing performance or durability. A 6B diesel engine was built and tested to calibrate analytical models for the aluminum cylinder head and aluminum block.
Development of the telescoping boom radiator from concept studies through a detailed design is reviewed. Particular attention is given to the scaling law for specific mass of cylindrical radiator geometries and radiator mass projections for large scale heat rejection systems. The concept of the telescoping, cylindrical radiator is identified as a leading candidate for rejecting tens of megawatts of thermal energy while maintaining a reasonable launch.
Hamilton Sundstrand is under contract with the NASA Johnson Space Center to develop a scalable, evaporative heat rejection system called the Multi-Fluid Evaporator (MFE). It is being designed to support the Orion Crew Module and to support future Constellation missions. The MFE would be used from Earth sea level conditions to the vacuum of space. The current Shuttle configuration utilizes an ammonia boiler and flash evaporator system to achieve cooling at all altitudes. The MFE system combines both functions into a single compact package with significant weight reduction and improved freeze-up protection. The heat exchanger core is designed so that radial flow of the evaporant provides increasing surface area to keep the back pressure low. The multiple layer construction of the core allows for efficient scale up to the desired heat rejection rate. The full scale MFE prototype will be constructed with four core sections that, combined with a novel control scheme, manage the risk of freezing the heat exchanger cores. A sub-scale MFE engineering development unit (EDU) has been built, and is identical to one of the four sections of a full scale prototype. The EDU has completed testing at Hamilton Sundstrand. The overall test objective was to determine the thermal performance of the EDU. The first set of tests simulated how each of the four sections of the prototype would perform by varying the chamber pressure, evaporant flow rate, coolant flow rate and coolant temperature. A second set of tests was conducted with an outlet steam header in place to verify that the outlet steam orifices prevent freeze-up in the core while also allowing the desired thermal turn-down ratio. This paper discusses the EDU tests and results.
Routine and complex EVAs call for more effective heat rejection systems in order to maximize mission productivity; an optimum EVA mobility unit (EMU) thermal control subsystem must require no expendables and introduce no contaminants into the environment, while conforming to minimum size limits and allowing easy regeneration. Attention is presently given to two thermal control subsystems, one of which can be integrated with the existing Space Shuttle Orbiter EMU to provide a 3-hour nonventing heat rejection capability, while the other can furnish the entire heat rejection capability requirement for an 8-hour Space Station EVA.
The Mars Science Laboratory (MSL), established as the most advanced rover to land on the surface of Mars to date, launched on November 26th, 2011 and arrived to the Martian Gale Crater during the night of August 5th, 2012 (PDT). MSL will investigate whether the landing region was ever suitable to support carbon-based life, and examine rocks, soil, and the atmosphere with a sophisticated suite of tools. This paper addresses the flight system requirement by which the vehicle transmitted indications of the following events using both X-band tones and UHF telemetry to allow identification of probable root causes should a mission anomaly have occurred: Heat-Rejection System (HRS) venting, completion of the cruise stage separation, turn to entry attitude, atmospheric deceleration, bank angle reversal commanded, parachute deployment, heatshield separation, radar ground acquisition, powered descent initiation, rover separation from the descent stage, and rover release. During Entry, Descent, and Landing (EDL), the flight system transmitted a UHF telemetry stream adequate to determine the state of the spacecraft (including the presence of faults) at 8 kbps initiating from cruise stage separation through at least one minute after positive indication of rover release on the surface of Mars. The flight system also transmitted X-band semaphore tones from Entry to Landing plus one minute although since MSL was occulted, as predicted, by Mars as seen from the Earth, Direct-To-Earth (DTE) communications were interrupted at approximately is approx. 5 min after Entry ( approximately 130 prior to Landing). The primary data return paths were through the Deep Space Network (DSN) for DTE and the existing Mars network of orbiting assets for UHF, which included the Mars Reconnaissance Orbiter (MRO), Mars Odyssey (ODY), and Mars Express (MEX) elements. These orbiters recorded the telemetry data stream and returned it back to Earth via the DSN. The paper also discusses the total power received during EDL and the robustness of the telecom design strategy used to ensure EDL communications coverage.
Basic guidelines are presented for a nuclear space power system which utilizes heat pipes to transport thermal power from a fast nuclear reactor to an out of core thermionic converter array. Design parameters are discussed for the nuclear reactor, heat pipes, thermionic converters, shields (neutron and gamma), waste heat rejection systems, and the electrical bus bar-cable system required to transport the high current/low voltage power to the processing equipment. Dimensions are compatible with shuttle payload bay constraints.
Technical improvements of a long life heat rejection system, suitable for long duration high power missions, that can be constructed and deployed in orbit is discussed. A mathematical model is formulated and a computer program developed which describes the transient priming characteristics of a dual passage heat pipe. An experimental test package is described for flight in the KC-135 Zero-g Aircraft, to be used to verify the modeling predictions.
A preliminary design study of a conceptual 6000-MW open-cycle gas-core nuclear rocket engine system was made. The engine has a thrust of 44,200 lb and a specific impulse of 4400 sec. The nuclear fuel is uranium-235 and the propellant is hydrogen. Critical fuel mass was calculated for several reactor configurations. Major components of the reactor (reflector, pressure vessel) and the waste heat rejection system were considered conceptually and were sized.
Reliable lightweight systems capable of providing electrical power at the magawatt level are a requirement for future manned space exploration missions. This can be achieved by the development of high temperature insulating materials which are not only capable of surviving the hostile space environment but can contribute to reducing the mass and weight of the heat rejection system. In this work, Apical, Upilex, Kapton, Teflon AF, and Peek polymers are characterized for AC and DC dielectric breakdown in air and in silicone oil at temperatures up to 250 C. The materials are also tested in terms of their dielectric constant and dissipation factor at high temperatures with an electrical stress of 60 Hz, 200 V/mil present. The effects of thermal aging on the properties of the films are determined after 15 hours of exposure to 200 and 250 C, each. The results obtained are discussed and conclusions are made concerning the suitability of these dielectrics for use in capacitors and cable insulations in high temperature environments.
A lightweight flexible radiator system for on-orbit cooling of space payloads is described. The radiator is packaged as a compact unit which can be attached to a vehicle structure or hatch prior to or after launch. On-orbit, it is deployed to provide the radiating surface needed for a specific experiment. The unit is being independently developed and qualified as a heat rejection system which will be ready for any spacecraft or experiment, and which will not require significant structural and systems accommodation. Design details and thermal vacuum test results are presented.