The thermal decomposition of oxygen difluoride in a flow system.
Thermal decomposition kinetics for oxygen difluoride shown not to represent unimolecular decomposition
Dauerman, L.↗
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A Study of the Effects of Altitude on Thermal Ice Protection System Performance
Thermal ice protection systems use heat energy to prevent a dangerous buildup of ice on an aircraft. As aircraft become more efficient, less heat energy is available to operate a thermal ice protections system. This requires that thermal ice protection systems be designed to more exacting standards so as to more efficiently prevent a dangerous ice buildup without adversely affecting aircraft safety. While the effects of altitude have always beeing taked into account in the design of thermal ice protection systems, a better understanding of these effects is needed so as to enable more exact design, testing, and evaluation of these systems.
Addy, Harold E., Jr.↗
Materials selection for long life in LEO: A critical evaluation of atomic oxygen testing with thermal atom systems
The use of thermal atom test methods as a materials selection and screening technique for low-Earth orbit (LEO) spacecraft is critically evaluated. The chemistry and physics of thermal atom environments are compared with the LEO environment. The relative reactivities of a number of materials determined to be in thermal atom environments are compared to those observed in LEO and in high quality LEO simulations. Reaction efficiencies measured in a new type of thermal atom apparatus are one-hundredth to one-thousandth those observed in LEO, and many materials showing nearly identical reactivities in LEO show relative reactivities differing by as much as a factor of 8 in thermal atom systems. A simple phenomenological kinetic model for the reaction of oxygen atoms with organic materials can be used to explain the differences in reactivity in different environments. Certain specific thermal test environments can be used as reliable materials screening tools. Using thermal atom methods to predict material lifetime in LEO requires direct calibration of the method against LEO data or high quality simulation data for each material.
Koontz, S. L.↗
Materials selection for long life in low earth orbit - A critical evaluation of atomic oxygen testing with thermal atom systems
The use of thermal atom test methods as a materials selection and screening technique for low-earth orbit (LEO) spacecraft is critically evaluated. The chemistry and physics of thermal atom environments are compared with the LEO environment. The relative reactivities of a number of materials determined in thermal atom environments are compared with those observed in LEO and in high-quality LEO simulations. Reaction efficiencies (cu cm/atom) measured in a new type of thermal atom apparatus are one-thousandth to one ten-thousandth those observed in LEO, and many materials showing nearly identical reactivities in LEO show relative reactivities differing by as much as a factor of eight in thermal atom systems. A simple phenomenological kinetic model for the reaction of oxygen atoms with organic materials can be used to explain the differences in reactivity in different environments. Certain speciic thermal atom test environments can be used as reliable materials screening tools.
Koontz, S. L.↗
Performance of two-layer thermal barrier systems on directionally solidified Ni-Al-Mo and comparative effects of alloy thermal expansion on system life
A promising two-layer thermal barrier coating system (TBS), Ni-16.4Cr-5.1A1-0.15Y/ZrO2-6.1Y2O3 (all in weight percent), was identified for directionally solidified Ni-Al-Mo (gamma/gamma' alpha). In cyclic furnace tests at 1095 C this system on gamma/gamma' alpha was better than Ni-16. 4Cr-5.1Al-0.15Y/ZrO2-7.8Y2O3 by about 50 percent. In natural gas - oxygen torch rig tests at 1250 C the ZrO2-6.1Y2O3 coating was better than the ZrO2-7.8Y2O3 coating by 95 percent, on MAR-M509 substrates and by 60 percent on gamma/gamma' alpha substrates. Decreasing the coefficient of thermal expansion of the substrate material from 17-18x10 to the -6 power/C (MAR-M200 + Hf and MAR-M509) to 11x10 to the -6 power/C (gamma/gamma' alpha) also resulted in improved TBS life. For example, in natural gas - oxygen torch rig tests at 1250 C, the life of Ni-16.4Cr-5.1Al-0.15Y/ZrO26.1Y2O3 was about 30 percent better on gamma/gamma' alpha than on MAR-M509 substrates. Thus compositional changes in the bond and thermal barrier coatings were shown to have a greater effect on TBS life than does the coefficient of thermal expansion.
Stecura, S.↗
Ballistic Performance of Porous-Ceramic, Thermal-Protection-Systems
Porous-ceramic, thermal protection systems are used heavily in current reentry vehicles like the Space Shuttle and are currently being proposed for the next generation of manned spacecraft, Orion. These materials insulate the structural components of a spacecraft against the intense thermal environments of atmospheric reentry. Furthermore, these materials are also highly exposed to space environmental hazards like meteoroid and orbital debris impacts. This paper discusses recent impact testing up to 9 km/s, and the findings of the influence of material equation-of-state on the simulation of the impact event to characterize the ballistic performance of these materials. These results will be compared with heritage models1 for these materials developed from testing at lower velocities. Assessments of predicted spacecraft risk based upon these tests and simulations will also be discussed.
Christiansen, E. L.↗
Improved Thermal-Insulation Systems for Low Temperatures
Improved thermal-insulation materials and structures and the techniques for manufacturing them are undergoing development for use in low-temperature applications. Examples of low-temperature equipment for which these thermal insulation systems could provide improved energy efficiency include storage tanks for cryogens, superconducting electric-power-transmission equipment, containers for transport of food and other perishable commodities, and cold boxes for low-temperature industrial processes. These systems could also be used to insulate piping used to transfer cryogens and other fluids, such as liquefied natural gas, refrigerants, chilled water, crude oil, or low-pressure steam. The present thermal-insulation systems are layer composites based partly on the older class of thermal-insulation systems denoted generally as multilayer insulation (MLI). A typical MLI structure includes an evacuated jacket, within which many layers of radiation shields are stacked or wrapped close together. Low-thermal-conductivity spacers are typically placed between the reflection layers to keep them from touching. MLI can work very well when a high vacuum level (less than 10(exp-4) torr) is maintained and utmost care is taken during installation, but its thermal performance deteriorates sharply as the pressure in the evacuated space rises into the soft vacuum range [pressures greater than 0.1 torr (greater than 13 Pa)]. In addition, the thermal performance of MLI is extremely sensitive to mechanical compression and edge effects and can easily decrease from one to two orders of magnitude from its ideal value even when the MLI is kept under high vacuum condition. The present thermal-insulation systems are designed to perform well under soft vacuum level, in particular the range of 1 to 10 torr. They are also designed with larger interlayer spacings to reduce vulnerability to compression (and consequent heat leak) caused by installation and use. The superiority of these systems is the synergistic effect of improvements in materials, design, and manufacture.
Fesmire, James E.↗
Displacements of Metallic Thermal Protection System Panels During Reentry
Bowing of metallic thermal protection systems for reentry of a previously proposed single-stage-to-orbit reusable launch vehicle was studied. The outer layer of current metallic thermal protection system concepts typically consists of a honeycomb panel made of a high temperature nickel alloy. During portions of reentry when the thermal protection system is exposed to rapidly varying heating rates, a significant temperature gradient develops across the honeycomb panel thickness, resulting in bowing of the honeycomb panel. The deformations of the honeycomb panel increase the roughness of the outer mold line of the vehicle, which could possibly result in premature boundary layer transition, resulting in significantly higher downstream heating rates. The aerothermal loads and parameters for three locations on the centerline of the windward side of this vehicle were calculated using an engineering code. The transient temperature distributions through a metallic thermal protection system were obtained using 1-D finite volume thermal analysis, and the resulting displacements of the thermal protection system were calculated. The maximum deflection of the thermal protection system throughout the reentry trajectory was 6.4 mm. The maximum ratio of deflection to boundary layer thickness was 0.032. Based on previously developed distributed roughness correlations, it was concluded that these defections will not result in tripping the hypersonic boundary layer.
Daryabeigi, Kamran↗
System Level Analysis of a Water PCM HX Integrated into Orion's Thermal Control System
In a cyclical heat load environment such as low Lunar orbit, a spacecraft's radiators are not sized to reject the full heat load requirement. Traditionally, a supplemental heat rejection device (SHReD) such as an evaporator or sublimator is used to act as a "topper" to meet the additional heat rejection demands. Utilizing a Phase Change Material (PCM) heat exchanger (HX) as a SHReD provides an attractive alternative to evaporators and sublimators as PCM HXs do not use a consumable, thereby leading to reduced launch mass and volume requirements. In continued pursuit of water PCM HX development an Orion system level analysis was performed using Thermal Desktop for a water PCM HX integrated into Orion's thermal control system in a 100km Lunar orbit. The study verified of the thermal model by using a wax PCM and analyzed 1) placing the PCM on the Internal Thermal Control System (ITCS) versus the External Thermal Control System (ETCS) 2) use of 30/70 PGW verses 50/50 PGW and 3) increasing the radiator area in order to reduce PCM freeze times. The analysis showed that for the assumed operating and boundary conditions utilizing a water PCM HX on Orion is not a viable option for any case. Additionally, it was found that the radiator area would have to be increased by at least 40% in order to support a viable water-based PCM HX.
Navarro, Moses↗
Weight Optimization of Active Thermal Management Using a Novel Heat Pump
Efficient lightweight power generation and thermal management are two important aspects for space applications. Weight is added to the space platforms due to the inherent weight of the onboard power generation equipment and the additional weight of the required thermal management systems. Thermal management of spacecraft relies on rejection of heat via radiation, a process that can result in large radiator mass, depending upon the heat rejection temperature. For some missions, it is advantageous to incorporate an active thermal management system, allowing the heat rejection temperature to be greater than the load temperature. This allows a reduction of radiator mass at the expense of additional system complexity. A particular type of active thermal management system is based on a thermodynamic cycle, developed by the authors, called the Solar Integrated Thermal Management and Power (SITMAP) cycle. This system has been a focus of the authors research program in the recent past (see Fig. 1). One implementation of the system requires no moving parts, which decreases the vibration level and enhances reliability. Compression of the refrigerant working fluid is accomplished in this scheme via an ejector.
Lear, William E.↗
Thermal protection system refurbishment cost study
Thermal protection system refurbishment cost study for space shuttles
Goldin, R. W.↗
Space shuttle heat pipe thermal control systems
Heat pipe (HP) thermal control systems designed for possible space shuttle applications were built and tested under this program. They are: (1) a HP augmented cold rail, (2) a HP/phase change material (PCM) modular heat sink and (3) a HP radiating panel for compartment temperature control. The HP augmented cold rail is similar to a standard two-passage fluid cold rail except that it contains an integral, centrally located HP throughout its length. The central HP core helps to increase the local power density capability by spreading concentrated heat inputs over the entire rail. The HP/PCM modular heat sink system consists of a diode HP connected in series to a standard HP that has a PCM canister attached to its mid-section. It is designed to connect a heat source to a structural heat sink during normal operation, and to automatically decouple from it and sink to the PCM whenever structural temperatures are too high. The HP radiating panel is designed to conductively couple the panel feeder HPs directly to a fluid line that serves as a source of waste heat. It is a simple strap-on type of system that requires no internal or external line modifications to distribute the heat to a large radiating area.
Alario, J.↗
In-flight testing of the space shuttle orbiter thermal control system
In-flight thermal control system testing of a complex manned spacecraft such as the space shuttle orbiter and the considerations attendant to the definition of the tests are described. Design concerns, design mission requirements, flight test objectives, crew vehicle and mission risk considerations, instrumentation, data requirements, and real-time mission monitoring are discussed. An overview of the tests results is presented.
Taylor, J. T.↗
Impact of Fluid Flow Pressure drop on Temperature of Components Controlled by Mechanically Pumped Fluid Loop Thermal Control System
In a typical thermal control system that employs a mechanically pumped fluid loop to control component temperatures, fluid flow pressure drop is minimized to ensure that adequate fluid flow can be provided by the pump. Engineers often are overly concerned about violating a maximum pressure drop, almost treating it as a “wall” that should not be crossed. The true “wall” is the control of the temperatures of the components served by the loop to ensure that their temperatures do not violate their allowed limits. As a case study, for assessing this “constraint,” sensitivity analyses were undertaken to understand if this is truly a major constraint, and to assess the impact of pressure drop exceedance on the controlled components’ temperatures. The flow impedance is varied as a parameter to understand how it affects the controlled components’ temperatures. One key finding was that the temperatures of the controlled components is relatively insensitive to the flow impedance increases. This finding is obviously most relevant to this particular case study; however, this also provides the process and guidance for assessing the performance of other pump/loop combinations in terms of their sensitivity to pressure drop impedances.
Mastropietro, Arthur J.↗
Study of the thermal design and performance of strapdown guidance and control systems
Thermal and mechanical design of strapdown inertial guidance and control systems
Sununu, J. H.↗
System Level Analysis of a Water PCM HX Integrated Into Orion's Thermal Control System Abstract
In a cyclical heat load environment such as low Lunar orbit, a spacecraft's radiators are not sized to reject the full heat load requirement. Traditionally, a supplemental heat rejection device (SHReD) such as an evaporator or sublimator is used to act as a "topper" to meet the additional heat rejection demands. Utilizing a Phase Change Material (PCM) heat exchanger (HX) as a SHReD provides an attractive alternative to evaporators and sublimators as PCM HXs do not use a consumable, thereby leading to reduced launch mass and volume requirements. In continued pursuit of water PCM HX development an Orion system level analysis was performed using Thermal Desktop for a water PCM HX integrated into Orion's thermal control system and in a 100km Lunar orbit. The study analyzed 1) placing the PCM on the Internal Thermal Control System (ITCS) versus the External Thermal Control System (ETCS) 2) use of 30/70 PGW verses 50/50 PGW and 3) increasing the radiator area in order to reduce PCM freeze times. The analysis showed that for the assumed operating and boundary conditions utilizing a water PCM HX on Orion is not a viable option. Additionally, it was found that the radiator area would have to be increased over 20% in order to have a viable water‐based PCM HX.
Navarro, Moses↗
Thermal Protection System with Staggered Joints
The thermal protection system disclosed herein is suitable for use with a spacecraft such as a reentry module or vehicle, where the spacecraft has a convex surface to be protected. An embodiment of the thermal protection system includes a plurality of heat resistant panels, each having an outer surface configured for exposure to atmosphere, an inner surface opposite the outer surface and configured for attachment to the convex surface of the spacecraft, and a joint edge defined between the outer surface and the inner surface. The joint edges of adjacent ones of the heat resistant panels are configured to mate with each other to form staggered joints that run between the peak of the convex surface and the base section of the convex surface.
Simon, Xavier D.↗
NASA Crew Exploration Vehicle, Thermal Protection System, Lessons Learned
The Orion (CEV) thermal protection system (TPS) advanced development project (ADP) was initiated in late 2006 to reduce developmental risk by significant investment in multiple heat shield architectural solutions that can meet the needs both the Low Earth orbit (LEO) and Lunar return missions. At the same time, the CEV TPS ADP was also charged with developing a preliminary design for the heat shield to meet the PDR requirement and at the time of the PDR, transfer the design to Lockheed- Martin, the prime contractor. We reported on the developmental activities of the first 18 months at the IPPW5 in Bordeaux, France, last summer. In June 08, at the time of the IPPW6, the CEV TPS ADP would have nearly completed the preparation for the Orion PDR and would be close to the original three-year mark. We plan to report on the progress at the Atlanta workshop. In the past year, Orion TPS ADP investment in TPS Technology, especially in PICA ablative Heat-shield design, development, testing and engineering (DDTE) has paid off in enabling MSL mission to switch from SLA 561 V heat shield to PICA heat shield. CEV TPS ADP considered SLA 561 V as a candidate for LEO missions and our testing identified failure modes in SLA and as a result, we dropped SLA for further evaluation. This close synergy between two projects is a highly visible example of how investment in technology areas can and does benefit multiple missions. In addition, CEV TPS ADP has been able to revive the Apollo ablative system namely AVCOAT honeycomb architecture as an alternate to the baseline PICA architecture and we plan to report the progress we have made in AVCOAT. CEV TPS ADP has invested considerable resources in developing analytical models for PICA and AVCOAT, material property measurements that is essential to the design of the heat-shield, in arcjet testing, in understanding the differences between different arc jet facilities, namely NASA Ames, NASA JSC and Air Force's AEDC, and in Non-Destructive Evaluation (NDE), and in integration of and manufacturing heat shield as a system. The capabilities of the two heat shield systems including failure modes via testing and analysis, once established, can serve the Probe Community and future mission designers to inner and outer planetary exploration very well. For example, missions to Venus, Mars and Titan can use either one of the system by selecting the mission design parameters that utilizes the full characteristics of these system to make use of system efficiency that will result in reduced heat shield mass, system robustness that will enhance mission success and cost. We plan to present significant progresses of the past three years and highlight the significant contributions CEV TPS ADP Project has made to advance the state of the art in Thermal Protection System technology that has and will continue to benefit future entry probe missions.
NASA Crew Exploration Vehicle↗