SEARCH · Engineering Papers
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.
Quote a phrase for an exact phrase match. Source license links do not imply unrestricted reuse.
This slide presentation reviews the thermal design for the Mars exploration rover project. It includes information on the spacecraft configuration, the cruise scenario, landing scenario, instrument package, thermal environment, and spacecraft schematics.
The efficiency, of the electronics plays a large part in determining the overall spacecraft power requirements to operate the cooler as well as determining the thermal dissipation characteristics of the various electrical components. The results of all the performance measurements are presented in the paper.
Effective thermal conductivities of Rene 41 panels suitable for advanced space transportation vehicle structures were determined analytically and experimentally for temperature ranges between 20.4K (423 F) and 1186K (1675 F). The cryogenic data were obtained using a cryostat whereas the high temperature data were measured using a heat flow meter and a comparative thermal conductivity instrument respectively. Comparisons were made between analysis and experimental data. Analytical methods appear to provide reasonable definition of the honeycomb panel effective thermal conductivities.
The new series of Geostationary Operational Environmental Satellites (GOES I,J,K,L/M) will have two earth viewing optical instruments which use passive, bi-metallic actuated, blade-type louvers for their thermal control. GOES is a three-axis stabilized spacecraft in geostationary orbit, which means that sun can enter the instrument apertures for several hours before and after local midnight. The solar heat absorbed in the instruments causes the louvers to open thus allowing the radiators to reject the heat. These louvered radiators are located on the north face of the instruments. To perform thermal analyses of the instruments, a modeling method was needed for predicting the thermal characteristics of the louvers. This task was complicated by the fact that sun shines on the louvers during the summer season. A Monte Carlo analytical technique was used to develop absorbed power and effective emissivity tables for the louvers. These tables could be used with the instrument math models to predict the thermal behavior of the instruments. Data from the Monte Carlo analysis showed that solar entrapment could have a significant effect on the heat rejection capability of the louvered radiators.
A detailed nodal computer model was developed to thermally represent the hardware, and sensitivity studies were performed to evaluate design parameters and orbital environmental effects of an instrument cooling system for IR detectors. Thermal-vacuum testing showed excellent performance of the system and a correspondence with math model predictions to within 3 K. Results show cold stage temperature sensitivity to cold patch backload, outer stage external surface emittance degradation, and cold stage emittance degradation, respectively. The increase in backload on the cold patch over the mission lifetime is anticipated to be less than 3.0 watts, which translates to less than a 3-degree increase in detector temperatures.
Explore the source record for details and available documents.
Based on the recent success of our strained-layer superlattice (SLS)-based infrared (IR) camera that performed Earth imaging from the International Space Station (ISS) in 2019 we have built, what we consider, to be the next generation multi-band SLS imaging system. The Compact Thermal Imager (CTI) was installed on the Robotic Refueling Mission 3 (RRM3) and attached to the exterior of the ISS. From this location we were able to capture 15 million images of a multitude of fires around the globe in 2019. This unexpected trove of data initiated quite a bit of scientific interest to further utilize this imaging capability but would include features to more precisely monitor terrestrial fires and other surface phenomena. To this end, we developed a technique to install specific bandpass filters directly onto the SLS detector hybrid assembly. Utilizing this technique we have built a CTI-2 camera system with two filters, 4 and 11µm, and have made a second detector assembly with six filter bands from 4-12µm. This second system will also be used to supplement Landsat remote imaging monitoring approximate land surface temperatures, monitor evapotranspiration, sea ice and glacier dynamics. The CTI-2 camera is based on a 1,024x1,024 (1kx1k) format SLS detector hybridized to a FLIR ISC0404 readout integrated circuit (ROIC). The six band SLS focal plane array is based on the 640x512 FLIR ISC 9803 ROIC. This camera system is based on the Landsat 8 and 9 Thermal IR Sensors (TIRS) instrument and one of its purposes is to perform ground truthing for the Landsat 8/9 data at higher spectral resolution. Both Landsat TIRS instruments are dual band thermal IR sensors centered on 11 and 12µm (each with about a 1µm bandpass). Both of our SLS systems utilize a Ricor K548 cryocooler. To streamline costs and development time we used commercial optics and both commercial and custom NASA electronic components. A primary feature of these camera systems is the incorporation of specific filters to collect fire data at ~3.9µm and thermal data at ~11µm. The CTI-2 instrument is designed for 37 m /pixel spatial resolution from 410km orbit (ISS orbit). In this paper, we will present the design and performance of the focal plane, optics, electronics and mechanical structure of the dual‑band CTI-2 and the focal plane performance of the six‑band focal plane.
The redesigned solid rocket motor (RSRM) flight set 360L003 was launched on March 13, 1989 as part of NASA space shuttle mission STS-29R. As was the case with flight sets 360L001 and 360L002 (STS-26R and STS-27R), both motors (360L003A and 360L003B) performed in an excellent manner. Evaluation of the ground environment instrumentation measurements verified thermal model analysis data and showed agreement with predicted environmental effects. The right-hand aft field joint primary heater failed during the countdown; the secondary heater was activated and performed as designed. All other field joint heaters and aft skirt thermal conditioning systems had no anomalies. Shuttle thermal imager infrared readings compared favorably with measured ground environment instrumentation data. No thermal launch commit criteria violations occurred at any time. Evaluation of the development flight instrumentation showed exceptional propulsion performance. All ballistic parameters closely matched the predicted values and were well within the required specification levels. Girth and biaxial strain gage measurements compared closely with corresponding gages on previous flight motors, static tests, and with preflight predictions. Adequate safety factors were verified. (Some ignition transient spiking was noted in a few girth gages; the spiking was determined not to be representative of actual case behavior, but an instrumentation phenomena.) The accelerometers again measured high vibration amplitude levels during the ignition transient and the reentry Max Q phases. Postflight inspection showed that all combustion gas was contained by the insulation in the field and case-to-nozzle joints. No anomalous insulation erosion patterns were found, and the seals that did directly contain motor pressure showed no heat effects, erosion, or blowby. All anomalies identified were a result of splashdown damage, with the exception of fretting in the case field joint interference (nonsealing) surfaces and a prelaunch field joint heater failure. The disposition of all anomalies and the complete results are reported.
Numerous separate effects irradiations of miniature nuclear fuel specimens have been conducted in the High Flux Isotope Reactor (HFIR) under the experimental platform designated as MiniFuel. MiniFuel is a static irradiation capability in which microstructural evolution and fuel performance phenomena are observed during postirradiation examination thereby offering a snapshot of the terminal fuel characteristics. This approach inherently requires fielding an irradiation where the experimental conditions are determined using predictive models and the pertinent outcomes are measured at the end of the test. Static irradiations can provide useful insights to the relationships between fuel performance and the pivotal irradiation conditions, namely temperature and burnup, but the ability to monitor fuel performance in situ would further support fuel development and qualification. To this end, an instrumented experiment design is being developed at Oak Ridge National Laboratory to capture thermal conductivity degradation and fission gas release during HFIR irradiation. These phenomena will be monitored using unique capsule designs that each target a different phenomenon. This paper details the thermal conductivity capsule (TCC) design and its expected performance envelope as determined using computer models. Each TCC will contain a miniature UO2 disc specimen (~0.5 mm thick × 5 mm diameter) sandwiched between metallic slugs with embedded thermocouples. The coupling of in situ temperature measurements, known thermal conductivity of the metallic components, and heat generation rates computed using high-fidelity neutronics models make the thermal conductivity measurement possible. This paper describes the reactor physics and heat transfer models used to predict the capsule’s performance and the methodology for calculating the fuel specimen’s thermal conductivity from the thermocouple measurements.
A highly adaptable and robust terahertz (THz) energy meter is designed and implemented to detect energetic THz pulses from high-intensity (>10 18 W/cm 2 ) laser–plasma interactions on the OMEGA EP. THz radiation from the laser driven target is detected by a shielded pyrometer. A second identical pyrometer is used for background subtraction. The detector can be configured to detect THz pulses in the 1 mm to 30 μm (0.3- to 10-THz) range and pulse energies from joules to microjoules via changes in filtration, aperture size, and position. Additional polarization selective filtration can also be used to determine the THz pulse polarization. The design incorporates significant radiation and electromagnetic pulse shielding to survive and operate within the OMEGA EP radiation environment. We describe the design, operational principle, calibration, and testing of the THz energy meter. The pyrometers were calibrated using a benchtop laser and show linear sensitivity to up to 1000 nJ of absorbed energy. Furthermore, the initial results from four OMEGA EP THz experiments detected up to ~15 μJ at the detector, which can correspond to hundreds of mJ depending on THz emission and reflection models.
A computerized tensimeter and experimental procedure for determination of the thermal decomposition temperature (T sub d) of perfluoro alkylethers were developed and tested. Both the apparatus and the procedure are described in detail. Results of testing with bis(2-ethylhexyl) phthalate and trimethylolpropane triheptanoate demonstrate that the reciprocal of the decomposition temperature is a linear function of the logarithm of the gas volume/heated liquid volume ratio. The T sub d obtained for each compound at a gas volume/heated liquid volume ration of one was similar to the value previously reported using an isoteniscope technique. Results of testing with a polymer of hexafluoropropylene oxide demonstrate that this instrument and procedure can be used to determine the T sub d of perfluoroalkylethers.
A computerized tensimeter and experimental procedure for determination of the thermal decomposition temperature (T sub d) of perfluoro alkylethers were developed and tested. Both the apparatus and the procedure are described in detail. Results of testing with bis(2-ethylhexyl) phthalate and trimethylolpropane triheptanoate demonstrate that the reciprocal of the decomposition temperature is a linear function of the logarithm of the gas volume/heated liquid volume ratio. The T sub d obtained for each compound at a gas volume/heated liquid volume ration of one was similar to the value previously reported using an isoteniscope technique. Results of testing with a polymer of hexafluoropropylene oxide demonstrate that this instrument and procedure can be used to determine the T sub d of perfluoro alkylethers.
This work presents a concept and implementation of a calibrated streaked spectral pyrometer (SSP) temperature diagnostic used in dynamically driven shock experiments on a two-stage gas gun. This system relies upon measuring the total system response using a tunable monochromator, a NIST-traceable calibrated power meter, and a SSP. The diagnostic performance is validated against previously measured temperatures of shock driven z-cut quartz at 99 GPa and 93 GPa. The results are found to agree with the literature to within 5% and are discussed in this manuscript. The experimental setup utilizes measurements from multiple SSP systems per sample, providing several independent measurements and substantially increasing confidence in the extrapolated shock driven sample temperature.
A simple combination of the Planck blackbody emission law, optical filters, and digital image processing is demonstrated to enable most commercial color cameras (still and video) to be used as an imaging pyrometer for flames and explosions. The hardware and data processing described take advantage of the color filter array (CFA) that is deposited on the surface of the light sensor array present in most digital color cameras. In this work, a triple-pass optical filter incorporated into the camera lens allows light in three 10-nm wide bandpass regions to reach the CFA/light sensor array. These bandpass regions are centered over the maxima in the blue, green, and red transmission regions of the CFA, minimizing the spectral overlap of these regions normally present. A computer algorithm is used to retrieve the blue, green, and red image matrices from camera memory and correct for remaining spectral overlap. A second algorithm calibrates the corrected intensities to a gray body emitter of known temperature, producing a color intensity correction factor for the camera/filter system. The Wien approximation to the Planck blackbody emission law is used to construct temperature images from the three color (blue, green, red) matrices. A short pass filter set eliminates light of wavelengths longer than 750 nm, providing reasonable accuracy (±10%) for temperatures between 1200 and 6000 K. The effectiveness of this system is demonstrated by measuring the temperature of several systems for which the temperature is known.
In support of the SPRITE concept, an integrated data acquisition system has been developed and fabricated for preliminary testing. The data acquisition system has been designed to condition traditional thermal protection system sensors, store their data to an on-board memory card, and in parallel, telemeter to an external system. In the fall of 2010, this system was integrated into a 14 in. diameter, 45 degree sphere cone probe instrumented with thermal protection system sensors. This system was then tested at the NASA Ames Research Center Aerodynamic Heating Facility's arc jet at approximately 170 W/sq. cm. The first test in December 2010 highlighted hardware design issues that were redesigned and implemented leading to a successful test in February 2011.
Spacecraft instruments requiring cryocoolers to cool infrared focal plane arrays have increased complexity in the overall design with respect to power and thermal management. The use of mechanical cryocoolers can drive spacecraft solar array and thermal radiator dimensions, and spacecraft designers must be prudent in keeping these dimensions feasible.
Multiconductor instrumentation cable in which the conducting wires are routed through two concentric copper tube sheaths, employing a compressed insulator between the conductors and between the inner and outer sheaths, is durable and easily installed in high thermal or nuclear radiation area. The double sheath is a barrier against moisture, abrasion, and vibration.
Lunar regolith was first identified as a major issue during the Apollo missions. There are three characteristics of lunar regolith that cause issues: small size, rough surface topology, and particle charge. The lack of a magnetosphere allows solar wind to bombard the lunar surface which results in the regolith becoming charged. These charged particles will adhere to most surfaces and degrade thermal control systems. The lack of atmosphere and surface liquids on the lunar surface results in the regolith’s small particle size, abrasive topology, and chemical reactivity. When the dust adheres to a surface, or the lunar lander exhaust kicks up dust regolith, it can impair instrument or thermal control system function, abrade optical surfaces, diminish range of motion, and mechanically degrade soft-goods such as gloves and astronaut suits. All of these examples were experienced during the Apollo missions. The objective of this study was to select and evaluate abrasion-resistant ceramic materials for application on the lunar landing equipment.