SEARCH · Engineering Papers

Results for “cold chamber”

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.

SourceResearch connectionFrom yearThrough year

Quote a phrase for an exact phrase match. Source license links do not imply unrestricted reuse.

At least 91 records · Page 5

All sources NASA NTRS DOE OSTI

DOE OSTI2024

The Significant Role of New Particle Composition and Morphology on the HNO 3 -Driven Growth of Particles down to Sub-10 nm

New particle formation and growth greatly influence air quality and the global climate. Recent CERN Cosmics Leaving OUtdoor Droplets (CLOUD) chamber experiments proposed that in cold urban atmospheres with highly supersaturated HNO 3 and NH 3 , newly formed sub-10 nm nanoparticles can grow rapidly (up to 1000 nm h –1 ). Here, we present direct observational evidence that in winter Beijing with persistent highly supersaturated HNO 3 and NH 3 , nitrate contributed less than ~14% of the 8–40 nm nanoparticle composition, and overall growth rates were only ~0.8–5 nm h –1 . To explain the observed growth rates and particulate nitrate fraction, the effective mass accommodation coefficient of HNO 3 (α HNO 3 ) on the nanoparticles in urban Beijing needs to be 2–4 orders of magnitude lower than those in the CLOUD chamber. We propose that the inefficient uptake of HNO 3 on nanoparticles is mainly due to the much higher particulate organic fraction and lower relative humidity in urban Beijing. To quantitatively reproduce the observed growth, we show that an inhomogeneous “inorganic core–organic shell” nanoparticle morphology might exist for nanoparticles in Beijing. Finally, this study emphasized that growth for nanoparticles down to sub-10 nm was largely influenced by their composition, which was previously ignored and should be considered in future studies on nanoparticle growth.

54 ENVIRONMENTAL SCIENCES↗

NASA NTRS1972

The design and operation of a planet emitted and albedo radiation simulator

A simulator is described, which provides a dynamic simulation of planetary albedo and planet-emitted thermal radiation. The simulator was designed for model tests of advanced thermal control concepts, the development of space hardware, and flight qualification of small payloads. The simulator was designed to be used in a vacuum chamber equipped with liquid nitrogen cold walls and a solar simulator. The simulator consists of two major components: a gimbaled model support system and an array of quartz infrared lamps. Tests indicate that the apparatus will satisfactorily simulate the secondary radiation on a 0.92-meter spherical body in near-earth orbits. Calculations indicate that other shapes can be accommodated without structural modifications.

Sweet, G. E.↗

NASA NTRS1984

Ignition characterization of the GOX/ethanol propellant combination

This paper describes the results of a study to define the ignition characteristics and thruster pulse mode capabilities of the GOX/ethanol propellant combination. Ignition limits were defined in terms of mixture ratio and cold flow pressure using a spark initiated torch igniter. Igniter tests were run over a wide range of cold flow pressure, propellant temperature and mixture ratio. The product of cold flow pressure and igniter chamber diameter was used to correlate mixture ratio regimes of ignition and nonignition. Engine ignition reliability and pulse mode capability were demonstrated using a 620 lbF thruster with an integrated torch igniter. The nominal chamber pressure and mixture ratio were 150 psia and 1.8, respectively, thruster tests were run over a wide range of chamber pressures and mixture ratios. The feasibility of thruster pulse mode operation with the non-hypergolic GOX/ethanol propellant combination was demonstrated.

Lawver, B. R.↗

NASA NTRS1995

Performance of the Satellite Test Assistant Robot in JPL's Space Simulation Facility

An innovative new telerobotic inspection system called STAR (the Satellite Test Assistant Robot) has been developed to assist engineers as they test new spacecraft designs in simulated space environments. STAR operates inside the ultra-cold, high-vacuum, test chambers and provides engineers seated at a remote Operator Control Station (OCS) with high resolution video and infrared (IR) images of the flight articles under test. STAR was successfully proof tested in JPL's 25-ft (7.6-m) Space Simulation Chamber where temperatures ranged from +85 C to -190 C and vacuum levels reached 5.1 x 10(exp -6) torr. STAR's IR Camera was used to thermally map the entire interior of the chamber for the first time. STAR also made several unexpected and important discoveries about the thermal processes occurring within the chamber. Using a calibrated test fixture arrayed with ten sample spacecraft materials, the IR camera was shown to produce highly accurate surface temperature data. This paper outlines STAR's design and reports on significant results from the thermal vacuum chamber test.

Mcaffee, Douglas↗

NASA NTRS1987

Development of a high-performance cryogenic radiator with V-groove radiation shields

A feasibility demonstration model of a new-technology, high-performance, cryogenic radiative cooler has been successfully tested in a helium-cooled vacuum chamber. Thermal isolation of the radiator cold stage from the warm spacecraft and instrument is achieved by a novel arrangement of lightweight radiation shields which form large V-groove cavities, and by the use of low-conductance structural supports. The vacuum chamber experiments demonstrate the thermal isolation capability of the radiator design. Analytical model predictions are all within 1 K of the measured test temperatures. Utilization of this design enables reasonably sized radiative coolers (less than 0.3 sq m) to achieve operating temperatures below 60 K at earth, Mars, and beyond for useful cooling loads below 75 mW. The design can be scaled to accommodate larger heat loads. The reduction in operating temperature of over 15 K relative to any radiative cooler flown in the past will allow the V-groove isolation radiator to meet the challenging instrument cooling requirements of many future space missions.

Bard, Steven↗

NASA NTRS2015

Propulsion System Testing for the Iodine Satellite (iSAT) Demonstration Mission

CUBESATS are relatively new spacecraft platforms that are typically deployed from a launch vehicle as a secondary payload, providing low-cost access to space for a wide range of end-users. These satellites are comprised of building blocks having dimensions of 10x10x10 cm cu and a mass of 1.33 kg (a 1-U size). While providing low-cost access to space, a major operational limitation is the lack of a propulsion system that can fit within a CubeSat and is capable of executing high delta v maneuvers. This makes it difficult to use CubeSats on missions requiring certain types of maneuvers (i.e. formation flying, spacecraft rendezvous). Recently, work has been performed investigating the use of iodine as a propellant for Hall-effect thrusters (HETs) 2 that could subsequently be used to provide a high specific impulse path to CubeSat propulsion. 3, 4 Iodine stores as a dense solid at very low pressures, making it acceptable as a propellant on a secondary payload. It has exceptionally high ρIsp (density times specific impulse), making it an enabling technology for small satellite near-term applications and providing the potential for systems-level advantages over mid-term high power electric propulsion options. Iodine flow can also be thermally regulated, subliming at relatively low temperature (less than 100 C) to yield I2 vapor at or below 50 torr. At low power, the measured performance of an iodine-fed HET is very similar to that of a state-of-the-art xenon-fed thruster. Just as importantly, the current-voltage discharge characteristics of low power iodine-fed and xenon-fed thrusters are remarkably similar, potentially reducing development and qualifications costs by making it possible to use an already-qualified xenon-HET PPU in an iodine-fed system. Finally, a cold surface can be installed in a vacuum test chamber on which expended iodine propellant can deposit. In addition, the temperature doesn't have to be extremely cold to maintain a low vapor pressure in the vacuum chamber (it is under 10(exp -6) torr at −75 C), making it possible to 'cryopump' the propellant with lower-cost recirculating refrigerant-based systems as opposed to using liquid nitrogen or low temperature gaseous helium cryopanels. An iodine-based system is not without its challenges. The primary challenge is that the entire feed system must be maintained at an elevated temperature to prevent the iodine from depositing (transitioning from the gas phase directly back into the solid phase), which will block the propellant feed lines. Furthermore, deposition will occur unless the temperature in the lines is not greater than the temperature of the propellant reservoir. The flow rate can be controlled by adjusting the heating applied to the reservoir, but as with any thermal control there is a relatively slow response to changes in the heating rate. In the present paper, we describe the propulsion and propellant feed system for the iodine satellite (iSAT) flight demonstration mission. The system is based around the Busek BHT-200 Hall thruster, which has been modified for chemical compatibility with iodine vapor. While the gross propellant flow rate is maintained by the heated propellant reservoir, the flow to the anode and cathode are adjusted using two heated Vacco proportional flow control valves (PFCV), which provide very fast response on the flow rate adjustment. The flight mission design layout will be presented, showing how the system will be packaged into the overall 12-U spacecraft and the techniques being employed to protect the remaining spacecraft hardware from the propulsion system (e.g., plasma impingement, iodine deposition, thermal loads). In addition to the flight system design, results of testing the thruster and cathode with both operating on iodine propellant are presented. The tests are conducted on a thrust stand (see Fig. 1) in a large vacuum chamber containing a beam dump chilled to below -100 C to 'cryopump' the propellant. The thruster performance during these tests is presented, with these data used to evaluate the feed system and guide further refinements. Results of relatively long duration testing are presented to demonstrate the capability to operate for the length of the iSAT mission and to perform a number of re-starts as will be required by the mission concept of operations.

Polzin, Kurt A.↗

NASA NTRS2015

Hardware in the Loop Testing of an Iodine-Fed Hall Thruster

CUBESATS are relatively new spacecraft platforms that are typically deployed from a launch vehicle as a secondary payload,1 providing low-cost access to space for a wide range of end-users. These satellites are comprised of building blocks having dimensions of 10x10x10 cm cu and a mass of 1.33 kg (a 1-U size). While providing low-cost access to space, a major operational limitation is the lack of a propulsion system that can fit within a CubeSat and is capable of executing high delta v maneuvers. This makes it difficult to use CubeSats on missions requiring certain types of maneuvers (i.e. formation flying, spacecraft rendezvous). Recently, work has been performed investigating the use of iodine as a propellant for Hall-effect thrusters (HETs) 2 that could subsequently be used to provide a high specific impulse path to CubeSat propulsion. Iodine stores as a dense solid at very low pressures, making it acceptable as a propellant on a secondary payload. It has exceptionally high ρIsp (density times specific impulse), making it an enabling technology for small satellite near-term applications and providing the potential for systems-level advantages over mid-term high power electric propulsion options. Iodine flow can also be thermally regulated, subliming at relatively low temperature ( less than100 C) to yield I2 vapor at or below 50 torr. At low power, the measured performance of an iodine-fed HET is very similar to that of a state-of-the-art xenon-fed thruster. Just as importantly, the current-voltage discharge characteristics of low power iodine-fed and xenon-fed thrusters are remarkably similar, potentially reducing development and qualifications costs by making it possible to use an already-qualified xenon-HET PPU in an iodine-fed system. Finally, a cold surface can be installed in a vacuum test chamber on which expended iodine propellant can deposit. In addition, the temperature doesn't have to be extremely cold to maintain a low vapor pressure in the vacuum chamber (it is under 10(exp -6) torr at -75 C), making it possible to 'cryopump' the propellant with lower-cost recirculating refrigerant-based systems as opposed to using liquid nitrogen or low temperature gaseous helium cryopanels. In the present paper, we describe testing performed using an iodine-fed 200 W Hall thruster mounted to a thrust stand and operated in conjunction with MSFCs Small Projects Rapid Integration and Test Environment (SPRITE) Portable Hardware In the Loop (PHIL) hardware. This work is performed in support of the iodine satellite (iSAT) project, which aims to fly a 200-W iodine-fed thruster on a 12-U CubeSat. The SPRITE PHIL hardware allows a given vehicle to do a checkout of its avionics algorithm by allowing it to monitor and feed data to simulated sensors and effectors in a digital environment. These data are then used to determine the attitude of the vehicle and a separate computer is used to interpret the data set and visualize it using a 3D graphical interface. The PHIL hardware allows the testing of the vehicles bus by providing 'real' hardware interfaces (in the case of this test a real RS422 bus) and specific components can be modeled to show their interactions with the avionics algorithm (e.g. a thruster model). For the iSAT project the PHIL is used to visualize the operating cycle of the thruster and the subsequent effect this thrusting has on the attitude of the satellite over a given period of time. The test is controlled using software running on an Andrews Space Cortex 160 flight computer. This computer is the current baseline for a full iSAT mission. While the test could be conducted with a lab computer and software, the team chose to exercise the propulsion system with a representative CubeSat-class computer. For purposes of this test, the "flight" software monitored the propulsion and PPU systems, controlled operation of the thruster, and provided thruster state data to the PHIL simulation. Commands to operate the thruster were initiated from an operator's workstation outside the vacuum chamber and passed through the Cortex 160 to exercise portions of the flight avionics. Two custom-designed pieces of electronics hardware have been designed to operate the propellant feed system. One piece of hardware is an auxiliary board that controls a latch valve, proportional flow control valves (PFCVs) and valve heaters as well as measuring pressures, temperatures and PFCV feedback voltage. An onboard FPGA provides a serial link for issuing commands and manages all lower level input-output functions. The other piece of hardware is a power distribution board, which accepts a standard bus voltage input and converts this voltage into all the different current-voltage types required to operate the auxiliary board. These electronics boards are located in the vacuum chamber near the thruster, exposing this hardware to both the vacuum and plasma environments they would encounter during a mission, with these components communicating to the flight computer through an RS-422 interface. The auxiliary board FPGA provides a 28V MOSFET switch circuit with a 20ms pulse to open or close the iodine propellant feed system latch valve. The FPGA provides a pulse width modulation (PWM) signal to a DC/DC boost converter to produce the 12-120V needed for control of the proportional flow control valve. There are eight MOSFET-switched heating circuits in the system. Heaters are 28V and located in the latch valve, PFCV, propellant tank and propellant feed lines. Both the latch valve and PFCV have thermistors built into them for temperature monitoring. There are also seven resistance temperature device (RTD) circuits on the auxiliary board that can be used to measure the propellant tank and feedline temperatures. The signals are conditioned and sent to an analog to digital converter (ADC), which is directly commanded and controlled by the FPGA.

Polzin, Kurt A.↗

NASA NTRS2016

Engineering Model Propellant Feed System Development for an Iodine Hall Thruster Demonstration Mission

CUBESATS are relatively new spacecraft platforms that are typically deployed from a launch vehicle as a secondary payload, providing low-cost access to space for a wide range of end-users. These satellites are comprised of building blocks having dimensions of 10x10x10 cu cm and a mass of 1.33 kg (a 1-U size). While providing low-cost access to space, a major operational limitation is the lack of a propulsion system that can fit within a CubeSat and is capable of executing high (Delta)v maneuvers. This makes it difficult to use CubeSats on missions requiring certain types of maneuvers (i.e. formation flying, spacecraft rendezvous). Recently, work has been performed investigating the use of iodine as a propellant for Hall-effect thrusters (HETs) 2 that could subsequently be used to provide a high specific impulse path to CubeSat propulsion. 3, 4 Iodine stores as a dense solid at very low pressures, making it acceptable as a propellant on a secondary payload. It has exceptionally high ρIsp (density times specific impulse), making it an enabling technology for small satellite near-term applications and providing the potential for systems-level advantages over mid-term high power electric propulsion options. Iodine flow can also be thermally regulated, subliming at relatively low temperature (< 100 C) to yield I2 vapor at or below 50 torr. At low power, the measured performance of an iodine-fed HET is very similar to that of a state-of-the-art xenon-fed thruster. Just as importantly, the current-voltage discharge characteristics of low power iodine-fed and xenon-fed thrusters are remarkably similar, potentially reducing development and qualifications costs by making it possible to use an already-qualified xenon-HET PPU in an iodine-fed system. Finally, a cold surface can be installed in a vacuum test chamber on which expended iodine propellant can deposit. In addition, the temperature doesn't have to be extremely cold to maintain a low vapor pressure in the vacuum chamber (it is under 10(exp −6) torr at −75 C), making it possible to 'cryopump' the propellant with lower-cost recirculating refrigerant-based systems as opposed to using liquid nitrogen or low temperature gaseous helium cryopanels. In the present paper, we describe the design and testing of the engineering model propellant feed system for iSAT (see Fig. 1). The feed system is based around an iodine propellant reservoir and two proportional control valves (PFCVs) that meter the iodine flow to the cathode and anode. The flow is split upstream of the PFCVs to both components can be fed from a common reservoir. Testing of the reservoir is reported to demonstrate that the design is capable of delivering the required propellant flow rates to operate the thruster. The tubing and reservoir are fabricated from hastelloy to resist corrosion by the heated gaseous iodine propellant. The reservoir, tubing, and PFCVs are heated to ensure the sublimed propellant will not re-deposit within the feed system. Heating is accomplished through a number of individual zones to control the overall power expended on heating the system and insulation is employed to minimize the amount of power used to heat the system prior to thruster operation.

Polzin, Kurt A.↗

NASA NTRS2001

Evaluation of Impinging Stream Vortex Chamber Concepts for Liquid Rocket Engine Applications

Trinh, Huu P.↗

NASA NTRS2002

Evaluation of Impinging Stream Vortex Chamber Concepts for Liquid Rocket Engine Applications

To pursue technology developments for future launch vehicles, NASA/Marshall Space Flight Center (MSFC) is examining vortex chamber concepts for liquid rocket engine applications. Past studies indicated that the vortex chamber schemes potentially have a number of advantages over conventional chamber methods. Due to the nature of the vortex flow, relatively cooler propellant streams tend to flow along the chamber wall. Hence, the thruster chamber can be operated without the need of any cooling techniques. This vortex flow also creates strong turbulence, which promotes the propellant mixing process. Consequently, the subject chamber concepts not only offer system simplicity, but also enhance the combustion performance. Test results have shown that chamber performance is markedly high even at a low chamber length-to-diameter ratio (LD). This incentive can be translated to a convenience in the thrust chamber packaging. Variations of the vortex chamber concepts have been introduced in the past few decades. These investigations include an ongoing work at Orbital Technologies Corporation (ORBITEC). By injecting the oxidizer tangentially at the chamber convergence and fuel axially at the chamber head end, Knuth et al. were able to keep the wall relatively cold. A recent investigation of the low L/D vortex chamber concept for gel propellants was conducted by Michaels. He used both triplet (two oxidizer orifices and one fuel orifice) and unlike impinging schemes to inject propellants tangentially along the chamber wall. Michaels called the subject injection scheme an Impinging Stream Vortex Chamber (ISVC). His preliminary tests showed that high performance, with an Isp efficiency of 9295, can be obtained. MSFC and the U. S. Army are jointly investigating an application of the ISVC concept for the cryogenic oxygen/hydrocarbon propellant system. This vortex chamber concept is currently tested with gel propellants at AMCOM at Redstone Arsenal, Alabama. A version of this concept for the liquid oxygen (LOX) hydrocarbon fuel (RP-1) system has been derived from the one for the gel propellant. An unlike impinging injector was employed to deliver the propellants to the chamber. MSFC is also conducting an alternative injection scheme, called the chasing injector, associated with this vortex chamber concept. In this injection technique, both propellant jets and their impingement point are in the same chamber cross-sectional plane. Long duration tests (approximately up to 15 seconds) will be conducted on the ISVC to study the thermal effects. This paper will report the progress of the subject efforts at NASA Marshall Space Flight Center. Thrust chamber performance and thermal wall compatibility will be evaluated. The chamber pressures, wall temperatures, and thrust will be measured as appropriate. The test data will be used to validate CFD models, which, in turn, will be used to design the optimum vortex chambers. Measurements in the previous tests showed that the chamber pressures vary significantly with radius. This is due to the existence of the vortices in the chamber flow field. Hence, the combustion efficiency may not be easily determined from chamber pressure. For this project, measured thrust data will be collected. The performance comparison will be in terms of specific impulse efficiencies. In addition to the thrust measurements, several pressure and temperature readings at various locations on the chamber head faceplate and the chamber wall will be made. The first injector and chamber were designed and fabricated based on the available data and experience gained during gel propellant system tests by the U.S. Army. The alternate injector for the ISVC was also fabricated. Hot-fire tests of the vortex chamber are about to start and are expected to complete in February of 2003 at the TS115 facility of MSFC.

Trinh, Huu P.↗

NASA NTRS1971

Heat pipe technology for advanced rocket thrust chambers

The application of heat pipe technology to the design of rocket engine thrust chambers is discussed. Subjects presented are: (1) evaporator wick development, (2) specific heat pipe designs and test results, (3) injector design, fabrication, and cold flow testing, and (4) preliminary thrust chamber design.

Rousar, D. C.↗

NASA NTRS1992

A new approach for performing contamination control bakeouts in JPL thermal vacuum test chambers

Contamination control requirements for the Wide Field/Planetary Camera II (WF/PC II) are necessarily stringent to protect against post-launch contamination of the sensitive optical surfaces, particularly the cold charge coupled device (CCD) imaging surfaces. Typically, thermal vacuum test chambers have employed a liquid nitrogen (LN2) cold trap to collect outgassed contaminants. This approach has the disadvantage of risking recontamination of the test article from shroud offgassing during post-test warmup of the chamber or from any shroud warming of even a few degrees during the bakeout process. By using an enclave, essentially a chamber within a chamber, configured concentrically and internally within an LN2 shroud, a method was developed, based on a design concept by Taylor, for preventing recontamination of test articles during bakeouts and subsequent post-test warmup of the vacuum chamber. Enclaves for testing WF/PC II components were designed and fabricated, then installed in three of JPL's Environmental Test Lab chambers. The design concepts, operating procedures, and test results of this development are discussed.

Johnson, Kenneth R.↗

NASA NTRS2013

Cold-end Subsystem Testing for the Fission Power System Technology Demonstration Unit

The Fission Power System (FPS) Technology Demonstration Unit (TDU) consists of a pumped sodium-potassium (NaK) loop that provides heat to a Stirling Power Conversion Unit (PCU), which converts some of that heat into electricity and rejects the waste heat to a pumped water loop. Each of the TDU subsystems is being tested independently prior to full system testing at the NASA Glenn Research Center. The pumped NaK loop is being tested at NASA Marshall Space Flight Center; the Stirling PCU and electrical controller are being tested by Sunpower Inc.; and the pumped water loop is being tested at Glenn. This paper describes cold-end subsystem setup and testing at Glenn. The TDU cold end has been assembled in Vacuum Facility 6 (VF 6) at Glenn, the same chamber that will be used for TDU testing. Cold-end testing in VF 6 will demonstrate functionality; validated cold-end fill, drain, and emergency backup systems; and generated pump performance and system pressure drop data used to validate models. In addition, a low-cost proof-of concept radiator has been built and tested at Glenn, validating the design and demonstrating the feasibility of using low-cost metal radiators as an alternative to high-cost composite radiators in an end-to-end TDU test.

Briggs, Maxwell↗

NASA NTRS2013

Cold-End Subsystem Testing for the Fission Power System Technology Demonstration Unit

The Fission Power System (FPS) Technology Demonstration Unit (TDU) consists of a pumped sodiumpotassium (NaK) loop that provides heat to a Stirling Power Conversion Unit (PCU), which converts some of that heat into electricity and rejects the waste heat to a pumped water loop. Each of the TDU subsystems is being tested independently prior to full system testing at the NASA Glenn Research Center. The pumped NaK loop is being tested at NASA Marshall Space Flight Center; the Stirling PCU and electrical controller are being tested by Sunpower Inc.; and the pumped water loop is being tested at Glenn. This paper describes cold-end subsystem setup and testing at Glenn. The TDU cold end has been assembled in Vacuum Facility 6 (VF 6) at Glenn, the same chamber that will be used for TDU testing. Cold-end testing in VF 6 will demonstrate functionality; validated coldend fill, drain, and emergency backup systems; and generated pump performance and system pressure drop data used to validate models. In addition, a low-cost proof-of concept radiator has been built and tested at Glenn, validating the design and demonstrating the feasibility of using low-cost metal radiators as an alternative to highcost composite radiators in an end-to-end TDU test.

Briggs, Mazwell↗

NASA NTRS1974

Formation of ion-acoustic shocks

The formation of an ion-acoustic shock in a numerically modeled two-chamber double plasma device is investigated for a plasma of cold ions and isothermal Boltzmann electrons. An initial potential ramp applied to the driver chamber launches an ion-acoustic pulse into the target chamber which steepens into a shock. The quasi-steady shock structure agrees with observed double plasma shocks. An upper limit of Mach 1.6 is observed independent of the potential ramp magnitude, in agreement with theory.

White, R. B.↗

DOE OSTI2021

Characterization and QC practice of 16-channel ADC ASIC at cryogenic temperature for Liquid Argon TPC front-end readout electronics system in DUNE experiment

ColdADC is a low-noise 16-channel analog-to-digital converter ASIC designed for cold readout electronics of Liquid Argon Time Projection Chambers (LArTPCs) in the Deep Underground Neutrino Experiment (DUNE). ColdADC was specifically designed for operation at cryogenic temperatures (77 K–89 K). Cold electronics is considered to be an enabling technology for liquid argon detectors in neutrino experiments. The main function of the chip is to digitize signals from the 16-channel charge-sensitive amplifier designed at BNL (LArASIC) and send digitized signals to the data aggregator and serializer chip (COLDATA). ColdADC operates with a resolution of 12 bits and a sampling rate of 2 MS/s per channel. A complete characterization of prototype chips was conducted at cryogenic temperature, involving power consumption, noise and linearity performance. In the first DUNE Far-Detector module, 384000 data channels will be read out, corresponding to 24000 ColdADC chips. Therefore, systematic testing of ColdADC chips at cryogenic temperature is vital for the quality and reliability of DUNE Single-Phase modules. We have developed a Quality Control (QC) test stand that includes ColdADC evaluation and characterization at cryogenic temperature. The QC procedure was applied to a first batch of 33 chips, corresponding to 528 data channels. The obtained results meet all the required DUNE specifications for ColdADC. Moreover, the testing outcome allowed us to identify possibilities of performance optimization that are currently being addressed in the design of the next version of the chip. The Quality Control procedure is proved to be successful and will be a reference design for future large-batch production testing of final ColdADC ASICs.

46 INSTRUMENTATION RELATED TO NUCLEAR SCIENCE AND ↗

NASA NTRS1995

Effects of Slag Ejection on Solid Rocket Motor Performance

In past firings of the Reusable Solid Rocket Motor (RSRM) both static test and flight motors have shown small pressure perturbations occurring primarily between 65 and 80 seconds. A joint NASA/Thiokol team investigation concluded that the cause of the pressure perturbations was the periodic ingestion and ejection of molten aluminum oxide slag from the cavity around the submerged nozzle nose which tends to trap and collect individual aluminum oxide droplets from the approach flow. The conclusions of the team were supported by numerous data and observations from special tests including high speed photographic films, real time radiography, plume calorimeters, accelerometers, strain gauges, nozzle TVC system force gauges, and motor pressure and thrust data. A simplistic slag ballistics model was formulated to relate a given pressure perturbation to a required slag quantity. Also, a cold flow model using air and water was developed to provide data on the relationship between the slag flow rate and the chamber pressure increase. Both the motor and the cold flow model exhibited low frequency oscillations in conjunction with periods of slag ejection. Motor and model frequencies were related to scaling parameters. The data indicate that there is a periodicity to the slag entrainment and ejection phenomena which is possibly related to organized oscillations from instabilities in the dividing streamline shear layer which impinges on the underneath surface of the nozzle.

Whitesides, R. Harold↗

DOE OSTI2020

Kilowatt Reactor Using Stirling TechnologY (KRUSTY) Cold Critical Measurements

For the Kilowatt Reactor Using Stirling TechnologY (KRUSTY) cold critical experiments, the KRUSTY component critical configuration was modified by the addition of parts that would be required for cold, warm, and hot critical experiments (including the vacuum chamber as well as the heat pipes and associated parts). Reactivity measurements were performed on the KRUSTY cold critical experimental configurations with the goal of obtaining reactivity-worth measurements on the beryllium oxide (BeO) reflector and the boron carbide (B 4 C) control rod parts. The resulting data are consistent and allow for accurate identification of the BeO and B 4 C part thicknesses required to achieve the excess reactivity needed for the KRUSTY warm and hot experimental configurations.

22 GENERAL STUDIES OF NUCLEAR REACTORS↗