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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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NASA NTRS1993

Hubble Space Telescope EVA Multisetting Torque Limiter redesign

The Hubble Space Telescope (HST) Multisetting Torque Limiter will be used by astronauts during the First Servicing Mission STS-61. It was designed to replace several individual torque limiters. This replacement reduces the number of tools required as well as reduce the time required for servicing. The Multisetting Torque Limiter design specified five separate torque settings: 2, 5.3, 9, 35 and 45 ft-lbs. This original units were tested at the 2, 5.3, and 9 ft-lbs settings and did not meet the design specifications. The 35 ft-lb and 45 ft-lb torque settings were not incorporated into the original flight design. The objective of the Goddard Space Flight Center (GSFC) redesign effort is to incorporate all five torque settings into a flight design which means the requirements for the first HST Servicing Mission. The first target of the redesign effort focused on the spring chamber and ball detent design. The original design used belleville springs in parallel; the new design changed this configuration to belleville springs stacked in series. The current engineering unit was successfully tested in ambient environment at all five settings. A flight design will soon undergo thermal vacuum testing to evaluate its performance at the temperature extremes.

Richards, Paul W.↗

NASA NTRSDate not supplied

A Summary of the NEXT-C Flight Thruster Proto-flight Testing

Aerojet Rocketdyne (AR), under contract to NASA Glenn Research Center, built, tested and delivered the first flight version of the NEXT-C (NASA Evolutionary Xenon Thruster – Commercial) engine system. This system is comprised of a 7 kW thruster and power processing unit which are the commercial variants of the NEXT prototype thruster and power processing unit. In addition to developing the NEXT-C engine system for potential future NASA science missions, AR is dedicated to ensuring that this propulsion system is viable for commercial users and a wider variety of commercial applications. This paper will focus on the performance of the first flight model thruster, SN001, and presents an overview of the proto-flight testing environments and data obtained. For this first flight thruster, the proto-flight test sequence involved characterizing the performance of the thruster, in addition to vibration and thermal vacuum testing of the thruster. Between each of these environmental tests, the performance of the thruster was verified, ensuring that the environmental exposures, i.e. vibration and TVAC, did not affect the thruster performance. Specifically, the proto-flight data obtained included discharge chamber, optics and neutralizer performance. Faraday probe and ExB data obtained in the regions downstream of the thruster are presented and used to calculate thrust, and therefore specific impulse and total efficiency. Lastly, data are presented showing how the centroid of the ion beam, assumed to be indicative of the thrust vector location, varies throughout the entire proto-flight test campaign. The proto-flight testing of the NEXT-C PPU is documented in a previously published paper by Bontempo4. Following the proto-flight testing of the thruster, the thruster and its PPU successfully completed a System Integrated Test (SIT) tailored to the DART Mission. After the completion of the proto-flight testing, the SN001 thruster and SN002 PPU were delivered to APL for use on the DART mission.

Ion thruster↗

NASA NTRS2015

Thermal Testing and Integration: Magnetospheric MultiScale (MMS) Observatories with Digital 1-Wire Sensors

Thermocouples require two thin wires to be routed out of the spacecraft to connect to the ground support equipment used to monitor and record the temperature data. This large number of wires that exit the observatory complicates integration and creates an undesirable heat path during testing. These wires exiting the spacecraft need to be characterized as a thermal short that will not exist during flight. To minimize complexity and reduce thermal variables from these ground support equipment (GSE) wires, MMS pursued a hybrid path for temperature monitoring, utilizing thermocouples and digital 1-wire temperature sensors. Digital 1-wire sensors can greatly reduce harness mass, length and complexity as they can be spliced together. For MMS, 350 digital 1-wire sensors were installed on the spacecraft with only 18 wires exiting as opposed to a potential 700 thermocouple wires. Digital 1-wire sensors had not been used in such a large scale at NASAGSFC prior to the MMS mission. During the MMS thermal vacuum testing a lessons learned matrix was formulated that will assist future integration of 1-wires into thermal testing and one day into flight.

Sensors↗

NASA NTRS2006

Solar Sails

The Solar Sail Propulsion investment area has been one of the three highest priorities within the In-Space Propulsion Technology (ISPT) Project. In the fall of 2003, the NASA Headquarters' Science Mission Directorate provided funding and direction to mature the technology as far as possible through ground research and development from TRL 3 to 6 in three years. A group of experts from government, industry, and academia convened in Huntsville, Alabama to define technology gaps between what was needed for science missions to the inner solar system and the current state of the art in ultra1ightweight materials and gossamer structure design. This activity set the roadmap for development. The centerpiece of the development would be the ground demonstration of scalable solar sail systems including masts, sails, deployment mechanisms, and attitude control hardware and software. In addition, new materials would be subjected to anticipated space environments to quantify effects and assure mission life. Also, because solar sails are huge structures, and it is not feasible to validate the technology by ground test at full scale, a multi-discipline effort was established to develop highly reliable analytical models to serve as mission assurance evidence in future flight program decision-making. Two separate contractor teams were chosen to develop the SSP System Ground Demonstrator (SGD). After a three month conceptual mission/system design phase, the teams developed a ten meter diameter pathfinder set of hardware and subjected it to thermal vacuum tests to compare analytically predicted structural behavior with measured characteristics. This process developed manufacturing and handling techniques and refined the basic design. In 2005, both contractor teams delivered 20 meter, four quadrant sail systems to the largest thermal vacuum chamber in the world in Plum Brook, Ohio, and repeated the tests. Also demonstrated was the deployment and articulation of attitude control mechanisms. In conjunction with these tests, the stowed sails were subjected to launch vibration and ascent vent tests. Other investments studied radiation effects on the solar sail materials, investigated spacecraft charging issues, developed shape measuring techniques and instruments, produced advanced trajectory modeling capabilities, and identified and resolved gossamer structure dynamics issues. Technology validation flight and application to a He1iophysics science mission is on the horizon.

Young, Roy↗

NASA NTRS2000

Geoscience Laser Altimeter System (GLAS) Final Test Report of DM LHP TV Testing

Two loop heat pipes (LHPs) are to be used for thermal control of the Geoscience Laser Altimeter System (GLAS), planned for flight in 2001. One LHP will be used to transport 100 W from a laser to the radiator, the other will transport 210 W from electronic boxes to the radiator. In order to verify the LHP design for the GLAS application, an LHP Development Model has been fabricated, and ambient and thermal vacuum tested. Two aluminum blocks of 15 kg and 30 kg, respectively, were attached to the LHP to simulate the thermal masses connected to the heat sources. A 20 W starter heater was installed on the evaporator to aid the loop startup. A new concept to thermally couple the vapor and liquid line was also incorporated in the LHP design. Such a thermal coupling would reduce the power requirement on the compensation chamber in order to maintain the loop set point temperature. To avoid freezing of the liquid in the condenser during cold cases, propylene was selected as the working fluid. The LHP was tested under reflux mode and with adverse elevation. Tests conducted included start-up, power cycle, steady state and transient operation during hot and cold cases, and heater power requirements for the set point temperature control of the LHP. Test results showed very successful operation of the LHP under all conditions. The 20 W starter heater proved necessary in order to start the loop when a large thermal mass was attached to the evaporator. The thermal coupling between the liquid line and the vapor line significantly reduced the heater power required for loop temperature control, which was less than 5 watts in all cases, including a cold radiator. The test also demonstrated successful operation with a propylene working fluid, with successful startups with condenser temperatures as low as 100 C. Furthermore, the test demonstrated accurate control of the loop operating temperature within +/- 0.2 C, and a successful shutdown of the loop during the survival mode of operation.

Baker, Charles↗

NASA NTRSDate not supplied

Carbothermal Reduction Demonstration: Laser Driven Reaction in a Thermal-Vacuum Environment and Project Status

Introduction: Lunar regolith is approximately 45% oxygen by mass. The majority of the oxygen is bound in silicate minerals. The carbothermal reduction process has been proven to be effective at removing oxygen from lunar regolith simulants [1]. The Carbothermal Reduction Demonstration (CaRD) project aims to increase the Technology Readiness Level (TRL) of a combined solar concentrator and carbothermal reduction systemin order to demonstrate this technology on the lunar surface. The CaRD project is divided into two design cycles, a brass board and prototype. The status of both design cycles will be discussed, as well as concepts for how this technology can be applied to the Artemis program in the future. Brassboard Vacuum Test: For the brassboard demonstration, a 2 kW Nd-YAG laser was used to heat lunar regolith simulant within a carbothermal reactor developed by Sierra Space. The reactor was placed inside of a 15 ft thermal vacuum chamber at the Johnson Space Center. The resulting reaction products were analyzed using both a gas chromotograph and mass spectrometer provided by Kennedy Space Center. Thermal data was also collected. Prototype Design: For the prototype, the CaRD team will perform another thermal vacuum test at JSC using the same interfaces and assets developed for the brassboard but will test a new carbothermal reactor design that Sierra Space is developing through the Carbothermal Oxygen Production Reactor (COPR) Tipping Point project that will include a means to autonomously move regolith in and out of the reactor. In addition, a deployable solar concentrator is being developed by Glenn Research Center using mirrors produced by Carbon Mirror Applications. The solar concentrator will be used to deliver solar energy into a carbothermal reactor to melt regolith and extract oxygen. Avionics and software for the concentrator are being developed by Kennedy Space Center. Future Concepts: An ongoing task within the CaRD project is to update models that can be used to determine the mass, power, and size of In-Situ Resource Utilization concepts at various scales. These models can now be used to analyze alternatives for future applications based on the latest available data.

A J Paz↗

NASA NTRS1976

Development of an inflatable radiator system

Conceptual designs of an inflatable radiator system developed for supplying short duration supplementary cooling of space vehicles are described along with parametric trade studies, materials evaluation/selection studies, thermal and structural analyses, and numerous element tests. Fabrication techniques developed in constructing the engineering models and performance data from the model thermal vacuum tests are included. Application of these data to refining the designs of the flight articles and to constructing a full scale prototype radiator is discussed.

Leach, J. W.↗

NASA NTRS1982

NiCd battery failure analysis

The failure of a nickel cadmium battery undergoing tests is discussed. Reasons for the complete destruction of the battery while undergoing preparation for thermal vacuum testing are given.

Sense, K. A.↗

NASA NTRS2012

The Integration and Test Program of the James Webb Space Telescope

The James Webb Space Telescope (JWST) project has entered into a comprehensive integration and test (I&T) program that over the coming years will assemble the various elements of the observatory (the Optical Telescope Element [OTE], the Integrated Science Instrument Module [ISIM], and the Spacecraft) and verify the readiness of the integrated system for launch. The I&T program as replanned for a 2018 launch readiness date has a number of interesting features. These include a streamlined ISIM cryo-vacuum test program at Goddard Space Flight Center, a streamlined OTIS (OTE + ISIM) test program at Johnson Space Center (JSC), the addition of a second Core cryo-vacuum thermal test, the enhancement of the Pathfinder program at JSC, and enhancement of the subsystem-level testing program for the MIRI cryo-cooler. These latter activities all serve to reduce the risk heading into the end-to-end optical and thermal testing of the telescope at JSC, leading to reduced cost and schedule risk for that critical activity. We report here on the overall I&T program for JWST and on the status of the hardware and plans that support it.

Kimble, Randy↗

NASA NTRS1992

Manned testing in a simulated space environment

A view of the facility and operational requirements involved in performing a manned thermal vacuum test is presented. The requirements fall into two major categories. The first category deals with placing the suited crewmen in a hazardous environment and assuring their safety. The second category deals with the constraints and special requirements involved with a suited crewman operating flight hardware in a 1-G environment. Design areas that deal with man rating a chamber, including fire suppression, emergency repress, emergency power, backups, reliable instrumentation and data systems, communications, television monitoring, biomedical monitoring, material compatibilities, and equipment supporting the Extravehicular Mobility Unit (EMU) are discussed. The operational issues that are peculiar to manned testing such as test rules, test procedures, test protocol, emergency drills, availability of hyperbaric facilities, test team training and certification engineering concerns for a safe mechanical and instrumentation buildup, hazard analysis, and Failure Modes and Effects Analysis are discussed. The constraints and special requirements involved with a suited crewman operating flight hardware in a 1-G environment are addressed.

Fender, Donna L.↗

NASA NTRS2008

Affect of Air Leakage into a Thermal-Vacuum Chamber on Helium Refrigeration Heat Load

NASA s Johnson Space Center (JSC) Building 32 houses two large thermal-vacuum chambers (Chamber A and Chamber B). Within these chambers are liquid nitrogen shrouds to provide a thermal environment and helium panels which operate at 20K to provide cryopumping. Some amount of air leakage into the chambers during tests is inevitable. This causes "air fouling" of the helium panel surfaces due to the components of the air that adhere to the panels. The air fouling causes the emittance of the helium panels to increase during tests. The increase in helium panel emittance increases the heat load on the helium refrigerator that supplies the 20K helium for those panels. Planning for thermal-vacuum tests should account for this increase to make sure that the helium refrigerator capacity will not be exceeded over the duration of a test. During a recent test conducted in Chamber B a known-size air leak was introduced to the chamber. Emittance change of the helium panels and the affect on the helium refrigerator was characterized. A description of the test and the results will be presented.

Garcia, Sam↗

NASA NTRS2020

The Origins Space Telescope Integration and Testing

This paper describes the plans, flows, key facilities, test beds, pathfinders, simulators and ground support equipment that could be used to fully integrate, functionally test and qualify the Origins Space Telescope (Origins). The Origins Observatory consists of the Spacecraft Bus Module (SBM) and the Cryogenic Payload Module (CPM), which comprises the telescope and three science instruments. The telescope is a Three Mirror Anastigmat (TMA) and is composed of four mirrors: three with optical power (the elliptical primary, hyperbolic secondary, and elliptical tertiary mirrors) and a flat field-steering mirror (FSM). The three science instruments spanning the wavelength range 2.8 to 588 μm provide the powerful new spectroscopic and imaging capabilities required to achieve the scientific objectives. The Origins Survey Spectrometer (OSS) uses six gratings in parallel to take multi-beam spectra simultaneously across the 25 to 588 μm window through long slits enabling deep 3D extra-galactic surveys. The Far-IR Imager/Polarimeter (FIP) provides imaging and polarimetric measurement capabilities at 50 and 250 μm. Its fast mapping enables rapid follow-up of transient or variable sources and efficient monitoring campaigns. The Mid-Infrared Spectrometer (MISC-T) provides spectroscopy simultaneously over 2.8 to 20 μm with exquisite stability and precision (<5 ppm between 2.8 to 10 μm, <20 ppm between 11 to 20 μm). All the instruments are delivered for Integration and Test fully qualified and calibrated. The Integration and Test program implemented at each level of assembly is discussed as well as the separation of thermal vacuum testing between the hot and cold zones of the Observatory. Portions of this manuscript appear in the Origins Study Final Report.

Origins↗

NASA NTRS2023

Using Mylar-Insulated Cryopumping Panels to Improve Vacuum Level During Warm Temperature Testing at JSC's Large Thermal Vacuum Facilities

The Johnson Space Center’s Space Environment Simulation Lab (SESL) has both Chamber A, the world’s largest purpose-built thermal vacuum chamber capable of creating deep space conditions, and Chamber B, the largest human rated thermal vacuum chamber. A unique design feature of these chambers is the gaseous helium cryopumping panels within the liquid nitrogen shroud. This shroud is used to bring the chamber to cryogenic temperatures while the cryopumping panels trap gasses using their large surface area in order to create a high vacuum environment of 5*10^-6 Torr. In preparation for the James Webb Space Telescope (JWST) flight test, a series of functionals required the chamber to run at higher temperatures, and therefore did not need active cooling from the liquid nitrogen shroud. During testing, cryopumping panels were used to mitigate contamination during this main shroud warm-up. One of the cryopumping panels in Chamber A was covered with several layers of aluminized mylar in order to thermally protect the zone from the warmed shroud. This strategy was effective and became part of operations during JWST testing. Currently, Chamber B has requests for both commercial and NASA space suit tests at both high and low temperatures to accurately generate thermal models. High temperature tests would greatly benefit from simultaneous shroud warming and the high vacuum environment provided by the cryopumping panels. This paper will quantitatively define the thermal loads on the panels used in previous Chamber A warm up sequences. Additionally, this report will perform a thermal analysis to judge the feasibility of adding layers of aluminized mylar to the cryopumping panels of Chamber B.

Cryogenics↗

NASA NTRSDate not supplied

Analytical Study of a Cryogenic Thermal Control Coating

This study focuses on the analysis of a cryogenic thermal control coating designed to reflect most of the Sun’s irradiance while emitting long wavelength radiation, thus enabling propulsion systems to maintain cryogenic temperatures while in deep space. This “solar white” coating is composed of yttrium oxide (Y2O3) which has properties that are well-suited for cryogenic propellant storage applications. To conduct proof of concept and material development, many tests have been run at both NASA Kennedy Space Center (KSC) and Glenn Research Center (GRC) to demonstrate the performance of the coating in the cryogenic regime in both powder and ceramic forms. At GRC, the Deep Space Solar Simulator (DS3) was used to run thermal vacuum tests where the coating was exposed to a deep space environment (~10 K) with some tests conducted under the illumination of a solar lamp. Using the test data, a thermal model of the DS3 test setup was developed to correlate the test results to predict the optical properties of the coating. Once the optical properties were determined, an additional thermal model, based on previous work by R.C. Young quist et al., was developed to verify the performance of a theoretical lunar, cryogenic propellant depot. The latter thermal model also demonstrates the usage of the HLS-UG-001 Human Landing System Lunar Thermal Analysis Guidebook published to baseline Artemis-related lunar studies.

Cryogenic↗

NASA NTRSDate not supplied

Investigation of Rectifier Diode Failures in the NEXT-C Power Processing Unit

NASA's Evolutionary Xenon Thruster-Commercial (NEXT-C) project is tasked with developing flight electric propulsion systems, including both thrusters and power processing units (PPUs). In 2018, the beam supply in a NEXT-C engineering prototype PPU experienced an output rectifier diode failure during development thermal-vacuum testing. A failure investigation led by NASA GRC identified the root cause of the failure as a thermal runaway caused by increased reverse recovery losses in the diodes when the PPU was run at its maximum operating temperature. Significant reverse recovery performance variations were identified in diodes with the same part number but manufactured by different vendors. The failure investigation was able to collect evidence of the increased reverse recovery and replicate the diode failures in a controlled laboratory environment.

George L Thomas↗

NASA NTRSDate not supplied

Investigation of Rectifier Diode Failures in the NEXT-C Power Processing Unit

George L. Thomas↗

NASA NTRSDate not supplied

Carbothermal Reduction Demonstration (CaRD)

Introduction: Lunar regolith is approximately 45% oxygen by mass. The majority of the oxygen is bound in silicate minerals. The carbothermal reduction process has been proven to be effective at removing oxygen from lunar regolith simulants. The Carbothermal Reduction Demonstration (CaRD) project aims to increase the Technology Readiness Level (TRL) of a combined solar concentrator and carbothermal reduction systemin order to demonstrate this technology on the lunar surface. The CaRD project is divided into two design cycles, a brass board and prototype. The status of both design cycles will be discussed, as well as concepts for how this technology can be applied to the Artemis program in the future. Brassboard Vacuum Test: For the brassboard demonstration, a 2 kW Nd-YAG laser was used to heat lunar regolith simulant within a carbothermal reactor developed by Sierra Space. The reactor was placed inside of a 15 ft thermal vacuum chamber at the Johnson Space Center. The resulting reaction products were analyzed using both a gas chromotograph and mass spectrometer provided by Kennedy Space Center. Thermal data was also collected. Prototype Design: For the prototype, the CaRD team will perform another thermal vacuum test at JSC using the same interfaces and assets developed for the brassboard but will test a new carbothermal reactor design that Sierra Space is developing through the Carbothermal Oxygen Production Reactor (COPR) Tipping Point project that will include a means to autonomously move regolith in and out of the reactor. In addition, a deployable solar concentrator is being developed by Glenn Research Center using mirrors produced by Carbon Mirror Applications. The solar concentrator will be used to deliver solar energy into a carbothermal reactor to melt regolith and extract oxygen. Avionics and software for the concentrator are being developed by Kennedy Space Center. Future Concepts: An ongoing task within the CaRD project is to update models that can be used to determine the mass, power, and size of In-Situ Resource Utilization concepts at various scales. These models can now be used to analyze alternatives for future applications based on the latest available data.

Aaron Paz↗

NASA NTRS1984

Control system testing

A three stage process of ground testing of the Space Telescope Pointing Control System is used for verification prior to on-orbit operation. First, development tests are conducted in a laboratory environment using flight/engineering model control sensor and actuators configured with an engineering model of the flight computer and data management system breadboards. These development tests validate the results of computer simulations predicting control system performance. Integration tests bring together flight system elements and software interfaced to a software simulation of vehicle dynamics to confirm closed loop performance. The final ground test phase, flight systems testing, is conducted on the fully assembled Space Telescope, verifies interfaces with the Fine Guidance Sensors and includes a thermal vacuum testing period. During the final test phase, the Point Control System is exercised with the dynamics simulator running in real time.

Whittler, W. H.↗