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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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At least 163 records · Page 9
All sourcesNASA NTRSDOE OSTI
NASA NTRS2014

Tank Applied Testing of Load-Bearing Multilayer Insulation (LB-MLI)

The development of long duration orbital cryogenic storage systems will require the reduction of heat loads into the storage tank. In the case of liquid hydrogen, complete elimination of the heat load at 20 K is currently impractical due to the limitations in lift available on flight cryocoolers. In order to reduce the heat load, without having to remove heat at 20 K, the concept of Reduced Boil-Off uses cooled shields within the insulation system at approximately 90 K. The development of Load-Bearing Multilayer Insulation (LB-MLI) allowed the 90 K shield with tubing and cryocooler attachments to be suspended within the MLI and still be structurally stable. Coupon testing, both thermal and structural was performed to verify that the LB-MLI should work at the tank applied level. Then tank applied thermal and structural (acoustic) testing was performed to demonstrate the functionality of the LB-MLI as a structural insulation system. The LB-MLI showed no degradation of thermal performance due to the acoustic testing and showed excellent thermal performance when integrated with a 90 K class cryocooler on a liquid hydrogen tank.

Thermal Insulation↗
NASA NTRS2014

Recent Ground Hold and Rapid Depressurization Testing of Multilayer Systems

In the development of flight insulation systems for large cryogenic orbital storage (spray on foam and multilayer insulation), testing need include all environments that are experienced during flight. While large efforts have been expended on studying, bounding, and modeling the orbital performance of the insulation systems, little effort has been expended on the ground hold and ascent phases of a mission. Historical cryogenic in-space systems that have flown have been able to ignore these phases of flight due to the insulation system being within a vacuum jacket. In the development phase of the Nuclear Mars Vehicle and the Shuttle Nuclear Vehicle, several insulation systems were evaluated for the full mission cycle. Since that time there had been minimal work on these phases of flight until the Constellation program began investigating cryogenic service modules and long duration upper stages. With the inception of the Cryogenic Propellant Storage and Transfer Technology Demonstration Mission, a specific need was seen for the data and as such, several tests were added to the Cryogenic Boil-off Reduction System liquid hydrogen test matrix to provide more data on a insulation system. Testing was attempted with both gaseous nitrogen (GN2) and gaseous helium (GHe) backfills. The initial tests with nitrogen backfill were not successfully completed due to nitrogen liquefaction and solidification preventing the rapid pumpdown of the vacuum chamber. Subsequent helium backfill tests were successful and showed minimal degradation. The results are compared to the historical data.

Space Environment Simulation↗
NASA NTRS2014

Tank Applied Testing of Load-Bearing Multilayer Insulation (LB-MLI)

The development of long duration orbital cryogenic storage systems will require the reduction of heat loads into the storage tank. In the case of liquid hydrogen, complete elimination of the heat load at 20 K is currently impractical due to the limitations in lift available on flight cryocoolers. In order to reduce the heat load, without having to remove heat at 20 K, the concept of Reduced Boil-Off uses cooled shields within the insulation system at approximately 90 K. The development of Load-Bearing Multilayer Insulation (LB-MLI) allowed the 90 K shield with tubing and cryocooler attachments to be suspended within the MLI and still be structurally stable. Coupon testing both thermally and structurally were performed to verify that the LB-MLI should work at the tank applied level. Then tank applied thermal and structural (acoustic) testing was performed to demonstrate the functionality of the LB-MLI as a structural insulation system. The LB-MLI showed no degradation of thermal performance due to the acoustic testing and showed excellent thermal performance when integrated with a 90 K class cryocooler on a liquid hydrogen tank.

Space Environment Simulation↗
NASA NTRS2014

Insulation Test Cryostat with Lift Mechanism

A multi-purpose, cylindrical thermal insulation test apparatus is used for testing insulation materials and systems of materials using a liquid boil-off calorimeter system for absolute measurement of the effective thermal conductivity (k-value) and heat flux of a specimen material at a fixed environmental condition (cold-side temperature, warm-side temperature, vacuum pressure level, and residual gas composition). The apparatus includes an inner vessel for receiving a liquid with a normal boiling point below ambient temperature, such as liquid nitrogen, enclosed within a vacuum chamber. A cold mass assembly, including the upper and lower guard chambers and a middle test vessel, is suspended from a lid of the vacuum canister. Each of the three chambers is filled and vented through a single feedthrough. All fluid and instrumentation feedthroughs are mounted and suspended from a top domed lid to allow easy removal of the cold mass. A lift mechanism allows manipulation of the cold mass assembly and insulation test article.

Fesmire, James E.↗
NASA NTRS2010

Supercritical CO2 Cleaning for Planetary Protection and Contamination Control

We have designed and built a prototype Supercritical CO? Cleaning (SCC) system at JPL. The key features of the system are: 1) the parts inside a high-pressure vessel can be rotated at high speeds; 2) the same thermodynamic condition is maintained during First-In First-Out flushing to keep solvent power constant; and 3) the boil-off during decompression is induced in a separate vessel downstream. Our goal is to demonstrate SCC's ability to remove trace amounts of microbial and organic contaminants down to parts per billion levels from spacecraft material surfaces for future astrobiology missions. The initial cleaning test results showed that SCC can achieve cleanliness levels of 0.01 microgram/cm(sup 2) or less for hydrophobic contaminants such as dioctyl phthalate and silicone and it is less effective in the removal and inactivation of the hydrophilic bacterial spores as expected. However, with the use of a polar co-solvent, the efficacy may improve dramatically. The same results were obtained using liquid CO?. This opens up the possibility of using subcritical cleaning conditions, which may prove to be more compatible with certain spacecraft hardware.

Lin, Ying↗
NASA NTRS2015

Measuring the Heat Load on the Flight ASTRO-H Soft Xray Spectrometer Dewar

The Soft Xray Spectrometer (SXS) instrument on-board the ASTRO-H X-ray mission is based on microcalorimeters operating at 50 mK. Low temperature is achieved by use of an adiabatic demagnetization refrigerator (ADR) cyclically operating up to a heat sink at either 1.2 K or 4.5 K. The 1.2 K heat sink is provided by a 40 liter superfluid helium dewar. The parasitic heat to the helium from supports, plumbing, wires, and radiation, and the cyclic heat dumped by the ADR operation determine the liquid helium lifetime. To measure this lifetime we have used various techniques to rapidly achieve thermal equilibrium and then measure the boil-off rate of the helium. We have measured a parasitic heat of 650 microwatts and a cyclic heat of 100 microwatts for a total of 750 microwatts. This closely matches the predicted heat load. Starting with a fill level at launch of more than 33 liters results in a lifetime of greater than 4 years for the liquid helium. The techniques and accuracy for this measurement will be explained in this paper.

cryogenic Equipment↗
NASA NTRS2015

Validation of a Computational Model for the SLS Core Stage Oxygen Tank Diffuser Concept and the Low Profile Diffuser - An Advanced Development Design for the SLS

The Low Profile Diffuser (LPD) project originated as an award from the Marshall Space Flight Center (MSFC) Advanced Development (ADO) office to the Main Propulsion Systems Branch (ER22). The task was created to develop and test an LPD concept that could produce comparable performance to a larger, traditionally designed, ullage gas diffuser while occupying a smaller volume envelope. Historically, ullage gas diffusers have been large, bulky devices that occupy a significant portion of the propellant tank, decreasing the tank volume available for propellant. Ullage pressurization of spacecraft propellant tanks is required to prevent boil-off of cryogenic propellants and to provide a positive pressure for propellant extraction. To achieve this, ullage gas diffusers must slow hot, high-pressure gas entering a propellant tank from supersonic speeds to only a few meters per second. Decreasing the incoming gas velocity is typically accomplished through expansion to larger areas within the diffuser which has traditionally led to large diffuser lengths. The Fluid Dynamics Branch (ER42) developed and applied advanced Computational Fluid Dynamics (CFD) analysis methods in order to mature the LPD design from and initial concept to an optimized test prototype and to provide extremely accurate pre-test predictions of diffuser performance. Additionally, the diffuser concept for the Core Stage of the Space Launch System (SLS) was analyzed in a short amount of time to guide test data collection efforts of the qualification of the device. CFD analysis of the SLS diffuser design provided new insights into the functioning of the device and was qualitatively validated against hot wire anemometry of the exterior flow field. Rigorous data analysis of the measurements was performed on static and dynamic pressure data, data from two microphones, accelerometers and hot wire anemometry with automated traverse. Feasibility of the LPD concept and validation of the computational model were demonstrated by the test data.

Brodnick, Jacob↗
NASA NTRS2015

The Safe Removal of Frozen Air from the Annulus of an LH2 Storage Tank

Large Liquid Hydrogen (LH2) storage tanks are vital infrastructure for NASA. Eventually, air may leak into the evacuated and perlite filled annular region of these tanks. Although the vacuum level is monitored in this region, the extremely cold temperature causes all but the helium and neon constituents of air to freeze. A small, often unnoticeable pressure rise is the result. As the leak persists, the quantity of frozen air increases, as does the thermal conductivity of the insulation system. Consequently, a notable increase in commodity boil-off is often the first indicator of an air leak. Severe damage can result from normal draining of the tank. The warming air will sublimate which will cause a pressure rise in the annulus. When the pressure increases above the triple point, the frozen air will begin to melt and migrate downward. Collection of liquid air on the carbon steel outer shell may chill it below its ductility range, resulting in fracture. In order to avoid a structural failure, as described above, a method for the safe removal of frozen air is needed. A thermal model of the storage tank has been created using SINDA/FLUINT modeling software. Experimental work is progressing in an attempt to characterize the thermal conductivity of a perlite/frozen nitrogen mixture. A statistical mechanics model is being developed in parallel for comparison to experimental work. The thermal model will be updated using the experimental/statistical mechanical data, and used to simulate potential removal scenarios. This paper will address methodologies and analysis techniques for evaluation of two proposed air removal methods.

cryogenic insulation↗
NASA NTRS2015

The Structural Heat Intercept-Insulation-Vibration Evaluation Rig (SHIVER)

NASA is currently investigating methods to reduce the boil-off rate on large cryogenic upper stages. Two such methods to reduce the total heat load on existing upper stages are vapor cooling of the cryogenic tank support structure and integration of thick multilayer insulation systems to the upper stage of a launch vehicle. Previous efforts have flown a 2-layer MLI blanket and shown an improved thermal performance, and other efforts have ground-tested blankets up to 70 layers thick on tanks with diameters between 2 3 meters. However, thick multilayer insulation installation and testing in both thermal and structural modes has not been completed on a large scale tank. Similarly, multiple vapor cooled shields are common place on science payload helium dewars; however, minimal effort has gone into intercepting heat on large structural surfaces associated with rocket stages. A majority of the vapor cooling effort focuses on metallic cylinders called skirts, which are the most common structural components for launch vehicles. In order to provide test data for comparison with analytical models, a representative test tank is currently being designed to include skirt structural systems with integral vapor cooling. The tank is 4 m in diameter and 6.8 m tall to contain 5000 kg of liquid hydrogen. A multilayer insulation system will be designed to insulate the tank and structure while being installed in a representative manner that can be extended to tanks up to 10 meters in diameter. In order to prove that the insulation system and vapor cooling attachment methods are structurally sound, acoustic testing will also be performed on the system. The test tank with insulation and vapor cooled shield installed will be tested thermally in the B2 test facility at NASAs Plumbrook Station both before and after being vibration tested at Plumbrooks Space Power Facility.

Cryogenic Fluid Management↗
NASA NTRS2015

Wrapped Multilayer Insulation

New NASA vehicles, such as Earth Departure Stage (EDS), Orion, landers, and orbiting fuel depots, need improved cryogenic propellant transfer and storage for long-duration missions. Current cryogen feed line multilayer insulation (MLI) performance is 10 times worse per area than tank MLI insulation. During each launch, cryogenic piping loses approximately 150,000 gallons (equivalent to $300,000) in boil-off during transfer, chill down, and ground hold. Quest Product Development Corp., teaming with Ball Aerospace, developed an innovative advanced insulation system, Wrapped MLI (wMLI), to provide improved thermal insulation for cryogenic feed lines. wMLI is high-performance multilayer insulation designed for cryogenic piping. It uses Quest's innovative discrete-spacer technology to control layer spacing/ density and reduce heat leak. The Phase I project successfully designed, built, and tested a wMLI prototype with a measured heat leak 3.6X lower than spiral-wrapped conventional MLI widely used for piping insulation. A wMLI prototype had a heat leak of 7.3 W/sq m, or 27 percent of the heat leak of conventional MLI (26.7 W/sq m). The Phase II project is further developing wMLI technology with custom, molded polymer spacers and advancing the product toward commercialization via a rigorous testing program, including developing advanced vacuuminsulated pipe for ground support equipment.

Dye, Scott A.↗
NASA NTRS2015

Wrapped Multilayer Insulation

New NASA vehicles, such as Earth Departure Stage (EDS), Orion, landers, and orbiting fuel depots, need improved cryogenic propellant transfer and storage for long-duration missions. Current cryogen feed line multilayer insulation (MLI) performance is 10 times worse per area than tank MLI insulation. During each launch, cryogenic piping loses approximately 150,000 gallons (equivalent to $300,000) in boil-off during transfer, chill down, and ground hold. Quest Product Development Corp., teaming with Ball Aerospace, developed an innovative advanced insulation system, Wrapped MLI (wMLI), to provide improved thermal insulation for cryogenic feed lines. wMLI is high-performance multilayer insulation designed for cryogenic piping. It uses Quest's innovative discrete-spacer technology to control layer spacing/ density and reduce heat leak. The Phase I project successfully designed, built, and tested a wMLI prototype with a measured heat leak 3.6X lower than spiral-wrapped conventional MLI widely used for piping insulation. A wMLI prototype had a heat leak of 7.3 W/m2, or 27 percent of the heat leak of conventional MLI (26.7 W/m2). The Phase II project is further developing wMLI technology with custom, molded polymer spacers and advancing the product toward commercialization via a rigorous testing program, including developing advanced vacuuminsulated pipe for ground support equipment.

Dye, Scott A.↗
NASA NTRS2016

After Math - Foamology and Flight Rationale

The Space Shuttle was developed by NASA to be a largely reusable launch system which could provide frequent access to low earth orbit. Like all previous launch systems, safe reentry for the crew and payload required the use of a thermal protection system (TPS). Unlike previous spacecraft though, the Shuttle's TPS was exposed from launch, making it sensitive to debris which could be generated by the vehicle on ascent. The most likely and potentially destructive source of debris was considered to be ice, which could build-up anywhere on the External Tank (ET) where there was exposed metal. Ice could form during ground operations after the cryogenic propellants had been loaded and then be knocked loose on ascent. In order to prevent both ice build-up and boil-off of the propellants, the entire ET and all protuberances (orbiter attach points, pressurization lines, propellant feed lines, etc.) were covered with a spray on foam insulation (SOFI) type TPS. Unfortunately the foam was also susceptible to liberation during ascent, and posed a debris risk of its own. During the early years of the Shuttle Program engineers spent a good deal of effort characterizing the amount of foam that was shed.

Steva, Thomas↗
NASA NTRS2016

Insulation Test Cryostat with Lift Mechanism

A multi-purpose, cylindrical thermal insulation test apparatus is used for testing insulation materials and systems of materials using a liquid boil-off calorimeter system for absolute measurement of the effective thermal conductivity (k-value) and heat flux of a specimen material at a fixed environmental condition (cold-side temperature, warm-side temperature, vacuum pressure level, and residual gas composition). An inner vessel receives liquid with a normal boiling point below ambient temperature, such as liquid nitrogen, enclosed within a vacuum chamber. A cold mass assembly, including upper and lower guard chambers and middle test vessel, is suspended from a lid of the vacuum canister. Each of the three chambers is filled and vented through a single feedthrough. All fluid and instrumentation feedthroughs are mounted and suspended from a top domed lid allowing easy removal of the cold mass. A lift mechanism allows manipulation of the cold mass assembly and insulation test article.

Fesmire, James E.↗
NASA NTRS2017

An Overview of Recent Cryogenic Fluid Management Developments

Long-term storage, supply, and transfer of cryogenic fluids are critical capabilities needed to advance the human exploration of space. Technologies and advanced development programs have been pursued to address issues likely to confront the designers and developers of future cryogenic fluid management (CFM) subsystems and propulsion systems. NASA and aerospace industries have continued to conduct research and development for the propulsion applications of cryogenic fluids. From the well known in-space applications, to new applications involving superconducting motors within multiple different aircraft, and a renewed interest in production of oxygen on Mars, NASA continues to probe cryogenic applications within propulsion. This article presents brief reviews of several of the current CFM efforts to support future space missions. NASA GRC is currently focusing on understanding some of the finer points in the application of multilayer insulation. GRC recently finished the activation of a new calorimeter that operates at 20 K with a warm boundary that can operate either around 90 K or at 300 K. Current testing is focused on investigating multiple different seam concepts, initially between temperatures of 300 K and 20 K, and between 20 K and 90 K. One of the larger recent NASA investments has been on the Structural Heat Intercept, Insulation, and Vibration Experiment Rig (SHIIVER). SHIIVER is a 4 m diameter tank that is approximately 1/2 scale of the planned upper stage of the new Space Launch System (SLS) rocket. SHIIVER is focused on demonstrating the thermal benefits of multilayer insulation on the tank domes and boil-off vapor cooling on structural cylinders that hold the tank in-line on the rocket. It will also quantify any damage that may incur during the acoustic environment of over 160 decibels that SLS will experience on its trip to Earth orbit. In support of the possible production of oxygen out of the Mars' atmosphere, a team comprised of four NASA centers (GRC, MSFC, JSC, and KSC) collaborated to investigate multiple different hardware combinations, refrigeration cycles, and integration techniques to minimize power and mass of the storage and liquefaction system. Moreover, high efficiency, high capacity cryocoolers are an element of CFM which is essential for achieving NASA's future long duration missions. Currently in development is a 20W at 20K unit for use with liquid hydrogen. It is scheduled for delivery to NASA in the summer of 2017. Two Phase I SBIRS have recently concluded resulting in the preliminary design of two different 90K units, each having a refrigeration capacity of approximately 150W. Once the development efforts are complete, these high capacity 90K units will be integrated into the design of NASA's Liquid Oxygen (LOX)/Methane applications such as the Lander and Ascent Vehicles, or possibly an in-space stage. CFM technology is critical to the success of missions to Mars, Planetary Exploration, and In-Situ Resource Utilization (ISRU) for cryogenic propellant production. NASA is focusing on the development of CFM technologies needed to provide necessary data and relevant experience to support informed decisions on implementation of design of cryogenic systems for long term space missions. Lessons learned from the described CFM developments would lead to enhanced safety and reliability and enabling technologies which could allow NASA to meet future space exploration goals.

Hedayat, A.↗
NASA NTRS2017

Cryogenic Fluid Management Technology Development Roadmaps

Advancement in Cryogenic Fluid Management (CFM) Technologies is essential for achieving NASA's future long duration missions. Propulsion systems utilizing cryogens are necessary to achieve mission success. Current State Of the Art (SOA) CFM technologies enable cryogenic propellants to be stored for several hours. However, some envisioned mission architectures require cryogens to be stored for two years or longer. The fundamental roles of CFM technologies are long term storage of cryogens, propellant tank pressure control and propellant delivery. In the presence of heat, the cryogens will "boil-off" over time resulting in excessive pressure buildup, off-nominal propellant conditions, and propellant loss. To achieve long term storage and tank pressure control, the CFM elements will intercept and/or remove any heat from the propulsion system. All functions are required to perform both with and without the presence of a gravitational field. Which CFM technologies are required is a function of the cryogens used, mission architecture, vehicle design and propellant tank size. To enable NASA's crewed mission to the Martian surface, a total of seventeen CFM technologies have been identified to support an In-Space Stage and a Lander/Ascent Vehicle. Recognizing that FY2020 includes a Decision Point regarding the In-Space Stage Architecture, a set of CFM Technology Development Roadmaps have been created identifying the current Technology Readiness Level (TRL) of each element, current technology "gaps", and existing technology development efforts. The roadmaps include a methodical approach and schedule to achieve a flight demonstration in FY2023, hence maturing CFM technologies to TRL 7 for infusion into the In-Space Stage Preliminary Design.

Stephens, J. R.↗
NASA NTRS2017

Investigation of Vapor Cooling Enhancements for Applications on Large Cryogenic Systems

The need to demonstrate and evaluate the effectiveness of heat interception methods for use on a relevant cryogenic propulsion stage at a system level has been identified. Evolvable Cryogenics (eCryo) Structural Heat Intercept, Insulation and Vibration Evaluation Rig (SHIIVER) will be designed with vehicle specific geometries (SLS Exploration Upper Stage (EUS) as guidance) and will be subjected to simulated space environments. One method of reducing structure-born heat leak being investigated utilizes vapor-based heat interception. Vapor-based heat interception could potentially reduce heat leak into liquid hydrogen propulsion tanks, increasing potential mission length or payload capability. Due to the high number of unknowns associated with the heat transfer mechanism and integration of vapor-based heat interception on a realistic large-scale skirt design, a sub-scale investigation was developed. The sub-project effort is known as the Small-scale Laboratory Investigation of Cooling Enhancements (SLICE). The SLICE aims to study, design, and test sub-scale multiple attachments and flow configuration concepts for vapor-based heat interception of structural skirts. SLICE will focus on understanding the efficiency of the heat transfer mechanism to the boil-off hydrogen vapor by varying the fluid network designs and configurations. Various analyses were completed in MATLAB, Excel VBA, and COMSOL Multiphysics to understand the optimum flow pattern for heat transfer and fluid dynamics. Results from these analyses were used to design and fabricate test article subsections of a large forward skirt with vapor cooling applied. The SLICE testing is currently being performed to collect thermal mechanical performance data on multiple skirt heat removal designs while varying inlet vapor conditions necessary to intercept a specified amount of heat for a given system. Initial results suggest that applying vapor-cooling provides a 50 heat reduction in conductive heat transmission along the skirt to the tank. The information obtained by SLICE will be used by the SHIIVER engineering team to design and implement vapor-based heat removal technology into the SHIIVER forward skirt hardware design.

small scale testing↗
NASA NTRS2017

Heat Loads Due to Small Penetrations in Multilayer Insulation Blankets

The main penetrations (supports and piping) through multilayer insulation systems for cryogenic tanks have been previously addressed by heat flow measurements. Smaller penetrations due to fasteners and attachments are now experimentally investigated. The use of small pins or plastic garment tag fasteners to each the handling and construction of multilayer insulation (MLI) blankets goes back many years. While it has long been understood that penetrations and other discontinuities degrade the performance of the MLI blanket, quantification of this degradation has generally been lumped into gross performance multipliers (often called degradation factors or scale factors). Small penetrations contribute both solid conduction and radiation heat transfer paths through the blanket. The conduction is down the stem of the structural element itself while the radiation is through the hole formed during installation of the pin or fastener. Analytical models were developed in conjunction with MLI perforation theory and Fouriers Law. Results of the analytical models are compared to experimental testing performed on a 10 layer MLI blanket with approximately 50 small plastic pins penetrating the test specimen. The pins were installed at 76-mm spacing inches in both directions to minimize the compounding of thermal effects due to localized compression or lateral heat transfer. The testing was performed using a liquid nitrogen boil-off calorimeter (Cryostat-100) with the standard boundary temperatures of 293 K and 78 K. Results show that the added radiation through the holes is much more significant than the conduction down the fastener. The results are shown to be in agreement with radiation theory for perforated films.

Structural attachments↗
NASA NTRS2017

Long-Term Cryogenic Propellant Storage on Mars with Hercules Propellant Storage Facility

This report details the process and results of roughly sizing the steady state, zero boil-off thermal and power parameters of the Hercules Propellant Storage Facility. For power analysis, isothermal and isobaric common bulkhead tank scenarios are considered. An estimated minimum power requirement of 8.3 kW for the Reverse Turbo-Brayton Cryocooler is calculated. Heat rejection concerns in soft vacuum Mars atmosphere are noted and potential solutions are proposed. Choice of coolant for liquid propellant conditioning and issues with current proposed cryocooler cycle are addressed; recommendations are made, e.g. adding a Joule-Thomson expansion valve after the Reverse Turbo-Brayton turbine in order to have two-phase, isothermal heat exchange through the Broad Area Cooling system. Issues with cross-country transfer lines from propellant storage to flight vehicle are briefly discussed: traditional vacuum jacketed lines are implausible, and Mars insulation needs to be developed.

Propellant Storage↗