Results for “cryogenic engineering” — Page 40

NASA NTRS2007

An Infrared Solution to a National Priority NASA Ice Detection and Measurement Problem

NASA has a serious problem with ice that forms on the cryogenic-filled Space Shuttle External Tank (ET) that could endanger the crew and vehicle. This problem has defied resolution in the past. To find a solution, a cooperative agreement was developed between NASA-Kennedy Space Center (KSC) and the U.S. Army-Tank-Automotive, armaments Research, Development & Engineering Center (TARDEC). This paper describes the need, initial investigation, solution methodology, and some results for a mobile near-IR ice detection and measurement system developed by MDA of Canada and jointly tested by the U.S. Army TARDEC and NASA. Performance results achieved demonstrate that the pre-launch inspection system has the potential to become a critical tool in addressing NASA's ice problem.

Meitzler, Thomas↗

NASA NTRS1989

Propulsion requirements for the supply of lunar-derived oxygen for a manned Mars mission

The use of lunar-derived liquid oxygen with hydrogen for robust manned explorations beyond earth is discussed. The propellant mixture ratio for optimal performance is near the current design point of r = 6. The results of NASA-sponsored studies of missions using cryogenic bipropellant H2-O2 for most functions reveal the potential benefits from a propellant terminal in cis-lunar space. The main propulsion technologies associated with Mars mission case studies include: in-space fluid transfer, zero-g ignition, oxygen-enriched propellant mixtures, throttling of main engines, and H2-O2 based auxiliary propulsion.

Wrobel, J. Richard↗

NASA NTRS2015

NASA James Webb Space Telescope Engineering of the Primary Mirror Segment Assemblies (PMSA) and the Primary Mirror Backplane Support Structure (PMBSS)

The design, engineering tests of the PMSAs PMBSS show that we have a robust system that not only meets but exceeds (better than) the design requirements for these components. In the next 2 years the Telescope Observatory will be subjected to a simulated launch environment (sine vibeacoustics) and operations tests at cryogenic temperatures. Launch is schedule for the last quarter of 2018.

PMBSS↗

NASA NTRS2015

Composite Cryotank Technologies and Development 2.4 and 5.5M out of Autoclave Tank Test Results

The Composite Cryotank Technologies and Demonstration (CCTD) project substantially matured composite, cryogenic propellant tank technology. The project involved the design, analysis, fabrication, and testing of large-scale (2.4-m-diameter precursor and 5.5-m-diameter) composite cryotanks. Design features included a one-piece wall design that minimized tank weight, a Y-joint that incorporated an engineered material to alleviate stress concentration under combined loading, and a fluted core cylindrical section that inherently allows for venting and purging. The tanks used out-of-autoclave (OoA) cured graphite/epoxy material and processes to enable large (up to 10-m-diameter) cryotank fabrication, and thin-ply prepreg to minimize hydrogen permeation through tank walls. Both tanks were fabricated at Boeing using automated fiber placement on breakdown tooling. A fluted core skirt that efficiently carried axial loads and enabled hydrogen purging was included on the 5.5-m-diameter tank. Ultrasonic inspection was performed, and a structural health monitoring system was installed to identify any impact damage during ground processing. The precursor and 5.5-m-diameter tanks were tested in custom test fixtures at the National Aeronautics and Space Administration Marshall Space Flight Center. The testing, which consisted of a sequence of pressure and thermal cycles using liquid hydrogen, was successfully concluded and obtained valuable structural, thermal, and permeation performance data. This technology can be applied to a variety of aircraft and spacecraft applications that would benefit from 30 to 40% weight savings and substantial cost savings compared to aluminum lithium tanks.

Jackson, Justin R.↗

NASA NTRSDate not supplied

Integral Channel Nozzles and Heat Exchangers using Additive Manufacturing Directed Energy Deposition NASA HR-1 Alloy

Heat exchangers for use in propulsion applications are very critical components because they must be efficient, compact and light and often operate with working fluids at extreme temperatures or pressures or both. Various components and systems use heat exchangers such as combustion chambers of gas turbines and internal combustion engines, fuel cells (air supply and thermal management), electric batteries (thermal management), evaporators and recuperators of waste-heat-to-power systems, and rocket engines. Even if the results are more generally applicable, the heat exchangers applications to which this study is more closely related are regeneratively cooled rocket nozzles and chambers, and repressurization systems for the launch vehicles. These components are often thin-walled and contain pressurized fluids, like propellants at cryogenic or elevated temperatures. Given that the environments that these propulsion components must endure are challenging, the manufacturing to meet these specifications often require long lead times due to specialty processes and unique tooling associated with the combined thin-wall integral channel and large-scale structures. Additive manufacturing (AM) offers programmatic advantages for reduction in processing time and cost in addition to various technical advantages, including the possibility to achieve enhanced hardware complexity targeted to superior performance, part consolidation, and the capability of processing of novel alloys. While AM is already being utilized for heat exchanger components in propulsion applications, almost all these AM components are made by means of Laser Powder Bed Fusion (L-PBF). L-PBF allows for fine features but is rather limited with respect to the overall size of the components that can be manufactured. Recent developments are maturing the Laser Powder Directed Energy Deposition (LP-DED) process which may be used, for example, to make integral channel thin-wall regeneratively-cooled rocket nozzles with diameters greater than 1 m. This paper highlights some integral channel heat exchanger demonstrator hardware applications of LP-DED, as well as the characterization of this process in combination with the use of the NASA HR-1 alloy. To properly utilize LP-DED for heat exchanger manufacturing, various aspects are being characterized such as geometry limitations, measurement of surface texture and geometric angled surfaces, surface enhancements for internal channels, and material evaluation. NASA HR-1 (FeNi-Cr) is a high strength hydrogen resistant superalloy developed for use in aerospace applications, such as heat exchangers. Some aspects and considerations about the design of heat exchangers are summarized together with data relevant to LP-DED manufacturing in combination with the NASA HR-1 alloy. Microchannels were successful deposited down to 2.54 mm and 1 mm wall thickness, wall angles of 30°, both with high reproducibility. It was also found that the areal surface roughness is highly dependent on the size of the powder feedstock used for deposition. The characterization of these LP-DED features is critical for fluid flow and heat transfer predictions as it can be exploited to enhance heat transfer at the cost of increased pressure drop.

Additive Manufacturing↗

NASA NTRSDate not supplied

Integral Channel Nozzles and Heat Exchangers using Additive Manufacturing Directed Energy Deposition NASA HR-1 Alloy

Heat exchangers for use in propulsion applications are very critical components because they must be efficient, compact and light and often operate with working fluids at extreme temperatures or pressures or both. Various components and systems use heat exchangers such as combustion chambers of gas turbines and internal combustion engines, fuel cells (air supply and thermal management), electric batteries (thermal management), evaporators and recuperators of waste-heat-to-power systems, and rocket engines. Even if the results are more generally applicable, the heat exchangers applications to which this study is more closely related are regeneratively cooled rocket nozzles and chambers, and repressurization systems for the launch vehicles. These components are often thin-walled and contain pressurized fluids, like propellants at cryogenic or elevated temperatures. Given that the environments that these propulsion components must endure are challenging, the manufacturing to meet these specifications often require long lead times due to specialty processes and unique tooling associated with the combined thin-wall integral channel and large-scale structures. Additive manufacturing (AM) offers programmatic advantages for reduction in processing time and cost in addition to various technical advantages, including the possibility to achieve enhanced hardware complexity targeted to superior performance, part consolidation, and the capability of processing of novel alloys. While AM is already being utilized for heat exchanger components in propulsion applications, almost all these AM components are made by means of Laser Powder Bed Fusion (L-PBF). L-PBF allows for fine features but is rather limited with respect to the overall size of the components that can be manufactured. Recent developments are maturing the Laser Powder Directed Energy Deposition (LP-DED) process which may be used, for example, to make integral channel thin-wall regeneratively-cooled rocket nozzles with diameters greater than 1 m. This paper highlights some integral channel heat exchanger demonstrator hardware applications of LP-DED, as well as the characterization of this process in combination with the use of the NASA HR-1 alloy. To properly utilize LP-DED for heat exchanger manufacturing, various aspects are being characterized such as geometry limitations, measurement of surface texture and geometric angled surfaces, surface enhancements for internal channels, and material evaluation. NASA HR-1 (FeNi-Cr) is a high strength hydrogen resistant superalloy developed for use in aerospace applications, such as heat exchangers. Some aspects and considerations about the design of heat exchangers are summarized together with data relevant to LP-DED manufacturing in combination with the NASA HR-1 alloy. Microchannels were successful deposited down to 2.54 mm and 1 mm wall thickness, wall angles of 30°, both with high reproducibility. It was also found that the areal surface roughness is highly dependent on the size of the powder feedstock used for deposition. The characterization of these LP-DED features is critical for fluid flow and heat transfer predictions as it can be exploited to enhance heat transfer at the cost of increased pressure drop.

additive manufacturing↗

NASA NTRSDate not supplied

CFD Modeling & Validation of Single Bubble Growth During Cavitation in Zero Gravity

Cryogenic Fluid Management (CFM) in zero gravity represents a vital area of research as NASA pursues future long-duration space missions. Success of these missions relies heavily on the ability to store and transfer the cryogenic liquid fuels efficiently and safely. In the Zero Boil-Off Tank (ZBOT-1) Experiment in 2018, unexpected and intense phase change occurred at the screened Liquid Acquisition Device (LAD) during tank pressure control that was later attributed to cavitation. The cavitation ultimately resulted in pump failure. Since cryogen cavitation during storage and transfer of propellants in microgravity has been a source of serious concern for the aerospace community, a series of drop tower experiments with liquid methane were also performed at the University of Bremen to carefully study cavitation from a single bubble nucleus in an engineered micron-sized wall cavity. As the test cell is depressurized during the 10-second drop, cavitation causes bubble growth from the nucleus. Here, in a collaboration between the Case Western Reserve University and University of Bremen, we present a two-phase Computational Fluid Dynamics (CFD) model of bubble growth from a single wall cavity during depressurization in microgravity. Both 2D axisymmetric and 3D parametric simulation case studies are performed to show the evolution of the bubble growth. The model uses a Volume of Fluid (VOF) approach with and without adaptive mesh refinement and adaptive time stepping to capture the growth dynamics using an explicit formulation with algebraic reconstruction of the interface. CFD results show good agreement with corresponding experiment. Based on the velocity, temperature, and volume fraction fields generated by the validation case studies, detailed discussion and analysis of fluid flow and heat transfer during bubble growth process are presented to delineate the intricacies of the cavitation process. Strengths and shortcomings of the present numerical approach in capturing this phenomenon are also highlighted.

Two Phase Flow↗

NASA NTRS2024

Characterization of NUW-LHT-5m, A Lunar Highland Simulant

A new simulant of the lunar highlands regolith, NUW-LHT-5M, was designed by NASA and manufactured by Washington Mills. The simulant was based on Apollo 16 data and is a member of the NU-LHT-series. NASA’s Marshall Space Flight Center and Johnson Space Center have already purchased 3 metric tons of the simulant for advanced engineering work. In support of engineering uses of the simulant, we provided measurements of the simulant including: mineral abundance and composition, liberation, X-ray fluorescence (XRF), ferrous iron, carbon, sulfur, 60 element inductively coupled plasma (ICP), loss on ignition, particle size, both 2D and 3D particle shape, specific surface area, shear, cohesion, internal friction, helium pycnometry, minimum index density, tap density, magnetic susceptibility, cryogenic and high temperature permittivity, visible and near-infrared (VNIR) and middle infra-red spectroscopy (MIR), differential scanning calorimetry (DSC), viscosity, thermal diffusivity, thermal conductivity, thermal gravimetric analysis (TGA), evolved gas analysis (EGA), and spark sintering. For the crystalline components the design of the simulant called for two rocks from the Stillwater Complex, Montana: 17.6 wt% norite, 37.7% anorthosite, and 4.7 wt% olivine from an unspecified commercial source. The other 40% of the simulant was a high calcium (An100), vesicular glass that Washington Mills made specifically for the simulant. Fabrication and quality control processes for both the glass and the simulant are described. Importantly, most of the graphs and tables presented herein provide values for both the new simulant and data for the older NASA mare simulant, JSC-1A. Finally, we discussed the current limitations of NUW-LT-5M and most other lunar regolith simulants to replicate the lunar material.

lunar regolith simulant↗

NASA NTRS1985

Mathematical models for space shuttle ground systems

Math models are a series of algorithms, comprised of algebraic equations and Boolean Logic. At Kennedy Space Center, math models for the Space Shuttle Systems are performed utilizing the Honeywell 66/80 digital computers, Modcomp II/45 Minicomputers and special purpose hardware simulators (MicroComputers). The Shuttle Ground Operations Simulator operating system provides the language formats, subroutines, queueing schemes, execution modes and support software to write, maintain and execute the models. The ground systems presented consist primarily of the Liquid Oxygen and Liquid Hydrogen Cryogenic Propellant Systems, as well as liquid oxygen External Tank Gaseous Oxygen Vent Hood/Arm and the Vehicle Assembly Building (VAB) High Bay Cells. The purpose of math modeling is to simulate the ground hardware systems and to provide an environment for testing in a benign mode. This capability allows the engineers to check out application software for loading and launching the vehicle, and to verify the Checkout, Control, & Monitor Subsystem within the Launch Processing System. It is also used to train operators and to predict system response and status in various configurations (normal operations, emergency and contingent operations), including untried configurations or those too dangerous to try under real conditions, i.e., failure modes.

Tory, E. G.↗

NASA NTRS2004

Numerical Modeling of Conjugate Heat Transfer in Fluid Network

Fluid network modeling with conjugate heat transfer has many applications in Aerospace engineering. In modeling unsteady flow with heat transfer, it is important to know the variation of wall temperature in time and space to calculate heat transfer between solid to fluid. Since wall temperature is a function of flow, a coupled analysis of temperature of solid and fluid is necessary. In cryogenic applications, modeling of conjugate heat transfer is of great importance to correctly predict boil-off rate in propellant tanks and chill down of transfer lines. In TFAWS 2003, the present author delivered a paper to describe a general-purpose computer program, GFSSP (Generalized Fluid System Simulation Program). GFSSP calculates flow distribution in complex flow circuit for compressible/incompressible, with or without heat transfer or phase change in all real fluids or mixtures. The flow circuit constitutes of fluid nodes and branches. The mass, energy and specie conservation equations are solved at the nodes where as momentum conservation equations are solved at the branches. The proposed paper describes the extension of GFSSP to model conjugate heat transfer. The network also includes solid nodes and conductors in addition to fluid nodes and branches. The energy conservation equations for solid nodes solves to determine the temperatures of the solid nodes simultaneously with all conservation equations governing fluid flow. The numerical scheme accounts for conduction, convection and radiation heat transfer. The paper will also describe the applications of the code to predict chill down of cryogenic transfer line and boil-off rate of cryogenic propellant storage tank.

Majumdar, Alok↗

NASA NTRS2017

Thermal Vacuum Testing of ICPTA RCS at Plum Brook B-2

Vacuum and thermal vacuum testing of the Integrated Cryogenic Propulsion Test Article (ICPTA) was performed at the Plum Brook B-2 facility as a part of a system checkout and facility characterization effort. Multiple test objectives included: integrated Reaction Control System (RCS) characterization, cold helium pressurization system characterization, modal propellant gaging experiment (Orion), CFM propellant loading experiments, main engine characterization. The ICPTA is a test bed for LOX/LCH4 technologies built in 2016 using new components and hardware from the former Morpheus vehicle and other projects.

Atwell, M. J.↗

NASA NTRS2017

Survey of Constellation-Era LOX/Methane Development Activities and Future Development Needs

NASA formed the Constellation Program in 2005 to achieve the objectives of maintaining American presence in low-Earth orbit, returning to the moon for purposes of establishing an outpost, and laying the foundation to explore Mars and beyond in the first half of the 21st century. The Exploration Technology Development Program (ETDP) was formulated to address the technology needs to address Constellation architecture decisions. The Propellants and Cryogenic Advanced Development (PCAD) project was tasked with risk mitigation of specific propulsion related technologies to support ETDP. Propulsion systems were identified as critical technologies owing to the high gear-ratio of lunar Mars landers Cryogenic propellants offer performance advantage over storables (NTOMMH) Mass savings translate to greater payload capacity In-situ production of propellant an attractive feature; methane and oxygen identified as possible Martian in-situ propellants New technologies were required to meet more difficult missions High performance LOX/LH2 deep throttle descent engines High performance LOX/LCH4 ascent main and reaction control system (RCS) engines The PCAD project sought to provide those technologies through Reliable ignition pulse RCS Fast start High efficiency engines Stable deep throttling.

Propulsion system components↗

NASA NTRS1965

Design of cryogenic fluid systems in the S-IV STAGE propulsion system.

Reliability design concepts in cryogenic fluid systems in Saturn S-IV stage propulsion system, noting hydrogen and LOX tank pressurization systems

SATURN S-4 STAGE↗

NASA NTRS2016

Jwst from Below: An Overview of the Construction of the James Webb Space Telescope, Interesting Metrology, and Cryogenic-Vacuum Testing

NASA's James Webb Space Telescope (JWST) is a 6.6m diameter, segmented, deployable telescope for cryogenic IR space astronomy (40K). The JWST Observatory includes the Optical Telescope Element (OTE) and the Integrated Science Instrument Module (ISIM) that contains four science instruments (SI) and the guider. The SIs are mounted to a composite metering structure. The SI and guider units are integrated to the ISIM structure and optically tested at NASA Goddard Space Flight Center as a suite using a telescope simulator (Optical Telescope Element SIMulator; OSIM). OSIM is a full field, cryogenic JWST telescope simulator. SI performance, including alignment and wavefront error, is evaluated using OSIM. This is an overview presentation to undergraduate students and other personnel at the University of Richmond, planned for 12 Oct, 2016. It uses material previously released by NASA on the Internet (e.g., via Flickr) or at engineering conferences (e.g., SPIE). This presentation provides an overview of the status of the project, with an emphasis on optics and measurement.

instrument development↗

NASA NTRS1981

Development and integration of the infrared telescope for Spacelab

A description is given of the design features of the Small Helium-Cooled Infrared Telescope (IRT) being developed for the 1983 Spacelab 2 mission and the problems encountered during the development of its (1) superfluid helium system requirements, (2) interfaces with Spacelab, and (3) Kennedy Space Center prelaunch operations requirements and restrictions. The IRT comprises cryogenic, IR optical, mechanical, and electronics subsystems. Attention is given the preflight management of superfluid helium by a vacuum maintenance assembly (VMA), which will operate almost continuously until launch and then be deactivated to fly as unavoidable ballast. The objective of IRT design and development is the meeting of IR science and engineering objectives with a relatively low-cost system.

Urban, E. W.↗

NASA NTRS1992

Modelling and experimental verification of a water alleviation system for the NASP

One possible low speed propulsion system for the National Aerospace Plane is a liquid air cycle engine (LACE). The LACE system uses the heat sink in the liquid hydrogen propellant to liquefy air in a heat exchanger which is then pumped up to high pressure and used as the oxidizer in a hydrogen liquid air rocket. The inlet airstream must be dehumidified or moisture could freeze on the cryogenic heat exchangers and block them. The main objective of this research has been to develop a computer simulation of the cold tube/antifreeze-spray water alleviation system and to verify the model with experimental data. An experimental facility has been built and humid air tests were conducted on a generic heat exchanger to obtain condensing data for code development. The paper describes the experimental setup, outlines the method of calculation used in the code, and presents comparisons of the calculations and measurements. Cause of discrepancies between the model and data are explained.

Vanfossen, G. James↗

NASA NTRS1995

Direct-Circulation Stirling-Cycle Refrigerator

In proposed cryogenic system, cold working fluid (helium) circulated directly from Stirling-cycle refrigerator machinery through loop of tubing to provide 15 W of cooling power at temperature of 80 K to scientific instruments or other devices located at some distance from machinery. (Separation of cooled devices from machinery may be dictated by engineering requirements including, need to isolate cooled devices from vibrations of machinery.) Advantages include less complexity, smaller size, lighter weight, and lower power consumption.

Ellison, Woody↗

NASA NTRS2014

Control of NASA's Space Launch System

The flight control system for the NASA Space Launch System (SLS) employs a control architecture that evolved from Saturn, Shuttle & Ares I-X while also incorporating modern enhancements. This control system, baselined for the first unmanned launch, has been verified and successfully flight-tested on the Ares I-X rocket and an F/A-18 aircraft. The development of the launch vehicle itself came on the heels of the Space Shuttle retirement in 2011, and will deliver more payload to orbit and produce more thrust than any other vehicle, past or present, opening the way to new frontiers of space exploration as it carries the Orion crew vehicle, equipment, and experiments into new territories. The initial 70 metric ton vehicle consists of four RS-25 core stage engines from the Space Shuttle inventory, two 5- segment solid rocket boosters which are advanced versions of the Space Shuttle boosters, and a core stage that resembles the External Tank and carries the liquid propellant while also serving as the vehicle's structural backbone. Just above SLS' core stage is the Interim Cryogenic Propulsion Stage (ICPS), based upon the payload motor used by the Delta IV Evolved Expendable Launch Vehicle (EELV).

VanZwieten, Tannen S.↗