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At least 433 records · Page 24
All sourcesNASA NTRSDOE OSTI
NASA NTRS2015

NASA Propulsion Sub-System Concept Studies and Risk Reduction Activities for Resource Prospector Lander

NASA's exploration roadmap is focused on developing technologies and performing precursor missions to advance the state of the art for eventual human missions to Mars. One of the key components of this roadmap is various robotic missions to Near-Earth Objects, the Moon, and Mars to fill in some of the strategic knowledge gaps. The Resource Prospector (RP) project is one of these robotic precursor activities in the roadmap. RP is a multi-center and multi-institution project to investigate the polar regions of the Moon in search of volatiles. The mission is rated Class D and is approximately 10 days, assuming a five day direct Earth to Moon transfer. Because of the mission cost constraint, a trade study of the propulsion concepts was conducted with a focus on available low-cost hardware for reducing cost in development, while technical risk, system mass, and technology advancement requirements were also taken into consideration. The propulsion system for the lander is composed of a braking stage providing a high thrust to match the lander's velocity with the lunar surface and a lander stage performing the final lunar descent. For the braking stage, liquid oxygen (LOX) and liquid methane (LCH4) propulsion systems, derived from the Morpheus experimental lander, and storable bi-propellant systems, including the 4th stage Peacekeeper (PK) propulsion components and Space Shuttle orbital maneuvering engine (OME), and a solid motor were considered for the study. For the lander stage, the trade study included miniaturized Divert Attitude Control System (DACS) thrusters (Missile Defense Agency (MDA) heritage), their enhanced thruster versions, ISE-100 and ISE-5, and commercial-off-the-shelf (COTS) hardware. The lowest cost configuration of using the solid motor and the PK components while meeting the requirements was selected. The reference concept of the lander is shown in Figure 1. In the current reference configuration, the solid stage is the primary provider of delta-V. It will generate 15,000-lbf of thrust with a single burn of ~ 80's seconds. The lander stage is a bi-propellant, pressure-regulated, pulsing liquid propulsion system to perform all other functions.

Trinh, Huu P.↗
NASA NTRS2015

SLS Test Stand Site Selection

Test site selection is a critical element of the design, development and production of a new system. With the advent of the new Space Launch System (SLS), the National Aeronautics and Space Administration (NASA) had a number of test site selection decisions that needed to be made early enough in the Program to support the planned Launch Readiness Date (LRD). This case study focuses on decisions that needed to be made in 2011 and 2012 in preparation for the April 2013 DPMC decision about where to execute the Main Propulsion Test that is commonly referred to as "Green Run." Those decisions relied upon cooperative analysis between the Program, the Test Lab and Center Operations. The SLS is a human spaceflight vehicle designed to carry a crew farther into space than humans have previously flown. The vehicle consists of four parts: the crew capsule, the upper stage, the core stage, and the first stage solid rocket boosters. The crew capsule carries the astronauts, while the upper stage, the core stage, and solid rocket boosters provide thrust for the vehicle. In other words, the stages provide the "lift" part of the lift vehicle. In conjunction with the solid rocket boosters, the core stage provides the initial "get-off-the-ground" thrust to the vehicle. The ignition of the four core stage engines and two solid rocket boosters is the first step in the launch portion of the mission. The solid rocket boosters burn out after about 2 minutes of flight, and are then jettisoned. The core stage provides thrust until the vehicle reaches a specific altitude and speed, at which point the core stage is shut off and jettisoned, and the upper stage provides vehicle thrust for subsequent mission trajectories. The integrated core stage primarily consists of a liquid oxygen tank, a liquid hydrogen tank, and the four core stage engines. For the SLS program, four RS-25 engines were selected as the four core stage engines. The RS-25 engine is the same engine that was used for Space Shuttle. The test plan for the integrated core stage was broken down into several segments: Component testing, system level testing, and element level testing. In this context, components are items such as valves, controllers, sensors, etc. Systems are items such as an entire engine, a tank, or the outer stage body. The core stage itself is considered to be an element. The rocket engines are also considered an element. At the program level, it was decided to perform a single green run test on the integrated core stage prior to shipment of it to Kennedy Space Center (KSC) for use in the EM-1 test flight of the SLS vehicle. A green run test is the first live fire of the new integrated core stage and engine elements - without boosters of course. The SLS Program had to decide where to perform SLS green run testing.

Crowe, Kathryn↗
NASA NTRS2019

Assembling and Testing NASA’s Space Launch System for First Flight

NASA is planning its next step toward human expansion into the solar system. The Space Launch System (SLS) (Figure 1) is a critical enabling component of that expansion. SLS payloads for its early missions include the Orion crew vehicle and components for Gateway, a lunar outpost orbiting the Moon that will facilitate research, technology and partnerships for eventual Mars missions. All major core stage hardware for test and flight completed structural manufacturing in 2018. The major components for the first flight vehicle are complete or approaching completion of internal equipment installation. The core stage forward join operation is also complete. The 10 booster segments needed for first flight have been cast and are ready to ship to NASA’s Kennedy Space Center (KSC) for mating and stacking. The four RS-25 core stage engines completed processing at NASA’s Stennis Space Center (SSC) and are ready for core stage integration. The Orion Stage Adapter (OSA) joined the Interim Cryogenic Propulsion Stage (ICPS) at KSC to await the rest of SLS flight hardware. Production and preparation of hardware for the second mission is also underway. Looking ahead to a busy 2019, liquid oxygen (LOX) tank, liquid hydrogen (LH2) tank and intertank structural testing will take place. This paper will discuss the current and planned status of SLS development in context of NASA’s overall exploration plans.

Askin, Bruce R.↗
NASA NTRS2022

Artemis I Countdown 101

Before the Artemis I mission launches on its way around the Moon, the launch team at Kennedy Space Center and supporting teams across the country will begin the launch countdown about two days before liftoff. The launch countdown contains "L Minus” and "T Minus" times. "L minus" indicates how far away we are from liftoff in hours and minutes and does not include built-in holds. “T minus” time is a sequence of events that are built into the launch countdown where counting and holds are inserted. The 730,000 gallons of propellants that will be loaded into the rocket are liquid oxygen (LO2) and liquid hydrogen (LH2). Pauses in the countdown, or "holds," are built into the countdown to allow the launch team to target a precise launch window, and to provide a cushion of time for certain tasks and procedures without impacting the overall schedule. For the Artemis I countdown, planned built-in holds vary in length and occur at the following times: L-8 hours, 40 minutes, and L-40 minutes.

Madison Elizabeth Tuttle↗
NASA NTRSDate not supplied

Nodal Modeling of Submerged Helium Injection Pressurization of a Cryogenic Propellant Tank

Subcooling of cryogenic propellant by helium injection is one of the most effective methods for suppressing bulk boiling and keeping subcooled propellant conditions for pre-launch, launch, and post-launch pressurization applications. For tank pressurization, submerged helium injection can substantially reduce helium consumption by infusing gaseous propellant into the tank ullage. This paper describes a thermodynamic model of the helium bubbling process in liquid oxygen to estimate the amount of oxygen vapor absorbed by the rising helium bubbles and the amount of subcooling of liquid oxygen due to evaporative heat and mass transfer. The physics of helium dissolution during the pressurization process is also modeled, primarily for liquid hydrogen propellant where the dissolution is more significant. The analyses were performed in a simulation model of tank pressurization built with Generalized Fluid System Simulation Program (GFSSP), a general-purpose flow network code developed at NASA/Marshall Space Flight Center. The numerical predictions of subcooling have been compared with the experimental data of Cho et al. which investigated the propellant subcooling effect as a function of system pressure, helium injection temperature, and flowrate for a non-drained submerged injection system. The numerical predictions of helium consumption have been compared with the test data from a NASA Centaur test vehicle which included both direct and submerged injection with draining of propellants. Hydrogen propellant testing of the Cryogenic Propellant Storage and Transfer Engineering Developmental Unit (CPST EDU) conducted at NASA/Glenn Research Facility was also analyzed. The data used for the model validation were taken in 1-g, but the model was developed to be applicable in both multi-g and micro-g environments.

tank pressurization↗
NASA NTRS1974

ASRDI oxygen technology survey. Volume 4: Low temperature measurement

Information is presented on temperature measurement between the triple point and critical point of liquid oxygen. The criterion selected is that all transducers which may reasonably be employed in the liquid oxygen (LO2) temperature range are considered. The temperature range for each transducer is the appropriate full range for the particular thermometer. The discussion of each thermometer or type of thermometer includes the following information: (1) useful temperature range, (2) general and particular methods of construction and the advantages of each type, (3) specifications (accuracy, reproducibility, response time, etc.), (4) associated instrumentation, (5) calibrations and procedures, and (6) analytical representations.

Sparks, L. L.↗
NASA NTRS2016

Thermal Propulsion Capture System Heat Exchanger Design

One of the biggest challenges of manned spaceflight beyond low earth orbit and the moon is harmful radiation that astronauts would be exposed to on their long journey to Mars and further destinations. Using nuclear energy has the potential to be a more effective means of propulsion compared to traditional chemical engines (higher specific impulse). An upper stage nuclear engine would allow astronauts to reach their destination faster and more fuel efficiently. Testing these engines poses engineering challenges due to the need to totally capture the engine exhaust. The Thermal Propulsion Capture System is a concept for cost effectively and safely testing Nuclear Thermal Engines. Nominally, hydrogen exhausted from the engine is not radioactive, but is treated as such in case of fuel element failure. The Thermal Propulsion Capture System involves injecting liquid oxygen to convert the hydrogen exhaust into steam. The steam is then cooled and condensed into liquid water to allow for storage. The Thermal Propulsion Capture System concept for ground testing of a nuclear powered engine involves capturing the engine exhaust to be cooled and condensed before being stored. The hydrogen exhaust is injected with liquid oxygen and burned to form steam. That steam must be cooled to saturation temperatures before being condensed into liquid water. A crossflow heat exchanger using water as a working fluid will be designed to accomplish this goal. Design a cross flow heat exchanger for the Thermal Propulsion Capture System testing which: Eliminates the need for water injection cooling, Cools steam from 5800 F to saturation temperature, and Is efficient and minimizes water requirement.

Richard, Evan M.↗
NASA NTRS2001

Evaluation of Design Concepts for Collapsible Cryogenic Storage Vessels

Future long-duration missions to Mars using in situ resource production to obtain oxygen from the Martian atmosphere for use as a propellant or for life support will require long term oxygen storage facilities. This report describes preliminary analysis of design concepts for lightweight, collapsible liquid oxygen storage tanks to be used on the surface of Mars. With storage at relatively low pressures, an inflatable tank concept in which the cryogen is stored within a fiber-reinforced Teflon FEP bladder is an efficient approach. The technology required for such a tank is well-developed through similar previous applications in positive expulsion bladders for zero-g liquid fuel rocket tanks and inflatable space habitat technology, though the liquid oxygen environment presents unique challenges. The weight of the proposed structure is largely dominated by the support structure needed to hold the tank off the ground and permit a vacuum insulation space to be maintained around the tank. In addition to the inflatable tank concept, telescoping tank concepts are studied. For a telescoping tank, the greatest difficulty is in making effective joints and seals. The use of shape memory alloy to produce a passive clamping ring is evaluated. Although the telescoping tank concepts are a viable option, it appears that inflatable tank concepts will be more efficient and are recommended.

Fleming, David C.↗
NASA NTRS1999

Propellant Feed System Leak Detection: Lessons Learned From the Linear Aerospike SR-71 Experiment (LASRE)

This paper presents pertinent results and assessment of propellant feed system leak detection as applied to the Linear Aerospike SR-71 Experiment (LASRE) program flown at the NASA Dryden Flight Research Center, Edwards, California. The LASRE was a flight test of an aerospike rocket engine using liquid oxygen and high-pressure gaseous hydrogen as propellants. The flight safety of the crew and the experiment demanded proven technologies and techniques that could detect leaks and assess the integrity of hazardous propellant feed systems. Point source detection and systematic detection were used. Point source detection was adequate for catching gross leakage from components of the propellant feed systems, but insufficient for clearing LASRE to levels of acceptability. Systematic detection, which used high-resolution instrumentation to evaluate the health of the system within a closed volume, provided a better means for assessing leak hazards. Oxygen sensors detected a leak rate of approximately 0.04 cubic inches per second of liquid oxygen. Pressure sensor data revealed speculated cryogenic boiloff through the fittings of the oxygen system, but location of the source(s) was indeterminable. Ultimately, LASRE was cancelled because leak detection techniques were unable to verify that oxygen levels could be maintained below flammability limits.

Hass, Neal↗
NASA NTRS1995

Atlas SOHO Booster and Centaur Erection

The launch vehicle for the Solar Heliospheric Observatory (SOHO) mission is a two stage Atlas-IIAS (Atlas/Centaur). The Atlas, consists of a solid rocket booster stage powered by four Thiokol Castor IVA solid rocket boosters (SRB) and a core vehicle stage (booster and sustainer) powered by Rocketdyne MA-5A liquid propellant engines (RP-1 fuel and liquid oxygen). The multiple firing Centaur is powered by two Pratt and Whitney (RL10A-4) liquid hydrogen and liquid oxygen engines with extendible nozzles. This video shows the erection of the Atlas booster and transportation (to 36-B launching pad) and erection of the Centaur.

Source record↗
NASA NTRSDate not supplied

Nodal Numerical Modeling of Submerged Helium Injection in a Cryogenic Propellant Tank

Subcooling of cryogenic propellant by helium injection is one of the most effective methods for suppressing bulk boiling and keeping subcooled propellant conditions for pre-launch, launch, and post-launch pressurization applications. For tank pressurization, submerged helium injection can substantially reduce helium consumption by infusing gaseous propellant into the tank ullage. This paper presents a mathematical model of the helium bubbling process in liquid oxygen to estimate the amount of oxygen vapor absorbed by the rising helium bubbles and the amount of subcooling of liquid oxygen due to evaporative heat and mass transfer. This mathematical model was incorporated in a simulation model of tank pressurization built with Generalized Fluid System Simulation Program (GFSSP), a general-purpose flow network code developed at NASA/Marshall Space Flight Center. The numerical predictions of subcooling have been compared with the experimental data of Cho et al. which investigated the propellant subcooling effect as a function of system pressure, helium injection temperature, and flowrate for a non-drained submerged injection system. The numerical predictions of helium consumption have been compared with the test data from a NASA Centaur test vehicle which included both direct and submerged injection with draining of propellants. The numerical model developed with GFSSP has been validated against two sets of experimental data and has been shown to predict both propellant subcooling and helium consumption to within 30% in most cases. The test data used for the model validation were taken in 1-g, but the mass diffusion model was developed to be applied in both 1-g and micro-g environments.

Cryogenics↗
NASA NTRS1977

Simulation of the heat transfer characteristics of LOX

In connection with proposals for a second generation shuttle vehicle it has been suggested that the engine regenerative coolant be the oxidizer rather than a fuel. The feasibility of such an approach depends on the suitability of oxygen for the cooling functions. The information currently available concerning the heat transfer characteristics of liquid oxygen (LOX) in the near critical region at elevated pressures is not sufficient for an evaluation. It is, therefore, proposed to make use of data from similar fluids for a simulation of the heat transfer characteristics of LOX. Graphs are presented which demonstrate that experimental heat transfer results for fluid nitrogen can be used qualitatively to simulate heat transfer to liquid oxygen. Quantitative agreement appears satisfactory provided the proper level or adjusting constant can be determined.

Hendricks, R. C.↗
NASA NTRS1989

Hybrid propulsion technology program: Phase 1, volume 4

The use of a liquid oxidizer-solid fuel hybrid propellant combination in booster rocket motors appears extremely attractive due to the integration of the best features of liquid and solid propulsion systems. The hybrid rocket combines the high performance, clean exhaust, and safety of liquid propellant engines with the low cost and simplicity of solid propellant motors. Additionally, the hybrid rocket has unique advantages such as an inert fuel grain and a relative insensitivity to fuel grain and oxidizer injection anomalies. The advantages mark the hybrid rocket as a potential replacement or alternative for current and future solid propellant booster systems. The issues are addressed and recommendations are made concerning oxidizer feed systems, injectors, and ignition systems as related to hybrid rocket propulsion. Early in the program a baseline hybrid configuration was established in which liquid oxygen would be injected through ports in a solid fuel whose composition is based on hydroxyl terminated polybutadiene (HTPB). Liquid oxygen remained the recommended oxidizer and thus all of the injector concepts which were evaluated assumed only liquid would be used as the oxidizer.

Claflin, S. E.↗
NASA NTRS1987

Protecting a Ball-Bearing-Deflection Monitor

Deflectometer probe monitors deflection of ball-bearing race in liquid oxygen with aid of small window or diaphragm. Diaphragm or window isolates optical deflectometer from liquid oxygen or other fluid in ball bearing. At high pressures, diaphragm integral part of housing preferable to window, since there would be no leakage.

Kuhr, George A.↗
NASA NTRS1990

A superconducting quenchgun for delivering lunar derived oxygen to lunar orbit

This study defines a superconducting electromagnetic lunar launcher and its necessary support systems. This launcher is to deliver lunar-derived liquid oxygen to lunar orbit. An Electromagnetic Launcher and Systems Sizing Model developed by the Large Scale Programs Institute scales the launcher and supporting systems as a function of liquid oxygen payload and time between launches. The impact an electromagnetic lunar launcher would have to the present NASA Office of Exploration Lunar Evolution Case Study is examined for three lunar oxygen producing scenarios.

Korsmeyer, David↗
NASA NTRS2017

Oxygen Candle Background for Subs and Space

"At any time and without warning, a submarine may have to remain submerged for several days on account of the presence of the enemy, or rough weather, or serious accident to the machinery. Fortunately such occurrences are rare; but every commanding officer must be prepared to meet such an emergency that will afford his men the greatest possible chance of survival." Reference (1) This quotation is taken from a review of submarine air purification technology published by the Bureau of Medicine and Surgery in 1919. At that time, the U.S. Navy had just begun to experiment with possible air purification devices and supplies of oxygen that might permit submarines to remain submerged longer that the untreated closed atmosphere would allow. Submariners were exposed to elevated levels of carbon dioxide and reduced levels of oxygen that would be considered completely unacceptable by current standards. It was a different world, but humans are still humans, and the requirements for safe and effective functioning in a self-contained environment are really unchanged. The maximum submergence time for submarines as published in that work was approximately 48 hours, reference (1) In early submarines, the preferred supply of oxygen was 1800 psig compressed gas bled into the boat as needed, references (1,2). The need for added oxygen was occasionally "measured" by the physiological impact on the crew rather than a reliable instrument, reference (1). The design of submarine oxygen supply was limited to approximately 25 day submerged operation, reference (2). It was not until 1958 that U.S. submarines were able to carry out dives beyond that period and necessitated new sources of oxygen, reference (2). A curious second source of oxygen at the time was compressed air vessels that were bled into the boat while "vitiated" air (air with reduced oxygen and elevated carbon dioxide) from the opposite end of the boat was compressed into waiting empty vessels, reference (1). The periodic use of pressurized air to control oxygen caused swings in ambient pressure that was uncomfortable to the crew, reference (1). As early as 1919, liquid oxygen was a commercially available product and its use on submarines was contemplated, reference (1). It is interesting that even at this date, 1919, the danger of oils or greases when exposed to compressed or liquid oxygen was recognized: "One precaution must be always taken. No oil or grease should be used in the gauge or fittings or a dangerous explosion may result."

Graf, John↗
NASA NTRS1987

Vibration-Free Vanes Direct Cryogenic Flow

Upstream fluid-dynamic surface prevents oscillations. Simple structural addition to liquid-oxygen manifold prevents vibration and allows increased flow. T-shaped manifold uses vanes to distribute liquid oxygen. Far from vanes, spoiler suppresses vibrations caused by unstable dynamic coupling between flow and elasticity of vanes.

Rao, Gadicherla V. R.↗
NASA NTRS2001

Overview of Rotating Cavitation and Cavitation Surge in the Fastrac Engine LOX Turbopump

Observations regarding rotating cavitation and cavitation surge experienced during the development of the Fastrac 60 Klbf engine turbopump are discussed. Detailed observations from the analysis of both water flow and liquid oxygen test data are offered. Scaling and general comparison of rotating cavitation between water flow and liquid oxygen testing are discussed. Complex data features linking the localized rotating cavitation mechanism of the inducer to system surge components are described in detail. Finally a description of a simple lumped-parameter hydraulic system model developed to better understand observed data is given.

Zoladz, Thomas↗