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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 55 records · Page 3
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NASA NTRSDate not supplied

The Space Superhighway: A Cost Analysis of an In-Space Logistics Resupply Network

The Space Superhighway is a concept for an in-space logistics network for payload delivery and propellant resupply within the Earth-Moon system. It is expected that in the next several years the number of spacecraft in cislunar space will increase significantly. Commercial launch service providers have significantly reduced the cost of access to space and reduced the barrier to entry for private companies to build and operate their own satellites. Many of these newer satellites are being designed to take advantage of future In-space Servicing, Assembly, and Manufacturing (ISAM) capabilities. In addition, the National Aeronautics and Space Administration (NASA) intends to develop and maintain a sustained human presence in cislunar space and on the surface of the Moon. The sustained lunar presence will then serve as a steppingstone for human missions to Mars. The combined demand for propellant resupply and payload delivery in cislunar space for NASA's Moon to Mars campaign in addition to commercial ventures will likely exceed 1,000 t in the upcoming decades. This work presents a concept for an in-space logistics network to fulfill that demand. The network consists of commercial launch vehicles, propellant tankers, orbital depots, and in-space electric propulsion tugs. This work specifically analyzes the cost of developing, producing, and operating such a network at varying levels of customer demand.

nrho↗
NASA NTRS2009

Orion Guidance and Control Ascent Abort Algorithm Design and Performance Results

During the ascent flight phase of NASA s Constellation Program, the Ares launch vehicle propels the Orion crew vehicle to an agreed to insertion target. If a failure occurs at any point in time during ascent then a system must be in place to abort the mission and return the crew to a safe landing with a high probability of success. To achieve continuous abort coverage one of two sets of effectors is used. Either the Launch Abort System (LAS), consisting of the Attitude Control Motor (ACM) and the Abort Motor (AM), or the Service Module (SM), consisting of SM Orion Main Engine (OME), Auxiliary (Aux) Jets, and Reaction Control System (RCS) jets, is used. The LAS effectors are used for aborts from liftoff through the first 30 seconds of second stage flight. The SM effectors are used from that point through Main Engine Cutoff (MECO). There are two distinct sets of Guidance and Control (G&C) algorithms that are designed to maximize the performance of these abort effectors. This paper will outline the necessary inputs to the G&C subsystem, the preliminary design of the G&C algorithms, the ability of the algorithms to predict what abort modes are achievable, and the resulting success of the abort system. Abort success will be measured against the Preliminary Design Review (PDR) abort performance metrics and overall performance will be reported. Finally, potential improvements to the G&C design will be discussed.

Proud, Ryan W.↗
NASA NTRSDate not supplied

Venus Sample Return Using In-situ Propellant

This project will pioneer a new approach to return a sample from the surface of Venus. At 450°C and 92 bar pressure, the Venus surface is the most hostile environment in the solar system. This project merges an innovative carbon monoxide rocket concept using propellant made from the Venus atmosphere with innovations in high-temperature technology and solar aircraft. We turn an ambitious mission into a reality. The mission will return a sample from the scorching surface of Venus, using a novel multi-step approach incorporating new technology. The high-temperature solar aircraft collects the sample from the surface and carries it above the main cloud deck to a balloon platform with a launch vehicle & propellant manufacturing plant. The carbon monoxide/oxygen monopropellant is manufactured from the Venus atmosphere.

Venus↗
NASA NTRS2010

Liquid Oxygen Propellant Densification Production and Performance Test Results With a Large-Scale Flight-Weight Propellant Tank for the X33 RLV

This paper describes in-detail a test program that was initiated at the Glenn Research Center (GRC) involving the cryogenic densification of liquid oxygen (LO2). A large scale LO2 propellant densification system rated for 200 gpm and sized for the X-33 LO2 propellant tank, was designed, fabricated and tested at the GRC. Multiple objectives of the test program included validation of LO2 production unit hardware and characterization of densifier performance at design and transient conditions. First, performance data is presented for an initial series of LO2 densifier screening and check-out tests using densified liquid nitrogen. The second series of tests show performance data collected during LO2 densifier test operations with liquid oxygen as the densified product fluid. An overview of LO2 X-33 tanking operations and load tests with the 20,000 gallon Structural Test Article (STA) are described. Tank loading testing and the thermal stratification that occurs inside of a flight-weight launch vehicle propellant tank were investigated. These operations involved a closed-loop recirculation process of LO2 flow through the densifier and then back into the STA. Finally, in excess of 200,000 gallons of densified LO2 at 120 oR was produced with the propellant densification unit during the demonstration program, an achievement that s never been done before in the realm of large-scale cryogenic tests.

Tomsik, Thomas M.↗
NASA NTRS1997

Fuels and Space Propellants for Reusable Launch Vehicles: A Small Business Innovation Research Topic and Its Commercial Vision

Under its Small Business Innovation Research (SBIR) program (and with NASA Headquarters support), the NASA Lewis Research Center has initiated a topic entitled "Fuels and Space Propellants for Reusable Launch Vehicles." The aim of this project would be to assist in demonstrating and then commercializing new rocket propellants that are safer and more environmentally sound and that make space operations easier. Soon it will be possible to commercialize many new propellants and their related component technologies because of the large investments being made throughout the Government in rocket propellants and the technologies for using them. This article discusses the commercial vision for these fuels and propellants, the potential for these propellants to reduce space access costs, the options for commercial development, and the benefits to nonaerospace industries. This SBIR topic is designed to foster the development of propellants that provide improved safety, less environmental impact, higher density, higher I(sub sp), and simpler vehicle operations. In the development of aeronautics and space technology, there have been limits to vehicle performance imposed by traditionally used propellants and fuels. Increases in performance are possible with either increased propellant specific impulse, increased density, or both. Flight system safety will also be increased by the use of denser, more viscous propellants and fuels.

Palaszewski, Bryan A.↗
NASA NTRS2014

A Monte Carlo Approach to Modeling the Breakup of the Space Launch System EM-1 Core Stage with an Integrated Blast and Fragment Catalogue

The Liquid Propellant Fragment Overpressure Acceleration Model (L-FOAM) is a tool developed by Bangham Engineering Incorporated (BEi) that produces a representative debris cloud from an exploding liquid-propellant launch vehicle. Here it is applied to the Core Stage (CS) of the National Aeronautics and Space Administration (NASA) Space Launch System (SLS launch vehicle). A combination of Probability Density Functions (PDF) based on empirical data from rocket accidents and applicable tests, as well as SLS specific geometry are combined in a MATLAB script to create unique fragment catalogues each time L-FOAM is run-tailored for a Monte Carlo approach for risk analysis. By accelerating the debris catalogue with the BEi blast model for liquid hydrogen / liquid oxygen explosions, the result is a fully integrated code that models the destruction of the CS at a given point in its trajectory and generates hundreds of individual fragment catalogues with initial imparted velocities. The BEi blast model provides the blast size (radius) and strength (overpressure) as probabilities based on empirical data and anchored with analytical work. The coupling of the L-FOAM catalogue with the BEi blast model is validated with a simulation of the Project PYRO S-IV destruct test. When running a Monte Carlo simulation, L-FOAM can accelerate all catalogues with the same blast (mean blast, 2 σ blast, etc.), or vary the blast size and strength based on their respective probabilities. L-FOAM then propagates these fragments until impact with the earth. Results from L-FOAM include a description of each fragment (dimensions, weight, ballistic coefficient, type and initial location on the rocket), imparted velocity from the blast, and impact data depending on user desired application. LFOAM application is for both near-field (fragment impact to escaping crew capsule) and far-field (fragment ground impact footprint) safety considerations. The user is thus able to use statistics from a Monte Carlo set of L-FOAM catalogues to quantify risk for a multitude of potential CS destruct scenarios. Examples include the effect of warning time on the survivability of an escaping crew capsule or the maximum fragment velocities generated by the ignition of leaking propellants in internal cavities.

Richardson, Erin↗
NASA NTRS1993

Advanced electric motor technology flux mapping

Design of electric motors which fulfill the needs of Thrust Vector Control (TVC) actuators used in large rocket propelled launch vehicles is covered. To accomplish this end the methodology of design is laid out in some detail. In addition a point design of a motor to fulfill the requirements of a certain actuator specified by MSFC is accomplished and reported upon. In the course of this design great stress has been placed on ridding the actuator of internally generated heat. To conduct the heat out of the motor use is made of the unique properties of the in house MSFC designed driving electronics. This property is that as along as they are operated in a quasi-linear mode the electronics nullify the effects of armature inductance as far as the phase of the armature current versus the rotor position is concerned. Actually the additional inductance due to the extended end turns in this design is of benefit because in the shorted armature failure mode the armature current in the fault (caused by the rotor flux sweeping past the armature) is diminished at a given rotor speed when compared to a more conventional motor with lower inductance. The magnetic circuit is analyzed using electromagnetic finite element methods.

Doane, George B., III↗
NASA NTRS2001

Trends and Options in the Disposal of Launch Vehicle Orbital Stages

An increasing number of agencies and organizations around the world provide guidelines for the disposal of launch vehicle stages placed in Earth orbit. The limitation of orbital lifetime following the satellite delivery mission, particularly in low Earth orbit, is viewed as one of the most important space debris mitigation measures. For higher altitude missions, orbital stages may be maneuvered into more rapidly decaying orbits or may be left in longer-term disposal orbits. This paper summarizes the recent disposal practices of all launch vehicle types and evaluates their compliance with existing national standards. Although the owners and operators of some satellite systems, e.g., Iridium and Globalstar, have levied orbital stage disposal requirements on launch service providers, in general, insufficient attention and communication is given to this topic. Sunsynchronous and geosynchronous missions may pose some of the most difficult challenges for the responsible disposal of orbital stages. A variety of disposal options are normally available, some of which may even influence the design of the spacecraft to be deployed. Also at issue is whether normal launch vehicle propellant reserves can be relied upon for post-mission disposal of orbital stages.

Johnson, Nicholas L.↗
NASA NTRS2009

Propellant Mass Fraction Calculation Methodology for Launch Vehicles and Application to Ares Vehicles

Propellant Mass Fraction (pmf) calculation methods vary throughout the aerospace industry. While typically used as a means of comparison between candidate launch vehicle designs, the actual pmf calculation method varies slightly from one entity to another. It is the purpose of this paper to present various methods used to calculate the pmf of launch vehicles. This includes fundamental methods of pmf calculation that consider only the total propellant mass and the dry mass of the vehicle; more involved methods that consider the residuals, reserves and any other unusable propellant remaining in the vehicle; and calculations excluding large mass quantities such as the installed engine mass. Finally, a historical comparison is made between launch vehicles on the basis of the differing calculation methodologies, while the unique mission and design requirements of the Ares V Earth Departure Stage (EDS) are examined in terms of impact to pmf.

Holt, James B.↗
NASA NTRS1999

Propellant Densification for Launch Vehicles: Simulation and Testing 1999

One of the many key technologies required to make single-stage to orbit an actuality, the technology of sub-cooling cryogenic propellants below their normal saturation temperatures and thereby making them more dense, is unquestionably on its way towards full and practical realization. The technology of Propellant Densification has been the subject of an extensive research and development program overseen by Lockheed Martin Michoud Space Systems and NASA Glenn Research Center over the past several years. This paper presents a status report of this research and development. Specifically examined within this paper is the status of the current and continuing efforts on the mathematical simulation of the in-tank propellant densification process currently baselined for the Lockheed Martin VentureStar Reusable Launch Vehicle (RLV). Keys to this modeling effort are an understanding and quantification of the effects of thermal stratification and the ability to capture the complex and unique multiple section tank geometries being proposed for future launch vehicles. A simulation that properly captures these phenomena has been developed by Lockheed Martin. Also discussed is the significant test program that has been undertaken in coordination with NASA Glenn Research Center. In this testing, the liquid hydrogen recirculation and densification process was simulated and the thermal stratification of the densified propellant was recorded throughout the tank. This testing marks the first time that such a process has been carried out within a multiple-lobe, flight-similar tank. The results from this testing have gone a long way towards grounding the mathematical models and towards demonstrating the readiness of the technology for near-term use. A further and even more ambitious test program examining the production and utilization of densified propellants is being planned for late-autumn 1999. An overview of these plans is presented.

Knowles, Timothy E.↗
NASA NTRS1966

Analysis of cryogenic propellant logistics.

Large launch vehicle cryogenic propellant logistics including storage and production capacity optimization, cost and heat loss analyses by computer simulation

PROPELLANT STORAGE↗
NASA NTRS2011

Experimental and Numerical Investigation of Reduced Gravity Fluid Slosh Dynamics for the Characterization of Cryogenic Launch and Space Vehicle Propellants

As space programs increasingly investigate various options for long duration space missions the accurate prediction of propellant behavior over long periods of time in microgravity environment has become increasingly imperative. This has driven the development of a detailed, physics-based understanding of slosh behavior of cryogenic propellants over a range of conditions and environments that are relevant for rocket and space storage applications. Recent advancements in computational fluid dynamics (CFD) models and hardware capabilities have enabled the modeling of complex fluid behavior in microgravity environment. Historically, launch vehicles with moderate duration upper stage coast periods have contained very limited instrumentation to quantify propellant stratification and boil-off in these environments, thus the ability to benchmark these complex computational models is of great consequence. To benchmark enhanced CFD models, recent work focuses on establishing an extensive experimental database of liquid slosh under a wide range of relevant conditions. In addition, a mass gauging system specifically designed to provide high fidelity measurements for both liquid stratification and liquid/ullage position in a micro-gravity environment has been developed. This pUblication will summarize the various experimental programs established to produce this comprehensive database and unique flight measurement techniques.

Walls, Laurie K.↗