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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 73 records · Page 4
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NASA NTRSDate not supplied

MUSTANG: A Workhorse for NASA Spaceflight Avionics

The Modular Unified Space Technology Avionics for Next Generation (MUSTANG) is a small integrated Avionics system including Command and Data Handling (C&DH), Power System Electronics (PSE), Attitude Control System Interfaces (ACS), and Propulsion Electronics. The MUSTANG Avionics Architecture is built upon many years of knowledge capture and lessons learned at the Goddard Space Flight Center. With a motivation towards modularity and keeping board redesign costs to a minimum, MUSTANG offers flexibility in features with a backplane-less design and allows the user to choose the options (cards) needed for their system. It incorporates a distributed power system that provides secondary power to all its subcomponents reducing the number of primary services needed for an Avionics. MUSTANG can be integrated into one system or divided into several smaller components. MUSTANG supports redundancy and cross-strap ability for a more robust and reliable Avionics system. A variation of MUSTANG exists for Instrument Electronics called iMUSTANG and allows the user to select functionality applicable to the instrument electronics. MUSTANG is not meant to replace Avionics for all spacecraft. There are limitations due to its relatively compact size, but the MUSTANG design has proven broadly applicable on many spacecraft and instrument bus avionics architectures.

MUSTANG↗
NASA NTRS1980

A low-cost molecular-leak value

Solenoid operated modular-leak and shutoff valve has been developed for small portable, automated, mass spectrometer used to measure trace constituents of air or other gases. Valve costs much less to produce than precision needle-in-foil type. Yet its performance closely matches that version.

Judson, C. M.↗
NASA NTRS1993

A multimission three-axis stabilized spacecraft flight dynamics ground support system

The Multimission Three-Axis Stabilized Spacecraft (MTASS) Flight Dynamics Support System (FDSS) has been developed in an effort to minimize the costs of ground support systems. Unlike single-purpose ground support systems, which attempt to reduce costs by reusing software specifically developed for previous missions, the multimission support system is an intermediate step in the progression to a fully generalized mission support system in which numerous missions may be served by one general system. The benefits of multimission attitude ground support systems extend not only to the software design and coding process, but to the entire system environment, from specification through testing, simulation, operations, and maintenance. This paper reports the application of an MTASS FDSS to multiple scientific satellite missions. The satellites are the Upper Atmosphere Research Satellite (UARS), the Extreme Ultraviolet Explorer (EUVE), and the Solar Anomalous Magnetospheric Particle Explorer (SAMPEX). Both UARS and EUVE use the multimission modular spacecraft (MMS) concept. SAMPEX is part of the Small Explorer (SMEX) series and uses a much simpler set of attitude sensors. This paper centers on algorithm and design concepts for a multimission system and discusses flight experience from UARS.

Langston, J.↗
NASA NTRS1996

An Advanced Photovoltaic Array Regulator Module

Current trends in satellite design are focused on developing small, reliable, and inexpensive spacecraft. To that end, a modular power management and distribution system is proposed which will help transition the aerospace industry towards an assembly line approach to building spacecraft. The modular system is based on an innovative DC voltage boost converter called the Series Connected Boost Unit (SCBU). The SCBU uses any isolating DC-DC converter and adds a unique series connection. This simple modification provides the SCBU topology with many advantages over existing boost converters. Efficiencies of 94-98%, power densities above 1,000 W/kg, and inherent fault tolerance are just a few of the characteristics presented. Limitations of the SCBU technology are presented, and it is shown that the SCBU makes an ideal photovoltaic an-ay regulator. A set of photovoltaic power system requirements are presented that can be applied to almost any low Earth orbit satellite. Finally, a modular design based on the series connected boost unit is outlined and functional descriptions of the components are given.

Button, Robert M.↗
NASA NTRS1979

Phase 1 of the First Small Power System Experiment (engineering Experiment No. 1). Volume 5: Supporting Analyses and Trade Studies

The development and design of a modular solar thermal power system for application in the 1 to 10 MWe range is described. The system is used in remote utility applications, small communities, rural areas, and for industrial uses. Thermal and stress analyses are performed on the collector subsystem, energy storage subsystem, energy transport subsystem, the power conversion subsystem, and the plant control subsystem.

Holl, R. J.↗
NASA NTRS2000

Flight Test of a Technology Transparent Light Concentration Panel on SMEX/WIRE

A flight experiment has demonstrated a modular solar concentrator that can be used as a direct substitute replacement for planar photovoltaic panels in spacecraft solar arrays. The Light Concentrating Panel (LCP) uses an orthogrid arrangement of composite mirror strips to form an array of rectangular mirror troughs that reflect light onto standard, high-efficiency solar cells at a concentration ratio of approximately 3:1. The panel area, mass, thickness, and pointing tolerance has been shown to be similar to a planar array using the same cells. Concentration reduces the panel's cell area by 2/3, which significantly reduces the cost of the panel. An opportunity for a flight experiment module arose on NASA's Small Explorer / Wide-Field Infrared Explorer (SMEX/WIRE) spacecraft, which uses modular solar panel modules integrated into a solar panel frame structure. The design and analysis that supported implementation of the LCP as a flight experiment module is described. Easy integration into the existing SMEX-LITE wing demonstrated the benefits of technology transparency. Flight data shows the stability of the LCP module after nearly one year in Low Earth Orbit.

Stern, Theodore G.↗
NASA NTRS1996

Metallized Gelled Propellants: Oxygen/RP-1/Aluminum Rocket Heat Transfer and Combustion Measurements

A series of rocket engine heat transfer experiments using metallized gelled liquid propellants was conducted. These experiments used a small 20- to 40-lb/f thrust engine composed of a modular injector, igniter, chamber and nozzle. The fuels used were traditional liquid RP-1 and gelled RP-1 with 0-, 5-, and 55-percentage by weight loadings of aluminum particles. Gaseous oxygen was used as the oxidizer. Three different injectors were used during the testing: one for the baseline O(2)/RP-1 tests and two for the gelled and metallized gelled fuel firings. Heat transfer measurements were made with a rocket engine calorimeter chamber and nozzle with a total of 31 cooling channels. Each chamber used a water flow to carry heat away from the chamber and the attached thermocouples and flow meters allowed heat flux estimates at each of the 31 stations. The rocket engine Cstar efficiency for the RP-1 fuel was in the 65-69 percent range, while the gelled 0 percent by weight RP-1 and the 5-percent by weight RP-1 exhibited a Cstar efficiency range of 60 to 62% and 65 to 67%, respectively. The 55-percent by weight RP-1 fuel delivered a 42-47% Cstar efficiency. Comparisons of the heat flux and temperature profiles of the RP-1 and the metallized gelled RP-1/A1 fuels show that the peak nozzle heat fluxes with the metallized gelled O2/RP-1/A1 propellants are substantially higher than the baseline O2/RP-1: up to double the flux for the 55 percent by weight RP-1/A1 over the RP-1 fuel. Analyses showed that the heat transfer to the wall was significantly different for the RP-1/A1 at 55-percent by weight versus the RP-1 fuel. Also, a gellant and an aluminum combustion delay was inferred in the 0 percent and 5-percent by weight RP-1/A1 cases from the decrease in heat flux in the first part of the chamber. A large decrease in heat flux in the last half of the chamber was caused by fuel deposition in the chamber and nozzle. The engine combustion occurred well downstream of the injector face based on the heat flux estimates from the temperature measurements.

Palaszewski, Bryan↗
NASA NTRS1997

Metallized Gelled Propellants: Oxygen/RP-1/Aluminum Rocket Engine Calorimeter Heat Transfer Measurements and Analysis

A set of analyses was conducted to determine the heat transfer characteristics of metallized gelled liquid propellants in a rocket engine. The analyses used the data from experiments conducted with a small 30- to 40-lbf thrust engine composed of a modular injector, igniter, chamber and nozzle. The fuels used were traditional liquid RP-1 and gelled RP-1 with 0-wt %, 5-wt%, and 55-wt% loadings of aluminum with silicon dioxide gellant, and gaseous oxygen as the oxidizer. Heat transfer was computed based on measurements using calorimeter rocket chamber and nozzle hardware with a total of 31 cooling channels. A gelled fuel coating formed in the 0-, 5- and 55-wt% engines, and the coating was composed of unburned gelled fuel and partially combusted RP-1. The coating caused a large decrease in calorimeter engine heat flux in the last half of the chamber for the 0- and 5-wt% RP-1/Al. This heat flux reduction effect was analyzed by comparing engine runs and the changes in the heat flux during a run as well as from run to run. Heat transfer and time-dependent heat flux analyses and interpretations are provided. The 5- and 55-wt% RP-1/Al fueled engines had the highest chamber heat fluxes, with the 5-wt% fuel having the highest throat flux. This result is counter to the predicted result, where the 55 wt% fuel has the highest combustion and throat temperature, and therefore implies that it would deliver the highest throat heat flux. The 5-wt% RP-1/Al produced the most influence on the engine heat transfer and the heat flux reduction was caused by the formation of a gelled propellant layer in the chamber and nozzle.

Palaszewski, Bryan↗
NASA NTRS1999

New Approaches to Parameterizing Convection

Many general circulation models (GCMs) currently use separate schemes for planetary boundary layer (PBL) processes, shallow and deep cumulus (Cu) convection, and stratiform clouds. The conventional distinctions. among these processes are somewhat arbitrary. For example, in the stratocumulus-to-cumulus transition region, stratocumulus clouds break up into a combination of shallow cumulus and broken stratocumulus. Shallow cumulus clouds may be considered to reside completely within the PBL, or they may be regarded as starting in the PBL but terminating above it. Deeper cumulus clouds often originate within the PBL with also can originate aloft. To the extent that our models separately parameterize physical processes which interact strongly on small space and time scales, the currently fashionable practice of modularization may be doing more harm than good.

Randall, David A.↗
NASA NTRS1998

Metallized Gelled Propellants: Heat Transfer of a Rocket Engine Fueled by Oxygen/RP-1/Aluminum Was Measured by a Calorimeter

A set of analyses was conducted to determine the heat transfer characteristics of metallized gelled liquid propellants in a rocket engine. These analyses used data from experiments conducted with a small 30- to 40-lbf thrust engine composed of a modular injector, igniter, chamber, and nozzle. The fuels used were traditional liquid RP-1 and gelled RP-1 with 0-, 5-, and 55-wt % loadings of aluminum (Al) with gaseous oxygen as the oxidizer. Heat transfer measurements were made with a calorimeter chamber and nozzle setup that had a total of 31 cooling channels. A gelled fuel coating, composed of unburned gelled fuel and partially combusted RP-1, formed in the 0-, 5- and 55-wt % engines. For the 0- and 5-wt % RP-1/Al, the coating caused a large decrease in calorimeter engine heat flux in the last half of the chamber. This heat flux reduction was analyzed by comparing engine firings and the changes in the heat flux during a firing at NASA Lewis Research Center's Rocket Laboratories. This work is part of an ongoing series of analyses of metallized gelled propellants.

Palaszewski, Bryan A.↗
NASA NTRS1997

Metallized Gelled Propellant Heat Transfer Tests Analyzed

A series of rocket engine heat transfer experiments using metallized gelled liquid propellants was conducted at the NASA Lewis Research Center. These experiments used a small 20- to 40-lbf thrust engine composed of a modular injector, an igniter, a chamber, and a nozzle. The fuels used were traditional liquid RP-1 and gelled RP-1 with 0-, 5-, and 55-wt % loadings of aluminum particles. Gaseous oxygen was used as the oxidizer. Heat transfer measurements were made with a rocket engine calorimeter chamber and nozzle with a total of 31 cooling channels. Each channel used water flow to carry heat away from the chamber and the attached thermocouples; flow meters allowed heat flux estimates at each of the 31 stations.

Palaszewski, Bryan A.↗
NASA NTRS2017

Space Technology Game Changing Development: NASA Affordable Vehicle Avionics (AVA)

Common Modular Avionics System for Nanolaunchers Offering Affordable Access to Space Small satellites are becoming ever more capable of performing valuable missions for both government and commercial customers. However, currently these satellites can be launched affordably only as secondary payloads. This makes it difficult for the small satellite mission to launch when needed, to the desired orbit, and with acceptable risk. What is needed is a class of low-cost launchers, so that launch costs to low-Earth orbit (LEO) are commensurate with payload costs. Several private and government-sponsored launch vehicle developers are working toward just that—the ability to affordably insert small payloads into LEO. But until now, cost of the complex avionics remained disproportionately high. AVA solves this problem. Significant contributors to the cost of launching nanosatellites to orbit are the avionics and software systems that steer and control the launch vehicles, sequence stage separation, deploy payloads, and telemeter data. The high costs of these guidance, navigation and control (GNC) avionics systems are due in part to the current practice of developing unique, single use hardware and software for each launch. High-performance, high-reliability inertial sensors components with heritage from legacy launchers also contribute to costs—but can low-cost commercial inertial sensors work just as well?

Aquilina, Rudy↗
NASA NTRS1990

Space Station Payload Adaptation System

The development and design of a system of containers for the efficient integration of Space Station payloads is described called the Space Station Payload Adaptation System (SSPAS). The SSPAS was developed to address the incorporation of multiple payloads, the use of a small payload carrier, large numbers of samples, and on-orbit servicing. SSPAS subsystems such as the Spacelab rack are modular and designed for integration into the 'Quick Is Beautiful' (QIB) facility. The QIB is designed to provide access to space for small- and medium-sized microgravity research projects and proof-of-concept investigations. The power-distribution and heat-rejection potential of the QIB are described, and an improved experiment-apparatus container is proposed. The SSPAS rack-mounting and container concepts are concluded to make up an efficent system that can effectively exploit the R&D potential of the Space Station.

Taylor, Kenneth R.↗
NASA NTRS1996

Program Simulates A Modular Manufacturing System

SSE computer program provides simulation environment for modeling manufacturing systems containing relatively small numbers of stations and operators. Designed to simulate manufacturing of apparel, also used in other manufacturing domains. Excellent for small or medium-size firms including those lacking expertise to develop detailed models or have only minimal knowledge in describing manufacturing systems and in analyzing results of simulations on mathematical models. User does not need to know simulation language to use SSE. Used to design new modules and to evaluate existing modules. Originally written in Turbo C v2.0 for IBM PC-compatible computers running MS-DOS and successfully implemented by use of Turbo C++ v3.0.

Schroer, Bernard J.↗
NASA NTRS1996

Another Program Simulates A Modular Manufacturing System

SSE5 computer program provides simulation environment for modeling manufacturing systems containing relatively small numbers of stations and operators. Designed to simulate manufacturing of apparel, also used in other manufacturing domains. Valuable for small or medium-size firms, including those lacking expertise to develop detailed mathematical models or have only minimal knowledge in describing manufacturing systems and in analyzing results of simulations on mathematical models. Two other programs available bundled together as SSE (MFS-26245). Each program models slightly different manufacturing scenario. Written in Turbo C v2.0 for IBM PC-series and compatible computers running MS-DOS and successfully compiled using Turbo C++ v3.0.

Schroer, Bernard J.↗
NASA NTRS2020

Soft-Switching GaN-Based Isolated Power Conversion System for Small Satellites with Wide Input Voltage Range

As we pursue the advancement of small satellites for space missions with more capabilities, there is a significant need for cutting-edge, modularly configurable, high density power converters. This article proposes a fixed switching frequency, high efficiency, compact isolated converter for sensitive loads such as radar, communication systems, or other instruments on small satellites.

Cook, Tom↗
NASA NTRS1991

Large constraint length high speed viterbi decoder based on a modular hierarchial decomposition of the deBruijn graph

A method of formulating and packaging decision-making elements into a long constraint length Viterbi decoder which involves formulating the decision-making processors as individual Viterbi butterfly processors that are interconnected in a deBruijn graph configuration. A fully distributed architecture, which achieves high decoding speeds, is made feasible by novel wiring and partitioning of the state diagram. This partitioning defines universal modules, which can be used to build any size decoder, such that a large number of wires is contained inside each module, and a small number of wires is needed to connect modules. The total system is modular and hierarchical, and it implements a large proportion of the required wiring internally within modules and may include some external wiring to fully complete the deBruijn graph. pg,14.

Collins, Oliver↗
NASA NTRS1972

Safety in earth orbit study. Volume 3: Analysis of tumbling spacecraft, escape and rescue

Detailed and supporting analyses of the hazards of a disabled spacecraft in a tumbling mode and of crew escape, rescue, and survivability are presented for earth orbital operations of the space shuttle program. Tumbling conditions, arresting tumbling by external means, and escape from a tumbling vehicle are discussed. Details on design concepts for the modular space station, integral tank orbiter, drop tank orbiter, and a small space vehicle are included. Results of the crew safety study indicate that the shuttle orbiter should be the primary vehicle for dealing with emergencies and one vehicle should be available for rapid emergency rescue. An Apollo command module should be carried in the orbiter cargo bay in case an orbiter is not available. Approaches for a quicker escape or rescue capability are also recommended.

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