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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 577 records · Page 32
All sourcesNASA NTRSDOE OSTI
DOE OSTI2025

SCORE (Synthesis of COnsists as Rolling Energy microgrids) (Final Technical Report)

The SCORE toolset is an open source, web-based application to assess the impact of new powering technologies on railroad performance, specifically technologies capable of both putting power into the rail (motive force) and taking power from the rail (regeneration). SCORE’s primary goal is generate trade studies to analyze different powering and train make-up options to minimize energy usage and greenhouse gas generation. At the core of generating these trade spaces is calculating the optimal powering policy for the train given the makeup of the train, the route, and time constraints. This paper presents details on the algorithms used in SCORE to compute this powering policies that is fast and accurate, discusses its implementation in an Energy-Longitudinal Train Dynamics (E-LTD) model, compares it to naïve approaches, and demonstrates its use across a variety of train/route pairs.

24 POWER TRANSMISSION AND DISTRIBUTION↗
NASA NTRS1985

Aerodynamic performance investigation of advanced mechanical suppressor and ejector nozzle concepts for jet noise reduction

Advanced Supersonic Transport jet noise may be reduced to Federal Air Regulation limits if recommended refinements to a recently developed ejector shroud exhaust system are successfully carried out. A two-part program consisting of a design study and a subscale model wind tunnel test effort conducted to define an acoustically treated ejector shroud exhaust system for supersonic transport application is described. Coannular, 20-chute, and ejector shroud exhaust systems were evaluated. Program results were used in a mission analysis study to determine aircraft takeoff gross weight to perform a nominal design mission, under Federal Aviation Regulation (1969), Part 36, Stage 3 noise constraints. Mission trade study results confirmed that the ejector shroud was the best of the three exhaust systems studied with a significant takeoff gross weight advantage over the 20-chute suppressor nozzle which was the second best.

Wagenknecht, C. D.↗
NASA NTRS1999

Advance Approach to Concept and Design Studies for Space Missions

Recent automated and advanced techniques developed at JPL have created a streamlined and fast-track approach to initial mission conceptualization and system architecture design, answering the need for rapid turnaround of trade studies for potential proposers, as well as mission and instrument study groups.

Team X Team A Missions Design Proposal Development↗
NASA NTRS1972

RAM environment control and life support subsystem.

The Research and Application Modules (RAMs) are shuttle payloads designed to provide economical manned facilities for earth orbital research and development. Support of the habitable environment in the sortie class RAMs can be accomplished by an independent ECLSS installation or by integration with the shuttle ECLSS. The RAM study evaluated advantages and disadvantages of various degrees of the two alternative installations. Selection of the optimum subsystem was made from an overall shuttle/RAM standpoint. These trade studies are presented and the resulting ECLSS support concept are discussed.

Arras, R. P.↗
NASA NTRS1991

Reusable Reentry Satellite (RRS): Telemetry, Tracking, and Command (TT/C) Coverage Tradeoff Study

The Reusable Reentry Satellite (RRS) Telemetry, Tracking & Command (TT&C) Coverage Tradeoff Study described herein was performed during Part 1 of the RRS Phase B contract. This report is one of several that describes the results of various trade studies performed to arrive at a recommended design for the RRS satellite system. The overall RRS Phase B Study objective is to design a relatively inexpensive satellite to access space for extended periods of time, with eventual recovery of experiments on Earth. The RRS will be capable of: (1) being launched by a variety of expendable launch vehicles; (2) operating in low earth orbit as a free flying unmanned laboratory; and (3) executing an independent atmospheric reentry and soft landing. The RRS will be designed to be refurbished and reused up to three times a year for a period of 10 years. The expected principal use for such a system is research on the effects of variable gravity (0 to 1.5 g) and radiation on small animals, plants, lower life forms, tissue samples, and materials processes. This Summary Report provides a description of the RRS Telemetry, Tracking & Command Analysis performed to study the ground station coverage available to meet RRS requirements. Concepts were considered that used off-the-shelf technology, had generous margins of capability, and high reliability, were easily maintained, cost effective, and were compatible with existing support networks. From this study, a recommended RRS TT&C configuration will be developed.

Source record↗
NASA NTRS1991

Design of the cryogenic system for the SAFIRE instrument

One of the primary goals of the NASA Mission to Planet Earth is to improve understanding of the ozone chemistry of the atmosphere over an extended period of time. The Spectroscopy of the Atmosphere using Far Infrared Emission (SAFIRE) instrument is being developed to conduct, for the first time, global measurements of the key ozone chemistry constituents in both the mid- and far-infrared spectral regions. Such remote, long-term observations are made possible by the recent development of compact long-life hybrid cryogenic dewars which are necessary to cool the sensitive detectors to the 3-4 K range. The success of this hybrid concept is based on the use of long-life Stirling cycle cryocoolers to intercept parasitic heat from the internal radiation shields of the superfluid helium dewar. Extensive system trade studies are required to optimize the mass, power, and lifetime of these space-borne cryogenic systems. The SAFIRE Cryogenic Subsystem is described, including the thermal performance trades leading to the chosen system configuration and the important dewar/cryocooler interface issues.

Lee, J. H.↗
NASA NTRS1995

Advanced controls for stability assessment of solar dynamics space power generation

In support of the power requirements for the Space Station Alpha (SSA), a joint program by the U.S. and Russia for a permanently manned space station to be launched into orbit by 1998, a robust control scheme is needed to assure the stability of the rotating machines that will be integrated into the power subsystem. A framework design and systems studies for modeling and analysis is presented. It employs classical d-q axes machine model with voltage/frequency dependent loads. To guarantee that design requirements and necessary trade studies are done, a functional analysis tool CORE is used for the study. This provides us with different control options for stability assessment. Initial studies and recommendations using advanced simulation tools are also presented. The benefits of the stability/control scheme for evaluating future designs and power management are discussed.

Momoh, James A.↗
DOE OSTI2024

TEAMER: Results of Investigating Structural Design Concepts and Alternative Materials for a Wave Power System

Included here are materials from a study on the design of a three-body Wave Energy Converter (WEC) utilizing a heave plate, dual Power Take Offs (PTOs), and single point mooring. A material trade study has been conducted to evaluate the effects of introducing various metallic and fiber-reinforced polymer (FRP) elements to the design. The baseline design is C-Power's patented k2 Wave Power System (WPS). This Testing Expertise and Access for Marine Energy Research (TEAMER) Request for Technical Support (RFTS) 9 effort resulted in alternative designs of the WPS center member, the nacelle. The analysis resulting from this study will inform the design of a lightweight structure with lower manufacturing and operating costs that increases energy conversion and power-to-weight ratio of the device. The dataset contains several components: finite element analysis (FEA) data, hand calculations, and CAD models. FEA files includes those used for and generated by finite element simulations. Calculations include hand calculations performed using Microsoft Excel, which assess core laminate properties, buckling, and other structural aspects in accordance with risk management DNV-RP-C201 standards. CAD materials include Solidworks files for parts, assemblies, and drawings of the evaluated models.

16 TIDAL AND WAVE POWER↗
NASA NTRS2019

Mars Sample Return mission concept status

This paper will provide an overview of current options and specific concepts for a potential Mars Sample Return (MSR) architecture being jointly studied by NASA and ESA. Overall objectives and mission options will be described, including the architecture’s constraints and notional timelines. The paper will highlight architecture-level trade studies, including specific elements and their status. The overall Sample Retrieval Lander (SRL) mission concept, including vehicle options will be described, including the Mars Ascent Vehicle (MAV), Sample Fetch Rover (being studied by ESA), Orbiting Sample container (OS), and tube transfer robotics systems. The concept and status of the Earth Return Orbiter (ERO) mission, being studied by ESA, and the Capture/Containment and Return System (CCRS) which would be the payload on the ERO, will be discussed.

Vijendran, Sanjay↗
NASA NTRS2019

Design of a Nozzle for the Spyder 2nd Stage Solid Rocket Motor

During the 2019 summer term, the author worked with a group of four interns to complete the preliminary design of a 2nd stage solid rocket motor for Up Aerospace’s Spyder Launch Vehicle. The Spyder vehicle is a four stage, solid fuel rocket designed as part of collaboration between NASA and Up Aerospace to develop a vehicle capable of delivering a 10 kg, 6U CubeSat into a 350 km, circular, low Earth orbit. As part of the agreement, NASA is tasked with designing high performance 2nd, 3rd, and 4th stages while Up Aerospace will provide the 1st stage, derived from the first stage of the company’s sub-orbital Spaceloft XL vehicle. Previous intern teams have designed the 3rd and 4th stages, which left the preliminary design of the 2nd stage motor to be completed this summer. The purpose of this report is to highlight a trade study which the author conducted to determine the nozzle geometry which would most benefit the performance of the 2nd stage motor. In this study, various nozzle parameters such as throat radius (RSI), expansion ratio, mass and their effects on the Isp and Delta V of the 2nd Stage were investigated. From this study, a nozzle geometry providing the necessary performance was chosen and implemented as part of the preliminary design of the 2nd stage motor. To mature the 2nd stage motor design, a trade space was needed to determine the nozzle configuration which would most benefit the performance of the 2nd stage. The trade space established did not only evaluate different expansion ratios for the same throat radius, but also investigated the possible performance gained from decreasing the throat radius to increase the expansion ratio and Isp capable of being delivered by the nozzle. Decreasing the throat radius would cause the chamber pressure to increase, consequently increasing the case and insulation mass required to safely operate a motor at higher pressures. To account for this factor, accurate estimates of inert mass first needed to be established. After doing so, the effects of varying nozzle expansion ratios, exit half angles, and subsequently length and mass were evaluated against motor and nozzle performance factors such as delta V and Isp. For this study, four throat radii ranging from 1.75” to 2.375” and consequently four different chamber pressures ranging 550 psia to 1200 psia were investigated. BACKGROUND To launch into Low Earth Orbit, a payload needs to be accelerated to the orbital velocity necessary to keep it from falling back to Earth. The change in velocity required between launch and orbital insertion is known as Delta V. The Delta V which a rocket or stage can deliver can be calculated using the Ideal Rocket equation, 𝛥𝑉=−𝑔0∗𝐼𝑠𝑝∗ln(𝑀𝑓𝑀𝑖) (3) Where 𝑔0 is the acceleration due to gravity at the earth’s surface, 𝐼𝑠𝑝 is the specific impulse of the rocket, 𝑀𝑖 is the initial mass of the rocket, and 𝑀𝑓 is the final mass of the rocket after burnout. From preliminary calculations beyond the scope of this paper, it was determined that 30500 ft/s of delta V would be required for a payload to be inserted into a 350 km circular orbit around the Earth. Using the known masses and Isp values of the 1st, 3rd, and 4th stages and equation 3, the delta V of each stage was calculated. The delta V required by the 2nd stage could then be found by taking the difference between the total delta V required and the delta V of the 1st, 3rd, and 4th stages. From this, the required delta V of the 2nd Stage was calculated to be 7340 ft/s. Specific impulse is an efficiency factor of the nozzle which defines the impulse delivered by the motor per unit of propellant weight. The main variables of a nozzle’s specific impulse investigated in this trade were exit cone half angle, throat radius, and expansion ratio which is affected by the throat radius. The expansion ratio, ε, of a nozzle is defined as the ratio between the nozzle exit area and throat area, and can be calculated using the equation, ε=𝑅𝑒𝑥𝑖𝑡2𝑅𝑠𝑖2 (2) Where 𝑅𝑒𝑥𝑖𝑡 is the radius of the nozzle’s exit and 𝑅𝑠𝑖 is the radius of the nozzle’s throat. A larger expansion ratio and smaller exit half angle will increase the Isp of a nozzle by allowing the gas to expand more and by allowing more of the exhaust gas to produce thrust in the direction of the motor’s central axis. A cross section view of the 2nd Stage motor with the major components annotated is provided in figure 1.

Bennett, Daniel↗
NASA NTRS1987

Astrometric Telescope Facility preliminary systems definition study. Volume 1: Executive summary

The Astrometric Telescope Facility (ATF) is a spaceborne observatory proposed for use on the Space Station (SS) as an Initial Operating Capability (IOC) payload. The primary objective of the ATF will be the search for extrasolar planetary systems and a detailed investigation of any discovered systems. In addition, it will have the capability of conducting other astrophysics investigations; e.g., measuring precise distances and motions of stars within our galaxy. The purposes of the study were to: (1) define mission and system requirements; (2) define a strawman system concept for the facility at the Prephase A level; (3) define the need for additional trade studies or technology development; and (4) estimate program cost for the strawman concept. It has been assumed for the study that the ATF will be a SS payload, will use a SS-provided Coarse Pointing System (CPS), will meet SS constraints, and will make maximum use of existing flight qualified designs or designs to be qualified by the SS program for general SS use.

Sobeck, Charlie↗
NASA NTRS2010

Life Cycle Cost Growth Study for the Discovery and New Frontiers Program Office

The D&NF Program Office LCC Management Study provides a detailed look at the drivers underlying cost overruns and schedule delays for five D&NF missions. While none of the findings are new, the study underlines the importance of continued emphasis on sound project management techniques: a clean project management structure with a clear definition of roles and responsibilities across the various partners in a project, an understanding of institutional standards and procedures and any differences among the partners, and the critical need for a comprehensive IMS that can be used easily and routinely to identify potential threats to the critical path. The study also highlights the continuing need for realistic estimates of the total LCC. Sufficient time and resources must be allocated early in a project to ensure that the appropriate trade studies and analyses are performed across all aspects of a mission: spacecraft, ground system, operations concept, and fault management, to ensure that proposed and confirmed costs truly reflect the resource requirements over the entire mission life cycle. These studies need to include a realistic review of the assumptions underlying the use of new technologies, the integration of heritage and new hardware and software into the total mission environment, and any development and test savings based on heritage technology and lessons learned. Finally, the LCC Management Study stresses the need to listen to, carefully consider, and take positive action regarding the issues raised during reviews by the expert review teams.

Barley, Bryan↗
NASA NTRS2011

An Evaluation of Ultra-High Pressure Regulator for Robotic Lunar Landing Spacecraft

The Robotic Lunar Lander Development (RLLD) Project Office at NASA Marshall Space Flight Center (MSFC) has studied several lunar surface science mission concepts. These missions focus on spacecraft carrying multiple science instruments and power systems that will allow extended operations on the lunar surface. Initial trade studies of launch vehicle options for these mission concepts indicate that the spacecraft design will be significantly mass-constrained. To minimize mass and facilitate efficient packaging, the notional propulsion system for these landers has a baseline of an ultra-high pressure (10,000 psig) helium pressurization system that has been used on Defense missiles. The qualified regulator is capable of short duration use; however, the hardware has not been previously tested at NASA spacecraft requirements with longer duration. Hence, technical risks exist in using this missile-based propulsion component for spacecraft applications. A 10,000-psig helium pressure regulator test activity is being carried out as part of risk reduction testing for MSFC RLLD project. The goal of the test activity is to assess the feasibility of commercial off-the-shelf ultra-high pressure regulator by testing with a representative flight mission profile. Slam-start, gas blowdown, water expulsion, lock-up, and leak tests are also performed on the regulator to assess performance under various operating conditions. The preliminary test results indicated that the regulator can regulate helium to a stable outlet pressure of 740 psig within the +/- 5% tolerance band and maintain a lock-up pressure less than +5% for all tests conducted. Numerous leak tests demonstrated leakage less than 10-3 standard cubic centimeters per second (SCCS) for internal seat leakage at lock-up and less than10-5 SCCS for external leakage through the regulator ambient reference cavity. The successful tests have shown the potential for 10,000 psig helium systems in NASA spacecraft and have reduced risk associated with hardware availability and hardware ability to meet RLL mission requirements.

Burnside, Christopher↗
NASA NTRS1972

Structural active cooling applications for the Space Shuttle.

Analytic and experimental studies have been conducted to evaluate a number of active cooling approaches to structural thermal protection for the Space Shuttle. The primary emphasis was directed toward the thermal protection system. Trade study results are presented for various heat shield material and TPS arrangements. Both metallic and reusable surface insulation (RSI) concepts were considered. Active systems heat sinks consisted of hydrogen, phase change materials, and expendable water. If consideration is given only to controlling the surface temperature, passive TPS was found to provide the most efficient system. Use of active cooling which incorporates some interior temperature control made the thermally less efficient RSI system more attractive.

Masek, R. V.↗
NASA NTRS2022

Lunar Communications Services with Emphasis on Commercialization

In mid-2020 Lockheed Marin Space (LM) and NASA’s Jet Propulsion Laboratory (JPL) formed a working group to study and address the need for a lunar communications network to service a growing lunar economy, beginning with the initial crewed and robotic mission needs in 2024. Multiple commercial relay architectures were studied and evaluated for platform requirements, communications capabilities, and commercial business viability in a growing ecosystem. The evolution of communications needs around the Moon were mapped to infrastructure build out. Trade studies were performed to evaluate reliability and lifetime requirements, and the viability of various data protocols. Mission concepts and operations plans were developed for both initial mission support and future autonomous network support. This work has led to Lockheed Martin and JPL to evaluate a single dedicated relay satellite in a frozen lunar orbit, with simultaneous coverage of the far-side and south pole landing sites for over 10 hours per day as a first step towards a comprehensive solution to lunar connectivity. This first satellite is compatible with an ESPA Grande launch volume and would be capable of launching on a rideshare into a variety of orbits and trajectories. The satellite is equipped with both high- and low-rate communications relay payloads with software-defined radios and a delay/disruption tolerant networking protocol. The system is designed to service initial and future mission needs in a commercialized manner, enabling a new class of missions to the Moon. The relay satellite leverages prior spacecraft platform work with on-orbit heritage in a lunar environment expected by 2023.

Davarian, Faramaz↗