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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 703 records · Page 39
All sourcesNASA NTRSDOE OSTI
NASA NTRS2014

Adapting a Large-Scale Multi-Mission Ground System for Low-Cost CubeSats

The majority of today's CubeSat fleet consists of Earth-orbiting missions that mostly use existing ground systems developed by universities because of availability, simplicity, and low-cost. The Interplanetary NanoSpacecraft Pathfinder In Relevant Environment (INSPIRE) mission is a revolutionary CubeSat mission that will launch a pair of CubeSats into deep space to study the feasibility of CubeSats beyond low-Earth orbit. This uncovers a new set of systems and software engineering challenges to the development of a robust and reliable ground system in a low-cost environment. In this paper, we discuss the approach to these challenges by using the Jet Propulsion Laboratory's (JPL) Advanced Multimission Operation System (AMMOS) Ground Data System (GDS) as well as the methodologies used to engineer the flight system to work with an existing ground system developed for large-scale missions. Specifically we will focus on the command and telemetry subsystem of AMMOS, the Multimission Data Processing and Control System. We conclude with a retrospective on the challenges encountered and a brief discussion on our efforts to provide AMMOS to support future deep space CubeSat missions.

Quach, William L.↗
NASA NTRS2007

Nuclear Thermal Propulsion Mars Mission Systems Analysis and Requirements Definition

This paper describes the Mars transportation vehicle design concepts developed by the Marshall Space Flight Center (MSFC) Advanced Concepts Office. These vehicle design concepts provide an indication of the most demanding and least demanding potential requirements for nuclear thermal propulsion systems for human Mars exploration missions from years 2025 to 2035. Vehicle concept options vary from large "all-up" vehicle configurations that would transport all of the elements for a Mars mission on one vehicle. to "split" mission vehicle configurations that would consist of separate smaller vehicles that would transport cargo elements and human crew elements to Mars separately. Parametric trades and sensitivity studies show NTP stage and engine design options that provide the best balanced set of metrics based on safety, reliability, performance, cost and mission objectives. Trade studies include the sensitivity of vehicle performance to nuclear engine characteristics such as thrust, specific impulse and nuclear reactor type. Tbe associated system requirements are aligned with the NASA Exploration Systems Mission Directorate (ESMD) Reference Mars mission as described in the Explorations Systems Architecture Study (ESAS) report. The focused trade studies include a detailed analysis of nuclear engine radiation shield requirements for human missions and analysis of nuclear thermal engine design options for the ESAS reference mission.

Mulqueen, Jack↗
NASA NTRS1993

Orbit transfer rocket engine technology program

An advanced near term (1990's) space-based Orbit Transfer Vehicle Engine (OTVE) system was designed, and the technologies applicable to its construction, maintenance, and operations were developed under Tasks A through F of the Orbit Transfer Rocket Engine Technology Program. Task A was a reporting task. In Task B, promising OTV turbomachinery technologies were explored: two stage partial admission turbines, high velocity ratio diffusing crossovers, soft wear ring seals, advanced bearing concepts, and a rotordynamic analysis. In Task C, a ribbed combustor design was developed. Possible rib and channel geometries were chosen analytically. Rib candidates were hot air tested and laser velocimeter boundary layer analyses were conducted. A channel geometry was also chosen on the basis of laser velocimeter data. To verify the predicted heat enhancement effects, a ribbed calorimeter spool was hot fire tested. Under Task D, the optimum expander cycle engine thrust, performance and envelope were established for a set of OTV missions. Optimal nozzle contours and quick disconnects for modularity were developed. Failure Modes and Effects Analyses, maintenance and reliability studies and component study results were incorporated into the engine system. Parametric trades on engine thrust, mixture ratio, and area ratio were also generated. A control system and the health monitoring and maintenance operations necessary for a space-based engine were outlined in Task E. In addition, combustor wall thickness measuring devices and a fiberoptic shaft monitor were developed. These monitoring devices were incorporated into preflight engine readiness checkout procedures. In Task F, the Integrated Component Evaluator (I.C.E.) was used to demonstrate performance and operational characteristics of an advanced expander cycle engine system and its component technologies. Sub-system checkouts and a system blowdown were performed. Short transitions were then made into main combustor ignition and main stage operation.

Gustafson, N. B.↗
NASA NTRS1991

An ultrareliable integrated digital computer for helicopters

The US Army Fault-Tolerant Architecture (AFTA) has been defined to meet real-time fault-tolerant processing requirements of future Army applications. AFTA is the enabling technology that will allow the Army to configure existing processors and other hardware to provide high throughput and ultrahigh reliability necessary for terrain-following/terrain-avoidance (TF/TA) and nap-of-the-earth (NOE) flight control and other advanced Army applications. A brief description of AFTA's relevant attributes and theory of operation is given. The systems engineering models which permit the configuration on an AFTA to meet mission requirements as they evolve are presented, along with an example of their use to engineer an AFTA configuration. The resulting AFTA configuration, its predicted throughput, reliability, availability, weight, power, volume, and life-cycle cost impact, and the effects of possible engineering tradeoffs are presented.

Harper, Richard E.↗
NASA NTRS1992

The Saturn V F-1 engine revisited

This paper describes how the Saturn V F-1 engine could be resurrected for potential use in a new heavy lift launch vehicle (HLLV). Performance, history, and reliability of the F-1 will be discussed and a 1990s F-1 engine configuration presented. The status of residual hardware, production capability, test-stand availability, and a general program plan to accomplish this restart will be presented. Similarities between restarting the F-1 program and restarting the Atlas and Delta engine programs will be discussed.

Shelton, B. W.↗
NASA NTRS1993

SPDE/SPRE final summary report

Mechanical Technology Incorporated (MTI) performed acceptance testing on the Space Power Research Engine (SPRE), which demonstrated satisfactory operation and sufficient reliability for delivery to NASA Lewis Research Center. The unit produced 13.5 kW PV power with an efficiency of 22 percent versus design goals of 28.8 kW PV power and efficiency of 28 percent. Maximum electric power was only 8 kWe due to lower alternator efficiency. One of the major shortcomings of the SPRE was linear alternator efficiency, which was only 70 percent compared to a design value of 90 percent. It was determined from static tests that the major cause for the efficiency shortfall was the location of the magnetic structure surrounding the linear alternator. Testing of an alternator configuration without a surrounding magnetic structure on a linear dynamometer confirmed earlier static test results. Linear alternator efficiency improved from 70 percent to over 90 percent. Testing of the MTI SPRE was also performed with hydrodynamic bearings and achieved full-stroke, stable operation. This testing indicated that hydrodynamic bearings may be useful in free piston Stirling engines. An important factor in achieving stable operation at design stroke was isolating a portion of the bearing length from the engine pressure variations. In addition, the heat pipe heater head design indicates that integration of a Stirling engine with a heat source can be performed via heat pipes. This design provides a baseline against which alternative designs can be measured.

Dochat, George↗
NASA NTRS2005

The Systems Engineering Process for Human Support Technology Development

Systems engineering is designing and optimizing systems. This paper reviews the systems engineering process and indicates how it can be applied in the development of advanced human support systems. Systems engineering develops the performance requirements, subsystem specifications, and detailed designs needed to construct a desired system. Systems design is difficult, requiring both art and science and balancing human and technical considerations. The essential systems engineering activity is trading off and compromising between competing objectives such as performance and cost, schedule and risk. Systems engineering is not a complete independent process. It usually supports a system development project. This review emphasizes the NASA project management process as described in NASA Procedural Requirement (NPR) 7120.5B. The process is a top down phased approach that includes the most fundamental activities of systems engineering - requirements definition, systems analysis, and design. NPR 7120.5B also requires projects to perform the engineering analyses needed to ensure that the system will operate correctly with regard to reliability, safety, risk, cost, and human factors. We review the system development project process, the standard systems engineering design methodology, and some of the specialized systems analysis techniques. We will discuss how they could apply to advanced human support systems development. The purpose of advanced systems development is not directly to supply human space flight hardware, but rather to provide superior candidate systems that will be selected for implementation by future missions. The most direct application of systems engineering is in guiding the development of prototype and flight experiment hardware. However, anticipatory systems engineering of possible future flight systems would be useful in identifying the most promising development projects.

Jones, Harry↗
NASA NTRS2011

Product Lifecycle Management and the Quest for Sustainable Space Exploration Solutions

Product Lifecycle Management (PLM) is an outcome of lean thinking to eliminate waste and increase productivity. PLM is inextricably tied to the systems engineering business philosophy, coupled with a methodology by which personnel, processes and practices, and information technology combine to form an architecture platform for product design, development, manufacturing, operations, and decommissioning. In this model, which is being implemented by the Marshall Space Flight Center (MSFC) Engineering Directorate, total lifecycle costs are important variables for critical decision-making. With the ultimate goal to deliver quality products that meet or exceed requirements on time and within budget, PLM is a powerful concept to shape everything from engineering trade studies and testing goals, to integrated vehicle operations and retirement scenarios. This briefing will demonstrate how the MSFC Engineering Directorate is implementing PLM as part of an overall strategy to deliver safe, reliable, and affordable space exploration solutions and how that strategy aligns with the Agency and Center systems engineering policies and processes. Sustainable space exploration solutions demand that all lifecycle phases be optimized, and engineering the next generation space transportation system requires a paradigm shift such that digital tools and knowledge management, which are central elements of PLM, are used consistently to maximum effect. Adopting PLM, which has been used by the aerospace and automotive industry for many years, for spacecraft applications provides a foundation for strong, disciplined systems engineering and accountable return on investment. PLM enables better solutions using fewer resources by making lifecycle considerations in an integrative decision-making process.

Caruso, Pamela W.↗
NASA NTRS2022

Reliability-Aware Requirements Development for Autonomy Software

Autonomy is the ability of a system to make decisions and take actions in the presence of uncertainty. Validating autonomy, therefore, is a matter of validating that the system makes intelligent decisions even when there might be discrepancies between the observed and the expected state of the world as well as when the possible outcome of each of the possible courses of action is uncertain. During the requirements engineering phase, the functions, including those that are required to be autonomous and their expected behavior are specified. However, in the case of autonomous software system, it is challenging to specify all possible scenarios that the software may encounter. Additionally, since autonomous behaviors cannot always be fully predetermined, it is difficult to reason about their completeness and correctness. While these challenges are also faced with non-autonomous system, they are more prevalent for autonomous system due to their complexity and emergent behavior. Our efforts have been three-fold. On one hand, we have developed a goal-based method for requirements decomposition. We have embedded this approach into a higher-level systems engineering framework and then developed methods for decision and reliability analysis within this framework. And finally, we use simulation analysis to validate the more theoretical methods and provide a feedback mechanism. The decision analysis approach (DA) helps in identifying and mitigating the run-time risks, by bringing to the forefront the uncertainties, decisions, interactions, and other factors that may cause autonomous software to make erroneous decisions. The DA framework is used to quantify, for each decision that the software can make, failure risk due to uncertainties. The information about the key decisions and circumstances that can cause autonomous system to make incorrect decisions are used to generate test scenarios to be run in a simulator to ensure that the system can handle error-prone circumstances. The results or outcomes from decisions, based on the simulated scenarios, can then be leveraged to further gain insight of the assured system. We combine a goal-based approach with simulation analysis to facilitate requirements development for autonomous systems and further provide a method for run-time systems level reliability considerations during the autonomous software development process. The DA and the reliability analysis methods provide a formalism for the consideration of uncertainty in the run time decision-making process for autonomous software. This formalism considers the uncertainty involved in the outcome of different courses of action, in terms of performance and cost, as well as a method to consider the system level reliability implications. System level reliability considerations for autonomous software decision making contribute to the satisfaction of the system level goals over the lifecycle of a given mission. We demonstrate our approach using a hypothetical rover path planning example. The requirements for our rover simulator are based on existing literature about Mars rovers and lessons learned from real world observations at JPL. The specifics of the autonomy design for the Mars rovers, that have not been previously cleared for external release, are not used for our demonstrations.

Lindvall, Mikael↗
NASA NTRS2014

Test Driven Development of Scientific Models

Test-Driven Development (TDD), a software development process that promises many advantages for developer productivity and software reliability, has become widely accepted among professional software engineers. As the name suggests, TDD practitioners alternate between writing short automated tests and producing code that passes those tests. Although this overly simplified description will undoubtedly sound prohibitively burdensome to many uninitiated developers, the advent of powerful unit-testing frameworks greatly reduces the effort required to produce and routinely execute suites of tests. By testimony, many developers find TDD to be addicting after only a few days of exposure, and find it unthinkable to return to previous practices.After a brief overview of the TDD process and my experience in applying the methodology for development activities at Goddard, I will delve more deeply into some of the challenges that are posed by numerical and scientific software as well as tools and implementation approaches that should address those challenges.

Overview of TDD Process↗
NASA NTRS2010

Flight Avionics Sequencing Telemetry (FAST) DIV Latching Display

The NASA Engineering (NE) Directorate at Kennedy Space Center provides engineering services to major programs such as: Space Shuttle, Inter national Space Station, and the Launch Services Program (LSP). The Av ionics Division within NE, provides avionics and flight control syste ms engineering support to LSP. The Launch Services Program is respons ible for procuring safe and reliable services for transporting critical, one of a kind, NASA payloads into orbit. As a result, engineers mu st monitor critical flight events during countdown and launch to asse ss anomalous behavior or any unexpected occurrence. The goal of this project is to take a tailored Systems Engineering approach to design, develop, and test Iris telemetry displays. The Flight Avionics Sequen cing Telemetry Delta-IV (FAST-D4) displays will provide NASA with an improved flight event monitoring tool to evaluate launch vehicle heal th and performance during system-level ground testing and flight. Flight events monitored will include data from the Redundant Inertial Fli ght Control Assembly (RIFCA) flight computer and launch vehicle comma nd feedback data. When a flight event occurs, the flight event is ill uminated on the display. This will enable NASA Engineers to monitor c ritical flight events on the day of launch. Completion of this project requires rudimentary knowledge of launch vehicle Guidance, Navigatio n, and Control (GN&C) systems, telemetry, and console operation. Work locations for the project include the engineering office, NASA telem etry laboratory, and Delta launch sites.

Moore, Charlotte↗
NASA NTRS2011

Heavy Lift Launch Capability with a New Hydrocarbon Engine

The Advanced Concepts Office at NASA's George C. Marshall Space Flight Center was tasked to define the thrust requirement of a new liquid oxygen rich staged combustion cycle hydrocarbon engine that could be utilized in a launch vehicle to meet NASA s future heavy lift needs. Launch vehicle concepts were sized using this engine for different heavy lift payload classes. Engine out capabilities for one of the heavy lift configurations were also analyzed for increased reliability that may be desired for high value payloads or crewed missions. The applicability for this engine in vehicle concepts to meet military and commercial class payloads comparable to current ELV capability was also evaluated.

Threet, Grady E., Jr.↗
NASA NTRS2010

Smart Sensor Demonstration Payload

Sensors are a critical element to any monitoring, control, and evaluation processes such as those needed to support ground based testing for rocket engine test. Sensor applications involve tens to thousands of sensors; their reliable performance is critical to achieving overall system goals. Many figures of merit are used to describe and evaluate sensor characteristics; for example, sensitivity and linearity. In addition, sensor selection must satisfy many trade-offs among system engineering (SE) requirements to best integrate sensors into complex systems [1]. These SE trades include the familiar constraints of power, signal conditioning, cabling, reliability, and mass, and now include considerations such as spectrum allocation and interference for wireless sensors. Our group at NASA s John C. Stennis Space Center (SSC) works in the broad area of integrated systems health management (ISHM). Core ISHM technologies include smart and intelligent sensors, anomaly detection, root cause analysis, prognosis, and interfaces to operators and other system elements [2]. Sensor technologies are the base fabric that feed data and health information to higher layers. Cost-effective operation of the complement of test stands benefits from technologies and methodologies that contribute to reductions in labor costs, improvements in efficiency, reductions in turn-around times, improved reliability, and other measures. ISHM is an active area of development at SSC because it offers the potential to achieve many of those operational goals [3-5].

Schmalzel, John↗
NASA NTRS1991

Advances in knowledge-based software engineering

The underlying hypothesis of this work is that a rigorous and comprehensive software reuse methodology can bring about a more effective and efficient utilization of constrained resources in the development of large-scale software systems by both government and industry. It is also believed that correct use of this type of software engineering methodology can significantly contribute to the higher levels of reliability that will be required of future operational systems. An overview and discussion of current research in the development and application of two systems that support a rigorous reuse paradigm are presented: the Knowledge-Based Software Engineering Environment (KBSEE) and the Knowledge Acquisition fo the Preservation of Tradeoffs and Underlying Rationales (KAPTUR) systems. Emphasis is on a presentation of operational scenarios which highlight the major functional capabilities of the two systems.

Truszkowski, Walt↗
NASA NTRS1983

Very Low Thrust Gaseous Oxygen-hydrogen Rocket Engine Ignition Technology

An experimental program was performed to determine the minimum energy per spark for reliable and repeatable ignition of gaseous oxygen (GO2) and gaseous hydrogen (GH2) in very low thrust 0.44 to 2.22-N (0.10 to 0.50-lb sub f) rocket engines or spacecraft and satellite attitude control systems (ACS) application. Initially, the testing was conducted at ambient conditions, with the results subsequently verified under vacuum conditions. An experimental breadboard electrical exciter that delivered 0.2 to 0.3 mj per spark was developed and demonstrated by repeated ignitions of a 2.22-N (0.50-lb sub f) thruster in a vacuum chamber with test durations up to 30 min.

Bjorklund, Roy A.↗
NASA NTRS1975

Program for refan JT8D engine design, fabrication and test, phase 2

The objective of the JT8D refan program was to design, fabricate, and test certifiable modifications of the JT8D engine which would reduce noise generated by JT8D powered aircraft. This was to be accomplished without affecting reliability and maintainability, at minimum retrofit cost, and with no performance penalty. The mechanical design, engine performance and stability characteristics at sea-level and altitude, and the engine noise characteristics of the test engines are documented. Results confirmed the structural integrity of the JT8D-109. Engine operation was stable throughout the airplane flight envelope. Fuel consumption of the test engines was higher than that required to meet the goal of no airplane performance penalty, but the causes were identified and corrected during a normal pre-certification engine development program. Compared to the baseline JT8D-109 engine, the acoustically treated JT8D-109 engine showed noise reductions of 6 PNdB at takeoff and 11 PNdB at a typical approach power setting.

Glass, J. A.↗
NASA NTRS2018

Chapter 12: Materials for Liquid Propulsion Systems

Earth to orbit launch vehicles are propelled by rocket engines and motors, both liquid and solid. This chapter will discuss liquid engines. The heart of a launch vehicle is its engine. The remainder of the vehicle (with the notable exceptions of the payload and guidance system) is an aero structure to support the propellant tanks which provide the fuel and oxidizer to feed the engine or engines. The basic principle behind a rocket engine is straightforward. The engine is a means to convert potential thermochemical energy of one or more propellants into exhaust jet kinetic energy. Fuel and oxidizer are burned in a combustion chamber where they create hot gases under high pressure. These hot gases are allowed to expand through a nozzle. The molecules of hot gas are first constricted by the throat of the nozzle (de-Laval nozzle) which forces them to accelerate; then as the nozzle flares outwards, they expand and further accelerate. It is the mass of the combustion gases times their velocity, reacting against the walls of the combustion chamber and nozzle, which produce thrust according to Newton's third law: for every action there is an equal and opposite reaction. Solid rocket motors are cheaper to manufacture and offer good values for their cost. Liquid propellant engines offer higher performance, that is, they deliver greater thrust per unit weight of propellant burned. They also have a considerably higher thrust to weigh ratio. Since liquid rocket engines can be tested several times before flight, they have the capability to be more reliable, and their ability to shut down once started provides an extra margin of safety. Liquid propellant engines also can be designed with restart capability to provide orbital maneuvering capability. In some instances, liquid engines also can be designed to be reusable. On the solid side, hybrid solid motors also have been developed with the capability to stop and restart. Solid motors are covered in detail in chapter 11. Liquid rocket engine operational factors can be described in terms of extremes: temperatures ranging from that of liquid hydrogen (-423 F) to 6000 F hot gases; enormous thermal shock (7000 F/sec); large temperature differentials between contiguous components; reactive propellants; extreme acoustic environments; high rotational speeds for turbo machinery and extreme power densities. These factors place great demands on materials selection and each must be dealt with while maintaining an engine of the lightest possible weight. This chapter will describe the design considerations for the materials used in the various components of liquid rocket engines and provide examples of usage and experiences in each.

Halchak, John A.↗