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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 163 records · Page 9
All sourcesNASA NTRSDOE OSTI
NASA NTRS2013

Testing of an Annular Linear Induction Pump for the Fission Surface Power Technology Demonstration Unit

Results of performance testing of an annular linear induction pump that has been designed for integration into a fission surface power technology demonstration unit are presented. The pump electromagnetically pushes liquid metal (NaK) through a specially-designed apparatus that permits quantification of pump performance over a range of operating conditions. Testing was conducted for frequencies of 40, 55, and 70 Hz, liquid metal temperatures of 125, 325, and 525 C, and input voltages from 30 to 120 V. Pump performance spanned a range of flow rates from roughly 0.3 to 3.1 L/s (4.8 to 49 gpm), and pressure heads of <1 to 104 kPa (<0.15 to 15 psi). The maximum efficiency measured during testing was 5.4%. At the technology demonstration unit operating temperature of 525 C the pump operated over a narrower envelope, with flow rates from 0.3 to 2.75 L/s (4.8 to 43.6 gpm), developed pressure heads from <1 to 55 kPa (<0.15 to 8 psi), and a maximum efficiency of 3.5%. The pump was supplied with three-phase power at 40 and 55 Hz using a variable-frequency motor drive, while power at 55 and 70 Hz was supplied using a variable-frequency power supply. Measured performance of the pump at 55 Hz using either supply exhibited good quantitative agreement. For a given temperature, the peak in efficiency occurred at different flow rates as the frequency was changed, but the maximum value of efficiency was relative insensitive within 0.3% over the frequency range tested, including a scan from 45 to 78 Hz. The objectives of the FSP technology project are as follows:5 • Develop FSP concepts that meet expected surface power requirements at reasonable cost with added benefits over other options. • Establish a nonnuclear hardware-based technical foundation for FSP design concepts to reduce overall development risk. • Reduce the cost uncertainties for FSP and establish greater credibility for flight system cost estimates. • Generate the key nonnuclear products to allow Agency decision makers to consider FSP as a viable option for potential future flight development. The pump must be compatible with the liquid NaK coolant and have adequate performance to enable a viable flight system. Idaho National Laboratory (INL) was tasked with the design and fabrication of an ALIP suitable for the FSP reference mission. Under the program, a quarter-scale FSP technology demonstration is under construction to test the end-to-end conversion of simulated nuclear thermal power to usable electrical power intended to raise the entire FSP system to Technology Readiness Level 6. An ALIP for this TDU was fabricated under the direction of the INL and shipped to NASA Marshall Space Flight Center (MSFC) for testing at representative operating conditions. This pump was designed to meet the requirements of the TDU experiment. The ALIP test circuit (ATC) at MSFC, previously used to conduct performance evaluation on another ALIP6 was used to test the present TDU pump for the FSP Technology Development program.

Polzin, K. A.↗
NASA NTRSDate not supplied

Quantifying Medical Risk on a Long Duration Lunar Mission: A Demonstration of NASA’s IMPACT Tradespace Analysis Tool

Background NASA’s human exploration spaceflight missions to the Moon and Mars present unprecedented challenges for in-mission medical care. The distance from Earth will mean increased mission durations, communication delays, limited to no resupply opportunities, and constraints on the medical evacuation of astronauts. Mass, volume, power, and data will be limited while higher demands will be placed on the crew to manage medical care. NASA’s Moon to Mars exploration strategy lays out increasingly complex Artemis missions both in terms of duration and operations. In these more challenging deep space missions, it is important to quantitatively estimate the human medical risk to inform a traditional heuristic approach to medical risk. Prior tools have been developed for missions in low Earth orbit, but a new tool is required to plan for future exploration missions. Methods IMPACT (Informing Mission Planning via Analysis of Complex Tradespaces) is a risk assessment tool developed by NASA to advance exploration mission medical system design by quantitatively estimating mission medical risk. IMPACT v1.0 includes a novel evidence library baselined to exploration environments; an expanded list of 119 medical conditions; the addition of medical resources; and the ability for rapid and iterative analysis. Medical system risk estimates include loss of crew life, consideration of the need for return to definitive care (medical evacuation), and an estimate of crew time affected due to medical conditions. A notional long duration lunar orbit and lunar surface design reference mission (DRM) was chosen with a 4-astronaut crew to represent a sustained exploration Artemis mission. Results/Discussion Overall, IMPACT successfully quantified medical risk and derived an optimized medical system to support crew on a long duration lunar mission. In this DRM, the calculated loss of crew life from a medical event was 0.008 events per mission, risk of potential need for evacuation was 0.30 events per mission, and cumulative crew time affected by medical conditions was 103 days. The medical conditions that most contributed to overall medical risk were decompression sickness, trauma conditions, and respiratory failure. The conditions that had the largest effects on crew performance included musculoskeletal injuries and lunar dust exposure. The IMPACT-generated medical system included resources that target the most common and highest risk conditions. This systematic analysis demonstrates the value of the IMPACT tool in medical system design for human exploration spaceflight missions.

Missions to Mars↗
NASA NTRS2024

Quantifying Medical Risk on a Long Duration Lunar Mission: A Demonstration of NASA’s IMPACT Tradespace Analysis Tool

Background NASA’s human exploration spaceflight missions to the Moon and Mars present unprecedented challenges for in-mission medical care. The distance from Earth will mean increased mission durations, communication delays, limited to no resupply opportunities, and constraints on the medical evacuation of astronauts. Mass, volume, power, and data will be limited while higher demands will be placed on the crew to manage medical care. NASA’s Moon to Mars exploration strategy lays out increasingly complex Artemis missions both in terms of duration and operations. In these more challenging deep space missions, it is important to quantitatively estimate the human medical risk to inform a traditional heuristic approach to medical risk. Prior tools have been developed for missions in low Earth orbit, but a new tool is required to plan for future exploration missions. Methods IMPACT (Informing Mission Planning via Analysis of Complex Tradespaces) is a risk assessment tool developed by NASA to advance exploration mission medical system design by quantitatively estimating mission medical risk. IMPACT v1.0 includes a novel evidence library baselined to exploration environments; an expanded list of 119 medical conditions; the addition of medical resources; and the ability for rapid and iterative analysis. Medical system risk estimates include loss of crew life, consideration of the need for return to definitive care (medical evacuation), and an estimate of crew time affected due to medical conditions. A notional long duration lunar orbit and lunar surface design reference mission (DRM) was chosen with a 4-astronaut crew to represent a sustained exploration Artemis mission. Results/Discussion Overall, IMPACT successfully quantified medical risk and derived an optimized medical system to support crew on a long duration lunar mission. In this DRM, the calculated loss of crew life from a medical event was 0.008 events per mission, risk of potential need for evacuation was 0.30 events per mission, and cumulative crew time affected by medical conditions was 103 days. The medical conditions that most contributed to overall medical risk were decompression sickness, trauma conditions, and respiratory failure. The conditions that had the largest effects on crew performance included musculoskeletal injuries and lunar dust exposure. The IMPACT-generated medical system included resources that target the most common and highest risk conditions. This systematic analysis demonstrates the value of the IMPACT tool in medical system design for human exploration spaceflight missions.

Missions to Mars↗
NASA NTRS1981

Solar Brayton engine/alternator set

Work on the Mod O solar Brayton engine/alternator set is redirected to utilize solarized components of the automotive advanced gas turbine (AGT). The new configuration is referred to as the Mod I. Commercialization of solar Brayton engines thus should be enhanced not only by relating the design to an engine expected to reach the high quantity, low cost production rates associated with the automotive market, but also by the potential the AGT components provide for growth of efficiency and power rating. This growth would be achieved through use of ceramics in later versions making operation possible at temperatures up to 2500 F. The longer program duration and higher cost of the Mod I is considered.

Six, L.↗
NASA NTRS2001

Search For Effects of an Electrostatic Potential on Clocks in the Frame of Reference of a Charged Particle

The Program Planning meeting was held on Jan 27-28 at the Washington University, Dept. of Physics. Experiment strategies and details were discussed and milestones mapped. Dr. Ringermacher, the program technical leader and Co-Principal Investigator, presented a Physics Department seminar on the theory and experiment in progress. Following the meeting Professor Conradi (CO-PI) and his students designed the NMR probe and electronics required for the experiment and submitted the drawings to the shop for fabrication. Dr. Cassenti, at UTRC, initiated analytic and computer analysis of the experimental conditions in order to anticipate possible classical effects. Dr. Ringermacher, at KRONOTRAN, re-evaluated the experiment interpretation from the theory and continued working to secure its underpinnings.

Ringermacher, Harry I.↗
NASA NTRS2007

NASA's RPS Design Reference Mission Set for Solar System Exploration

NASA's 2006 Solar System Exploration (SSE) Strategic Roadmap identified a set of proposed large Flagship, medium New Frontiers and small Discovery class missions, addressing key exploration objectives. These objectives respond to the recommendations by the National Research Council (NRC), reported in the SSE Decadal Survey. The SSE Roadmap is down-selected from an over-subscribed set of missions, called the SSE Design Reference Mission (DRM) set. Missions in the Flagship and New Frontiers classes can consider Radioisotope Power Systems (RPSs), while small Discovery class missions are not permitted to use them, due to cost constraints. In line with the SSE DRM set and the SSE Roadmap missions, the RPS DRM set represents a set of missions, which can be enabled or enhanced by RPS technologies. At present, NASA has proposed the development of two new types of RPSs. These are the Multi-Mission Radioisotope Thermoelectric Generator (MMRTG), with static power conversion; and the Stirling Radioisotope Generator (SRG), with dynamic conversion. Advanced RPSs, under consideration for possible development, aim to increase specific power levels. In effect, this would either increase electric power generation for the same amount of fuel, or reduce fuel requirements for the same power output, compared to the proposed MMRTG or SRG. Operating environments could also influence the design, such that an RPS on the proposed Titan Explorer would use smaller fins to minimize heat rejection in the extreme cold environment; while the Venus Mobile Explorer long-lived in-situ mission would require the development of a new RPS, in order to tolerate the extreme hot environment, and to simultaneously provide active cooling to the payload and other electric components. This paper discusses NASA's SSE RPS DRM set, in line with the SSE DRM set. It gives a qualitative assessment regarding the impact of various RPS technology and configuration options on potential mission architectures, which could support NASA's RPS technology development planning, and provide an understanding of fuel need trades over the next three decades.

solar system↗
NASA NTRS2011

Development of an Open Rotor Cycle Model in NPSS Using a Multi-Design Point Approach

NASA's Environmentally Responsible Aviation Project and Subsonic Fixed Wing Project are focused on developing concepts and technologies which may enable dramatic reductions to the environmental impact of future generation subsonic aircraft (Refs. 1 and 2). The open rotor concept (also referred to as the Unducted Fan or advanced turboprop) may allow the achievement of this objective by reducing engine emissions and fuel consumption. To evaluate its potential impact, an open rotor cycle modeling capability is needed. This paper presents the initial development of an open rotor cycle model in the Numerical Propulsion System Simulation (NPSS) computer program which can then be used to evaluate the potential benefit of this engine. The development of this open rotor model necessitated addressing two modeling needs within NPSS. First, a method for evaluating the performance of counter-rotating propellers was needed. Therefore, a new counter-rotating propeller NPSS component was created. This component uses propeller performance maps developed from historic counter-rotating propeller experiments to determine the thrust delivered and power required. Second, several methods for modeling a counter-rotating power turbine within NPSS were explored. These techniques used several combinations of turbine components within NPSS to provide the necessary power to the propellers. Ultimately, a single turbine component with a conventional turbine map was selected. Using these modeling enhancements, an open rotor cycle model was developed in NPSS using a multi-design point approach. The multi-design point (MDP) approach improves the engine cycle analysis process by making it easier to properly size the engine to meet a variety of thrust targets throughout the flight envelope. A number of design points are considered including an aerodynamic design point, sea-level static, takeoff and top of climb. The development of this MDP model was also enabled by the selection of a simple power management scheme which schedules propeller blade angles with the freestream Mach number. Finally, sample open rotor performance results and areas for further model improvements are presented.

Hendricks, Eric S.↗
NASA NTRS2011

Benefits of Power and Propulsion Technology for a Piloted Electric Vehicle to an Asteroid

NASA's goal for human spaceflight is to expand permanent human presence beyond low Earth orbit (LEO). NASA is identifying potential missions and technologies needed to achieve this goal. Mission options include crewed destinations to LEO and the International Space Station; high Earth orbit and geosynchronous orbit; cis-lunar space, lunar orbit, and the surface of the Moon; near-Earth objects; and the moons of Mars, Mars orbit, and the surface of Mars. NASA generated a series of design reference missions to drive out required functions and capabilities for these destinations, focusing first on a piloted mission to a near-Earth asteroid. One conclusion from this exercise was that a solar electric propulsion stage could reduce mission cost by reducing the required number of heavy lift launches and could increase mission reliability by providing a robust architecture for the long-duration crewed mission. Similarly, solar electric vehicles were identified as critical for missions to Mars, including orbiting Mars, landing on its surface, and visiting its moons. This paper describes the parameterized assessment of power and propulsion technologies for a piloted solar electric vehicle to a near-Earth asteroid. The objective of the assessment was to determine technology drivers to advance the state of the art of electric propulsion systems for human exploration. Sensitivity analyses on the performance characteristics of the propulsion and power systems were done to determine potential system-level impacts of improved technology. Starting with a "reasonable vehicle configuration" bounded by an assumed launch date, we introduced technology improvements to determine the system-level benefits (if any) that those technologies might provide. The results of this assessment are discussed and recommendations for future work are described.

Mercer, Carolyn R.↗
NASA NTRS2012

Benefits of Power and Propulsion Technology for a Piloted Electric Vehicle to an Asteroid

NASA s goal for human spaceflight is to expand permanent human presence beyond low Earth orbit (LEO). NASA is identifying potential missions and technologies needed to achieve this goal. Mission options include crewed destinations to LEO and the International Space Station; high Earth orbit and geosynchronous orbit; cis-lunar space, lunar orbit, and the surface of the Moon; near-Earth objects; and the moons of Mars, Mars orbit, and the surface of Mars. NASA generated a series of design reference missions to drive out required functions and capabilities for these destinations, focusing first on a piloted mission to a near-Earth asteroid. One conclusion from this exercise was that a solar electric propulsion stage could reduce mission cost by reducing the required number of heavy lift launches and could increase mission reliability by providing a robust architecture for the long-duration crewed mission. Similarly, solar electric vehicles were identified as critical for missions to Mars, including orbiting Mars, landing on its surface, and visiting its moons. This paper describes the parameterized assessment of power and propulsion technologies for a piloted solar electric vehicle to a near-Earth asteroid. The objective of the assessment was to determine technology drivers to advance the state of the art of electric propulsion systems for human exploration. Sensitivity analyses on the performance characteristics of the propulsion and power systems were done to determine potential system-level impacts of improved technology. Starting with a "reasonable vehicle configuration" bounded by an assumed launch date, we introduced technology improvements to determine the system-level benefits (if any) that those technologies might provide. The results of this assessment are discussed and recommendations for future work are described.

Mercer, Carolyn R.↗
NASA NTRS2004

Microthrust Propulsion of the LISA Mission

We present the most recent propulsion requirements for the Laser Interferometer Space Antenna (LISA) Mission and describe potential microth ruster technology that can meet these requirements. LISA consists of three spacecraft in heliocentric orbits, forming a triangle with 5x l 0 (exp 6) km sides that are the arms of three Michelson-type interferometers. Reflective proof masses provide the reference surfaces at the end of the interferometer arms as part of the Gravitational Referenc e Sensor (GRS) designed to detect gravitational waves. The microthrus t propulsion system will be part of the Disturbance Reduction System (DRS), which is responsible for maintaining each spacecraft position w ithin approximately 10 nm around the proof masses. To provide the nec essary sensitivity, the GRS must not experience spurious acceleration s >15 (exp -10) m/ s(exp 2) in the 0.1 mHz to 1 Hz bandwidth, requiring precision formation flying and drag-free operation of the LISA spa cecraft. This leads to the following microthruster performance requir ements: a thrust range of 2-30 Micro N, a thrust resolution < 0.1 Mic ro N, and thrust noise <0.1 Hz(exp -1/2) over the LISA measurement bandwidth. The microthruster must provide this performance for 5 years c ontinuously, contain 10 years worth of propellant, and not disrupt th e science measurements. Potential microthruster technologies include Colloid, Field Emission Electric Propulsion (FEEP), and precision cold gas microthrusters. Each of these technologies is described in detai l with focus on the NASA microthruster development of the Busek Collo id Micro-Newton Thruster (CMNT).

Field Emission Electric Propulsion (FEEP)↗
NASA NTRS1982

Radiometric considerations in remote sensing systems

All sensors systems designed to acquire quantitative data undergo radiometric calibration. The types and potential accuracies of calibration are discussed as well as the needs for calibration in the practical application of sensors. The recent and ongoing experience with the shuttle multispectral infrared radiometer is used as a reference.

Goetz, A. F. H.↗
NASA NTRS1988

Problems in fluid dynamics

A scheme was developed for the parametric differentiation and integration of gas dynamics equations. A numerical integration of the gas dynamics equations is necessarily performed for a specific set of parameter values. The linear variational equations are obtained by differentiating the exact equations with respect to each of the relevant parameters. The resulting matrix of flow quantities is referred to as the Jacobi matrix. The subsequent procedure is then straightforward. The method was tested for two dimensional supersonic flow past an airfoil, with airfoil thickness, camber, and angle of attack varied. This approach has great potential value for rapidly assessing the effect of design changes. The other focus of the work was on problems in fluid stability, bifurcations, and turbulence.

Sirovich, Lawrence↗
NASA NTRS1992

Advanced aeroservoelastic stabilization techniques for hypersonic flight vehicles

Advanced high performance vehicles, including Single-Stage-To-Orbit (SSTO) hypersonic flight vehicles, that are statically unstable, require higher bandwidth flight control systems to compensate for the instability resulting in interactions between the flight control system, the engine/propulsion dynamics, and the low frequency structural modes. Military specifications, such as MIL-F-9490D and MIL-F-87242, tend to limit treatment of structural modes to conventional gain stabilization techniques. The conventional gain stabilization techniques, however, introduce low frequency effective time delays which can be troublesome from a flying qualities standpoint. These time delays can be alleviated by appropriate blending of gain and phase stabilization techniques (referred to as Hybrid Phase Stabilization or HPS) for the low frequency structural modes. The potential of using HPS for compensating structural mode interaction was previously explored. It was shown that effective time delay was significantly reduced with the use of HPS; however, the HPS design was seen to have greater residual response than a conventional gain stablized design. Additional work performed to advance and refine the HPS design procedure, to further develop residual response metrics as a basis for alternative structural stability specifications, and to develop strategies for validating HPS design and specification concepts in manned simulation is presented. Stabilization design sensitivity to structural uncertainties and aircraft-centered requirements are also assessed.

Chan, Samuel Y.↗
NASA NTRS1992

Lunar base thermal management/power system analysis and design

A compilation of several lunar surface thermal management and power system studies completed under contract and IR&D is presented. The work includes analysis and preliminary design of all major components of an integrated thermal management system, including loads determination, active internal acquisition and transport equipment, external transport systems (active and passive), passive insulation, solar shielding, and a range of lunar surface radiator concepts. Several computer codes were utilized in support of this study, including RADSIM to calculate radiation exchange factors and view factors, RADIATOR (developed in-house) for heat rejection system sizing and performance analysis over a lunar day, SURPWER for power system sizing, and CRYSTORE for cryogenic system performance predictions. Although much of the work was performed in support of lunar rover studies, any or all of the results can be applied to a range of surface applications. Output data include thermal loads summaries, subsystem performance data, mass, and volume estimates (where applicable), integrated and worst-case lunar day radiator size/mass and effective sink temperatures for several concepts (shielded and unshielded), and external transport system performance estimates for both single and two-phase (heat pumped) transport loops. Several advanced radiator concepts are presented, along with brief assessments of possible system benefits and potential drawbacks. System point designs are presented for several cases, executed in support of the contract and IR&D studies, although the parametric nature of the analysis is stressed to illustrate applicability of the analysis procedure to a wide variety of lunar surface systems. The reference configuration(s) derived from the various studies will be presented along with supporting criteria. A preliminary design will also be presented for the reference basing scenario, including qualitative data regarding TPS concerns and issues.

Mcghee, Jerry R.↗
NASA NTRS1993

NASCAP programmer's reference manual

The NASA Charging Analyzer Program (NASCAP) is a computer program designed to model the electrostatic charging of complicated three-dimensional objects, both in a test tank and at geosynchronous altitudes. This document is a programmer's reference manual and user's guide. It is designed as a reference to experienced users of the code, as well as an introduction to its use for beginners. All of the many capabilities of NASCAP are covered in detail, together with examples of their use. These include the definition of objects, plasma environments, potential calculations, particle emission and detection simulations, and charging analysis.

Mandell, M. J.↗
NASA NTRS2000

Steady-State Computation of Constant Rotational Rate Dynamic Stability Derivatives

Dynamic stability derivatives are essential to predicting the open and closed loop performance, stability, and controllability of aircraft. Computational determination of constant-rate dynamic stability derivatives (derivatives of aircraft forces and moments with respect to constant rotational rates) is currently performed indirectly with finite differencing of multiple time-accurate computational fluid dynamics solutions. Typical time-accurate solutions require excessive amounts of computational time to complete. Formulating Navier-Stokes (N-S) equations in a rotating noninertial reference frame and applying an automatic differentiation tool to the modified code has the potential for directly computing these derivatives with a single, much faster steady-state calculation. The ability to rapidly determine static and dynamic stability derivatives by computational methods can benefit multidisciplinary design methodologies and reduce dependency on wind tunnel measurements. The CFL3D thin-layer N-S computational fluid dynamics code was modified for this study to allow calculations on complex three-dimensional configurations with constant rotation rate components in all three axes. These CFL3D modifications also have direct application to rotorcraft and turbomachinery analyses. The modified CFL3D steady-state calculation is a new capability that showed excellent agreement with results calculated by a similar formulation. The application of automatic differentiation to CFL3D allows the static stability and body-axis rate derivatives to be calculated quickly and exactly.

Park, Michael A.↗
NASA NTRS2016

Nuclear Systems Kilopower Overview

The Nuclear Systems Kilopower Project was initiated by NASAs Space Technology Mission Directorate Game Changing Development Program in fiscal year 2015 to demonstrate subsystem-level technology readiness of small space fission power in a relevant environment (Technology Readiness Level 5) for space science and human exploration power needs. The Nuclear Systems Kilopower Project consists of two elements. The primary element is the Kilopower Prototype Test, also called the Kilopower Reactor Using Stirling Technology(KRUSTY) Test. This element consists of the development and testing of a fission ground technology demonstrator of a 1 kWe fission power system. A 1 kWe system matches requirements for some robotic precursor exploration systems and future potential deep space science missions, and also allows a nuclear ground technology demonstration in existing nuclear test facilities at low cost. The second element, the Mars Kilopower Scalability Study, consists of the analysis and design of a scaled-up version of the 1 kWe reference concept to 10 kWe for Mars surface power projected requirements, and validation of the applicability of the KRUSTY experiment to key technology challenges for a 10 kWe system. If successful, these two elements will lead to initiation of planning for a technology demonstration of a 10 kWe fission power capability for Mars surface outpost power.

Space Nuclear Power↗
NASA NTRS1983

Fluorine disposal

A preliminary design of an F2 dispoal system for HELSTF is presented along with recommendations on operational policy and identification of potential operational problems. The analysis is based on sizing a system to handle two different modes of the HELSTF Fluorine Flow System (one operational and one catastrophic). This information should serve both as a guide to a final detailed design for HELSTF as well as a reference for subsequent monitoring and/or modification of the system which consists of a charcoal reactor followed by a dry soda lime scrubber.

Rakow, A.↗