Results for “electrical” — Page 16

NASA NTRS2014

High-Power Solar Electric Propulsion for Future NASA Missions

NASA has sought to utilize high-power solar electric propulsion as means of improving the affordability of in-space transportation for almost 50 years. Early efforts focused on 25 to 50 kilowatt systems that could be used with the Space Shuttle, while later efforts focused on systems nearly an order of magnitude higher power that could be used with heavy lift launch vehicles. These efforts never left the concept development phase in part because the technology required was not sufficiently mature. Since 2012 the NASA Space Technology Mission Directorate has had a coordinated plan to mature the requisite solar array and electric propulsion technology needed to implement a 30 to 50 kilowatt solar electric propulsion technology demonstration mission. Multiple solar electric propulsion technology demonstration mission concepts have been developed based on these maturing technologies with recent efforts focusing on an Asteroid Redirect Robotic Mission. If implemented, the Asteroid Redirect Vehicle will form the basis for a capability that can be cost-effectively evolved over time to provide solar electric propulsion transportation for a range of follow-on mission applications at power levels in excess of 100 kilowatts.

Mission Applications↗

NASA NTRS2013

Guide to Flow Measurement for Electric Propulsion Systems

In electric propulsion (EP) systems, accurate measurement of the propellant mass flow rate of gas or liquid to the thruster and external cathode is a key input in the calculation of thruster efficiency and specific impulse. Although such measurements are often achieved with commercial mass flow controllers and meters integrated into propellant feed systems, the variability in potential propellant options and flow requirements amongst the spectrum of EP power regimes and devices complicates meter selection, integration, and operation. At the direction of the Committee on Standards for Electric Propulsion Testing, a guide was jointly developed by members of the electric propulsion community to establish a unified document that contains the working principles, methods of implementation and analysis, and calibration techniques and recommendations on the use of mass flow meters in laboratory and spacecraft electric propulsion systems. The guide is applicable to EP devices of all types and power levels ranging from microthrusters to high-power ion engines and Hall effect thrusters. The establishment of a community standard on mass flow metering will help ensure the selection of the proper meter for each application. It will also improve the quality of system performance estimates by providing comprehensive information on the physical phenomena and systematic errors that must be accounted for during the analysis of flow measurement data. This paper will outline the standard methods and recommended practices described in the guide titled "Flow Measurement for Electric Propulsion Systems."

spacecraft electric propulsion (EP) systems↗

NASA NTRS2015

Xenon Acquisition Strategies for High-Power Electric Propulsion NASA Missions

Solar electric propulsion (SEP) has been used for station-keeping of geostationary communications satellites since the 1980s. Solar electric propulsion has also benefitted from success on NASA Science Missions such as Deep Space One and Dawn. The xenon propellant loads for these applications have been in the 100s of kilograms range. Recent studies performed for NASA's Human Exploration and Operations Mission Directorate (HEOMD) have demonstrated that SEP is critically enabling for both near-term and future exploration architectures. The high payoff for both human and science exploration missions and technology investment from NASA's Space Technology Mission Directorate (STMD) are providing the necessary convergence and impetus for a 30-kilowatt-class SEP mission. Multiple 30-50- kilowatt Solar Electric Propulsion Technology Demonstration Mission (SEP TDM) concepts have been developed based on the maturing electric propulsion and solar array technologies by STMD with recent efforts focusing on an Asteroid Redirect Robotic Mission (ARRM). Xenon is the optimal propellant for the existing state-of-the-art electric propulsion systems considering efficiency, storability, and contamination potential. NASA mission concepts developed and those proposed by contracted efforts for the 30-kilowatt-class demonstration have a range of xenon propellant loads from 100s of kilograms up to 10,000 kilograms. This paper examines the status of the xenon industry worldwide, including historical xenon supply and pricing. The paper will provide updated information on the xenon market relative to previous papers that discussed xenon production relative to NASA mission needs. The paper will discuss the various approaches for acquiring on the order of 10 metric tons of xenon propellant to support potential near-term NASA missions. Finally, the paper will discuss acquisitions strategies for larger NASA missions requiring 100s of metric tons of xenon will be discussed.

Ion Propulsion↗

NASA NTRS2016

Temperature Dependence of Electrical Resistance of Woven Melt-Infiltrated SiCf/SiC Ceramic Matrix Composites

Recent studies have successfully shown the use of electrical resistance (ER)measurements to monitor room temperature damage accumulation in SiC fiber reinforced SiC matrix composites (SiCf/SiC) Ceramic Matrix Composites (CMCs). In order to determine the feasibility of resistance monitoring at elevated temperatures, the present work investigates the temperature dependent electrical response of various MI (Melt Infiltrated)-CVI (Chemical Vapor Infiltrated) SiC/SiC composites containing Hi-Nicalon Type S, Tyranno ZMI and SA reinforcing fibers. Test were conducted using a commercially available isothermal testing apparatus as well as a novel, laser-based heating approach developed to more accurately simulate thermomechanical testing of CMCs. Secondly, a post-test inspection technique is demonstrated to show the effect of high-temperature exposure on electrical properties. Analysis was performed to determine the respective contribution of the fiber and matrix to the overall composite conductivity at elevated temperatures. It was concluded that because the silicon-rich matrix material dominates the electrical response at high temperature, ER monitoring would continue to be a feasible method for monitoring stress dependent matrix cracking of melt-infiltrated SiC/SiC composites under high temperature mechanical testing conditions. Finally, the effect of thermal gradients generated during localized heating of tensile coupons on overall electrical response of the composite is determined.

Ceramic Matrix Composites↗

NASA NTRS2016

Energy Conversion and Storage Requirements for Hybrid Electric Aircraft

Among various options for reducing greenhouse gases in future large commercial aircraft, hybrid electric option holds significant promise. In the hybrid electric aircraft concept, gas turbine engine is used in combination with an energy storage system to drive the fan that propels the aircraft, with gas turbine engine being used for certain segments of the flight cycle and energy storage system being used for other segments. The paper will provide an overview of various energy conversion and storage options for hybrid electric aircraft. Such options may include fuel cells, batteries, super capacitors, multifunctional structures with energy storage capability, thermoelectric, thermionic or a combination of any of these options. The energy conversion and storage requirements for hybrid electric aircraft will be presented. The role of materials in energy conversion and storage systems for hybrid electric aircraft will be discussed.

Energy Conversion↗

NASA NTRS2019

Thermal Analysis of Potted Litz Wire for High-Power-Density Aerospace Electric Machines

Increasing the power density and efficiency of electric machines (motors and generators) is integral to bringing Electrified Aircraft (EA) to commercial realization. To that end an effort to create a High Efficiency Megawatt Motor (HEMM) with a goal of exceeding 98% efficiency and 1.46 MW of power has been undertaken at the NASA Glenn Research Center. Of the motor components the resistive losses in the stator windings are by far the largest contributor (34%) to total motor loss. The challenge is the linear relationship between resistivity and temperature, making machine operation sensitive to temperature increases. In order to accurately predict the thermal behavior of the stator the thermal conductivity of the Litz wire-potting-electrical insulation system must be known. Unfortunately, this multi material system has a wide range of thermal conductivities (0.1 W/m-K â€“ 400 W/m-K) and a high anisotropy (axial vs transverse) making the prediction of the transverse thermal conductivity an in turn the hot spot temperatures in the windings is difficult. In order to do this a device that simulates the thermal environment found in the HEMM stator was designed. This device is not unlike the motorettes (little motors) that are described in IEEE standards for testing electrical insulation lifetimes or other electric motor testing. However, because the HEMM motor design includes significant rotor electrical and thermal considerations the term motorette was not deemed appropriate. Instead statorette (or little stator) was adopted as the term for this test device. This paper discussed the design, thermal heat conjugate analysis (thermal model), manufacturing and testing of HEMM's statorette. Analysis of the results is done by thermal resistance network model and micro thermal model and is compared to analytical predictions of thermal conductivity of the insulated and potted Litz wire system.

Motor↗

NASA NTRS2021

Martian Moons and Space Transportation Using Chemical and Electric Propulsion Options

Using chemical and nuclear electric propulsion for the exploration of the Martian moons will be investigated. Both oxygen/hydrogen chemical propulsion and nuclear electric propulsion with 500 kilowatt electric (kWe) to 10 megawatt electric (MWe) reactors will be assessed. The initial masses, propellant masses, and trip times for a variety of space vehicle payload masses will be compared. For high energy orbital transfer, the nuclear electric propulsion vehicles required a small fraction of the propellant mass over oxygen/hydrogen orbital transfer vehicles (OTVs). The moons, Phobos and Deimos, may hold resources for refueling future space vehicles. In-situ resource utilization (ISRU) can be a powerful method of reducing Earth dependence on space vehicle propellants, liquid water, and breathing gases. Historical studies have identified the potential of water in carbonaceous chondrites on the moons. The moon-derived propellants OTVs that move payloads between the moons and to other important operational Mars orbits. Also, the propellants have been suggested to support reusable Mars landers. To extract the water, the mined mass, its volume and the mining time were estimated. The water mass fraction may be as low as 2x10 −4 . Very large masses were needed to be extracted for up to 100 MT of water.

In-situ resource utilization↗

NASA NTRS2021

Compass Final Report: Nuclear Electric Propulsion (NEP)-Chemical Vehicle 1.2

Many previous studies have examined sending crews to and from Mars. The most economical involved a ‘conjunction’ class whereby the crew spends around 500 days on Mars waiting for a ‘cheap’ return. The total mission time results in over a 1000-day mission duration (about 3 years). Given the current experience level of only one year on the International Space Station (ISS), it of interest to reduce that time to only two years, thus reducing risk and minimizing required Mars surface infrastructure. The Phase 1.1 Study goal was stated as follows, “Determine the feasibility of a two-year roundtrip class Mars mission concept of operation that enables boots on Mars no later than 2036.” While the Phase1 study did show feasibility for the NEP-Chemical option, the 2036 Opposition opportunity was found to stress the schedule due to proposed technology development schedules. A 2039 Opposition (which requires even more energy than the 2036 case) was chosen as representative for Phase 1.2. Phase 1.2 also sought to further refine the concept, building on the feasibility, but addressing several challenges brought by the red team and habitat team. Given the date of 2039, nearer term technologies, primarily nuclear thermal and nuclear electric were deemed as the most viable for these missions. As will be shown, the energy required to perform such a mission in only two years (for the 2039 opportunity at least) is about three times that of the three-year conjunction mission. The rocket equation shows that this mission would then require several times the propellant of the three-year mission unless the specific impulse (ISP) of the propulsion system can be increased. Based on lunar needs, a limit of five Space Launch System (SLS) launchers with 8.4m fairings was imposed for the piloted transportation portion of the mission, limiting the size of the system. When using nuclear electric propulsion, the main limiting factor was packaging the required radiator area. The higher Isp nuclear electric propulsion (NEP) system option is described herein but with a twist: in order to keep the size of radiators packageable in one SLS and use proven reactor power system technology (~1200K reactor outlet temperature and superalloy-class Brayton) the NEP system had to be combined with a chemical propulsion system. This combination of electric propulsion and high thrust chemical was found to be useful in previous design studies combining solar electric propulsion (SEP) and chemical propulsion. Such a combination allowed the low-thrust system to provide significant change in velocity (∆V) during the interplanetary portions of the mission, thereby notably reducing the ∆V required by the high thrust system to capture and depart from the Mars gravity well. Here the high thrust ‘impulsive’ system is more efficient due to the Oberth Effect. A plethora of trades, both at the mission and system level, as well as the subsystem level were performed to develop these vehicle concepts. An entire family of NEP-Chemical transportation vehicles is described herein. The main driver and the primary focus was the piloted vehicle but additional concepts for cargo were performed using the same ‘building blocks’ in order to reduce costs and provide commonality.

Nuclear Power↗

NASA NTRSDate not supplied

NASA Reference Motor Designs for Electric Vertical Takeoff and Landing Vehicles

Electric and hybrid electric vertical takeoff and landing vehicles require high performance and high reliability electric motor drivetrains. Failure analysis of NASA’s Revolutionary Vertical Lift Technologies’ reference vehicles pointed to current electric motor drivetrain reliability being below what is needed to meet the expected stringent reliability requirements for Urban Air Mobility vehicles. In this paper, design studies are carried out for UAM vehicle electric motors to produce reference designs. The primary intent of these reference motor designs is to provide guidance for UAM motor reliability model development and technology advancement. They additionally provide high fidelity motor sizing information for vehicle designers and references for different technologies or motor topologies to be traded against.

Electric Motor Urban Air Mobility↗

NASA NTRSDate not supplied

Considering Turbofan Operability in Hybrid Electric Aircraft Propulsion System Design

This paper explores the design of a hybrid electric aircraft propulsion system that uses a turbofan to power an electric system. In such a system, the gas turbine will experience a loss of power generation as altitude increases, however the electric system will not. This difference results in designs that may over size the electric system at high altitude or under size at low altitude. Two studies are performed within this paper. The first looks at extracting power from the engine for use with electric aircraft propulsion at cruise and the second reviews a design of an engine that uses thrust assist for takeoff. Both studies look at the effects of changing altitude on the amount of power extraction or insertion that can be taken from the turbofan as dictated by operability limits. Results of the paper show that low-pressure compressor surge margin and high-pressure compressor speed can be pushed to unaccepted limits with large scale power extraction or insertion, however these issues can be mitigated by adding power extraction or insertion at off design operating points to compensate. Additionally, the benefits of thrust assist are quantified for this configuration demonstrating a reduction in thrust specific fuel consumption at cruise of over 5%.

Gas Turbine↗

NASA NTRS2023

Considering Turbofan Operability in Hybrid Electric Aircraft Propulsion System Design

This paper explores the design of a hybrid electric aircraft propulsion system that uses a turbofan to power an electric system. In such a system, the gas turbine will experience a loss of power generation as altitude increases, however the electric system will not. This difference results in designs that may over size the electric system at high altitude or under size at low altitude. Two studies are performed within this paper. The first looks at extracting power from the engine for use with electric aircraft propulsion at cruise and the second reviews a design of an engine that uses thrust assist for takeoff. Both studies look at the effects of changing altitude on the amount of power extraction or insertion that can be taken from the turbofan as dictated by operability limits. Results of the paper show that low-pressure compressor surge margin and high-pressure compressor speed can be pushed to unaccepted limits with large scale power extraction or insertion, however these issues can be mitigated by adding power extraction or insertion at off design operating points to compensate. Additionally, the benefits of thrust assist are quantified for this configuration demonstrating a reduction in thrust specific fuel consumption at cruise of over 5%.

Gas turbine↗

NASA NTRSDate not supplied

Considering Turbofan Operability in Hybrid Electric Aircraft Propulsion System Design

This paper explores the design of a hybrid electric aircraft propulsion system that uses a turbofan to power an electric system. In such a system, the gas turbine will experience a loss of power generation as altitude increases, however the electric system will not. This difference results in designs that may over size the electric system at high altitude or under size at low altitude. Two studies are performed within this paper. The first looks at extracting power from the engine for use with electric aircraft propulsion at cruise and the second reviews a design of an engine that uses thrust assist for takeoff. Both studies look at the effects of changing altitude on the amount of power extraction or insertion that can be taken from the turbofan as dictated by operability limits. Results of the paper show that low-pressure compressor surge margin and high-pressure compressor speed can be pushed to unaccepted limits with large scale power extraction or insertion, however these issues can be mitigated by adding power extraction or insertion at off design operating points to compensate. Additionally, the benefits of thrust assist are quantified for this configuration demonstrating a reduction in thrust specific fuel consumption at cruise of over 5%.

Gas turbine↗

NASA NTRSDate not supplied

Scaling Electric Machines to a Megawatt and Material Options

Megawatt (MW) electric aircraft propulsion (EAP) is seen as a significant contributor toward achieving the goals set forth by the Sustainable Flight National Partnership. A large part of enabling MW EAP is developing specific-power-dense electric machines. As specific-power-dense electric machines are scaled up from kW to MW power levels, the thermal stresses on the machines increase in both magnitude and performance-affecting characteristics. This is particularly true for the stators of these machines. Analysis via thermal resistance network modeling and multiscale modeling reveals that increasing amounts of heat will be trapped in the stator windings as the power levels increase. The challenges this presents can be addressed through material advancements whereby materials gain multifunctionality. Specifically, the electrical insulation and potting materials, along with the electrical conductor, that compose the stator slot must work together (gain multifunctionality) to relieve the increased thermal stress. Materials research at the NASA Glenn Research Center points to some useful solutions in this trade space.

Electric Machine↗

NASA NTRS2023

Scaling Electric Machines to a Megawatt and Material Options

Megawatt (MW) electric aircraft propulsion (EAP) is seen as a significant contributor toward achieving the goals set forth by the Sustainable Flight National Partnership. A large part of enabling MW EAP is developing specific-power-dense electric machines. As specific-power-dense electric machines are scaled up from kW to MW power levels, the thermal stresses on the machines increase in both magnitude and performance-affecting characteristics. This is particularly true for the stators of these machines. Analysis via thermal resistance network modeling and multiscale modeling reveals that increasing amounts of heat will be trapped in the stator windings as the power levels increase. The challenges this presents can be addressed through material advancements whereby materials gain multifunctionality. Specifically, the electrical insulation and potting materials, along with the electrical conductor, that compose the stator slot must work together (gain multifunctionality) to relieve the increased thermal stress. Materials research at the NASA Glenn Research Center points to some useful solutions in this trade space.

Electric Machine↗

NASA NTRS2023

Considering Turbofan Operability in Hybrid Electric Aircraft Propulsion System Design

This paper explores the design of a hybrid electric aircraft propulsion system that uses a turbofan to power an electric system. In such a system, the gas turbine will experience a loss of power generation as altitude increases, however the electric system will not. This difference results in designs that may over size the electric system at high altitude or under size at low altitude. Two studies are performed within this paper. The first looks at extracting power from the engine for use with electric aircraft propulsion at cruise and the second reviews a design of an engine that uses thrust assist for takeoff. Both studies look at the effects of changing altitude on the amount of power extraction or insertion that can be taken from the turbofan as dictated by operability limits. Results of the paper show that low-pressure compressor surge margin and high-pressure compressor speed can be pushed to unaccepted limits with large scale power extraction or insertion, however these issues can be mitigated by adding power extraction or insertion at off design operating points to compensate. Additionally, the benefits of thrust assist are quantified for this configuration demonstrating a reduction in thrust specific fuel consumption at cruise of over 5%.

Gas Turbine↗

NASA NTRS2024

Mission and Vehicle-Level Updates for the Parallel Electric-Gas Architecture with Synergistic Utilization Scheme (PEGASUS) Concept Aircraft

NASA created the PEGASUS concept with the goal of lowering mission energy (a surrogate for operating cost) compared to other regional aircraft by leveraging electrified aircraft propulsion (EAP). Since its inception, researchers have explored multiple facets of PEGASUS in varying fidelity but have not completed a rigorous, integrated design. The goal of this memorandum is to provide an updated design using recent studies and improved methods. This memorandum explores the initial vehicle concept and concept of operations, while considering ways to improve both the mission concept of operations and the integrated vehicle-level performance. Additionally, the design and analysis methodologies for EAP-enabled aircraft concepts are improved in several areas. This research incorporates new propulsion-airframe integration and wing weight surrogates to model the impacts of wingtip propulsors on the configuration. Detailed weight and balance calculations enable calculating dynamic stability and flight qualities within the conceptual design environment. Ultimately, the vehicle is optimized to reduce "well-to-wake" equivalent CO2, CO2e, rather than minimizing either fuel (or total energy) consumption or maximum takeoff weight. Using fuel/energy or takeoff weight leads to conflicting optima for hybrid-electric aircraft. Two aircraft are developed to provide points of comparison for PEGASUS: an advanced conventional turboprop vehicle and a hybrid-electric variant. The results show that the PEGASUS concept can reduce CO2e relative to the advanced turboprop or a hybrid-electric propulsion architecture, albeit with an increase in maximum takeoff weight. PEGASUS's maximum takeoff weight is 55% heavier than the advanced conventional turboprop but releases 18% less CO2e for a 400 nmi mission. Over the same mission, PEGASUS's maximum takeoff weight is 47% heavier than the comparator hybrid-electric vehicle but releases 12% less CO2e. This study shows that the PEGASUS configuration reduces CO2e through its use of wingtip propulsors and that its benefit is not solely a result of switching to a hybrid-electric propulsion architecture. PEGASUS achieves this reduction in CO2e while maintaining satisfactory Level 1 or 2 flight qualities for all of its longitudinal- and lateral-directional modes.

Hybrid-electric↗

NASA NTRSDate not supplied

Multidisciplinary Optimization of a Transonic Truss Braced Wing Aircraft with Hybrid-Electric Propulsion

Next generation aircraft concepts have subsystems that are increasingly inter-connected. This necessitates advanced design tools using gradient based optimization to properly account for strong subsystem coupling in these aircraft. Additionally, these design tools require accurate modeling capability to support the increased interest in electrified propulsion. Design studies for these aircraft must consider the trade-offs between electric propulsion and turbojet engines at all points in the flight envelope to determine the optimal balance between propulsion options. This paper describes the development of an electric propulsion subsystem model designed to work within Aviary - an open source aircraft design tool. We will demonstrate the electric propulsion model operating with Aviary by showing results from an optimized Transonic Truss-Braced Wing (TTBW) concept using assisted electric propulsion during climb. First, we present results with and without electrification to show the overall system impacts of hybrid electric propulsion during climb. We also present two distinct battery models compatible with this optimization framework, and compare results using both. Next, we ran the same problem using a slightly different cell type, and demonstrate a considerable change in the result. Finally, we vary the cell energy density of the batteries, and provide an illustrative trend for the system-level impact for improving cell technology.

optimization↗

NASA NTRSDate not supplied

Subscale Hardware-In-The-Loop Results for Hybrid Electric Turbofan Controls Use Cases

NASA is investigating hybrid electric turbine engine systems for commercial transport aircraft due to the potentially significant improvements hybrid electric technology offers in performance, fuel consumption, and operational and design flexibility. Recently, the technology has been tested at full scale in partnership with industry and advanced to Technology Readiness Level 4. This presentation will focus on a recent subscale hardware-in-the-loop test of an open source turbofan engine model developed by NASA. The Advanced Geared Turbofan 30,000 lbf – electrified (AGTF30-e) engine is used as a reference model to demonstrate control system design and use cases for an example mild hybrid electric system with no large-scale energy storage. This model is run in real-time in NASA’s Hybrid Propulsion Emulation Rig (HyPER) and is used to drive an emulation of the turbomachinery system using subscale electric machines. This dynamic scaled shaft emulation interacts with a subscale (<100 kW) hybrid system consisting of electric machines, motor controllers, and a programmable electronic load. Specific use cases demonstrated include the use of Turbine Electrified Energy Management to improve operation during transients, megawatt-scale power extraction from the AGTF30-e, and power transfer between engine spools. Results related to the effectiveness of hybrid systems are qualitatively compared to results from industry testing.

Hybrid↗