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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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All sourcesNASA NTRSDOE OSTI
NASA NTRS2011

Thermal Management Tools for Propulsion System Trade Studies and Analysis

Energy-related subsystems in modern aircraft are more tightly coupled with less design margin. These subsystems include thermal management subsystems, vehicle electric power generation and distribution, aircraft engines, and flight control. Tighter coupling, lower design margins, and higher system complexity all make preliminary trade studies difficult. A suite of thermal management analysis tools has been developed to facilitate trade studies during preliminary design of air-vehicle propulsion systems. Simulink blocksets (from MathWorks) for developing quasi-steady-state and transient system models of aircraft thermal management systems and related energy systems have been developed. These blocksets extend the Simulink modeling environment in the thermal sciences and aircraft systems disciplines. The blocksets include blocks for modeling aircraft system heat loads, heat exchangers, pumps, reservoirs, fuel tanks, and other components at varying levels of model fidelity. The blocksets have been applied in a first-principles, physics-based modeling and simulation architecture for rapid prototyping of aircraft thermal management and related systems. They have been applied in representative modern aircraft thermal management system studies. The modeling and simulation architecture has also been used to conduct trade studies in a vehicle level model that incorporates coupling effects among the aircraft mission, engine cycle, fuel, and multi-phase heat-transfer materials.

McCarthy, Kevin↗
NASA NTRS1997

Performance of an Electro-Hydrostatic Actuator on the F-18 Systems Research Aircraft

An electro-hydrostatic actuator was evaluated at NASA Dryden Flight Research Center, Edwards, California. The primary goal of testing this actuator system was the flight demonstration of power-by-wire technology on a primary flight control surface. The electro-hydrostatic actuator uses an electric motor to drive a hydraulic pump and relies on local hydraulics for force transmission. This actuator replaced the F-18 standard left aileron actuator on the F-18 Systems Research Aircraft and was evaluated throughout the Systems Research Aircraft flight envelope. As of July 24, 1997 the electro-hydrostatic actuator had accumulated 23.5 hours of flight time. This paper presents the electro-hydrostatic actuator system configuration and component description, ground and flight test plans, ground and flight test results, and lessons learned. This actuator performs as well as the standard actuator and has more load capability than required by aileron actuator specifications of McDonnell- Douglas Aircraft, St. Louis, Missouri. The electro-hydrostatic actuator system passed all of its ground tests with the exception of one power-off test during unloaded dynamic cycling.

Robert Navarro↗
NASA NTRS2021

Thermal Systems Modeling of Chemical Integrated Power Source (CHIPS) to Survive Lunar Night Environments

This paper presents the results of the systems level thermal modeling for a conceptual Chemical Heat Integrated Power Source (CHIPS). This proposed system offers a combined thermal and electrical power source to support survival of spacecraft operating in extreme low temperature lunar environments, without the use of radioisotope-based sources. A conceptual design study has been completed for this system, that uses heat generated by an exothermic chemical reaction in place of radioisotope or electrical heat sources. The goal of the study was to evaluate the feasibility of such as system through thermodynamic and chemical analysis and thermal modeling, and to identify technology gaps to inform a technology development and maturation plan. The specific technical objectives focused on delivery of 90-100 Wth thermal power and 30-40 We electrical power for 336 hours to a representative Commercial Lunar Payload Services (CLPS) lander, to support lunar surface survival and limited operations through a lunar night. The total system mass was targeted at ≤50 kg. A highly exothermic chemical reaction system is used to generate on-board electrical power for spacecraft systems, and thermal power to maintain critical spacecraft/lander systems within their allowable flight temperature (AFTs). Based on the very high energy content of the chemical reaction system, a much higher amount of heat per unit mass can be delivered to the spacecraft, relative to a rechargeable lithium-ion battery and electrical heater(s). Since radioisotope heaters or generators are not used, the system will be orders of magnitude lower in cost than a radioisotope heating/power unit, without the attendant regulatory complexities. As part of the CHIPS concept, a fraction of the thermal power generated is converted to electrical power via an appropriate thermal-to-electric converter technology (such as a free piston Stirling converter or a thermoelectric generator module), to provide power to critical loads. This approach is ideally suited to support operation of commercial landers (e.g., via the CLPS program). In most cases, these landers are only designed to operate during a portion of the lunar day, with no provision for survival through the lunar night. By supporting lunar night survival, a mission can be extended through the lunar night and at least into another lunar day, thus turning a nominal eight-day mission into a 36-day mission. Therefore, the current system demonstration will focus on scalability to support at least 336 hours (one lunar night) of continuous thermal and electrical power generation. Although initially targeted to support lunar equatorial landings, the technology is extensible to missions at the lunar poles, other extreme environments in the Solar System or even air-independent applications on Earth (e.g., ocean exploration).

West, William↗
NASA NTRS2021

Thermal Systems Modeling of Chemical Heat Integrated Power Source (CHIPS) to Survive Lunar Night Environments

This paper presents the results of the systems level thermal modeling for a conceptual Chemical Heat Integrated Power Source (CHIPS). This proposed system offers a combined thermal and electrical power source to support survival of spacecraft operating in extreme low temperature lunar environments, without the use of radioisotope-based sources. A conceptual design study has been completed for this system, that uses heat generated by an exothermic chemical reaction in place of radioisotope or electrical heat sources. The goal of the study was to evaluate the feasibility of such a system through thermodynamic and chemical analysis and thermal modeling, and to identify technology gaps to inform a technology development and maturation plan. The specific technical objectives focused on delivery of 90 to 100 Wth thermal power and 30 to 40 We electrical power for 336 hours to a representative Commercial Lunar Payload Services (CLPS) lander, to support lunar surface survival and limited operations through a lunar night. The total system mass was targeted at ≤50 kg. A highly exothermic chemical reaction system is used to generate on-board electrical power for spacecraft systems, and thermal power to maintain critical spacecraft/lander systems within their allowable flight temperature (AFTs). Based on the very high energy content of the chemical reaction system, a much higher amount of heat per unit mass can be delivered to the spacecraft, relative to a rechargeable lithium-ion battery and electrical heater(s). Since radioisotope heaters or generators are not used, the system will be orders of magnitude lower in cost than a radioisotope heating/power unit, without the attendant regulatory complexities. As part of the CHIPS concept, a fraction of the thermal power generated is converted to electrical power via an appropriate thermal-to-electric converter technology (such as a free piston Stirling converter or a thermoelectric generator module), to provide power to critical loads. This approach is ideally suited to support operation of commercial landers (e.g., via the CLPS program). In most cases, these landers are only designed to operate during a portion of the lunar day, with no provision for survival through the lunar night. By supporting lunar night survival, a mission can be extended through the lunar night and at least into another lunar day, thus turning a nominal eight-day mission into a 36-day mission. Therefore, the current system demonstration will focus on scalability to support at least 336 hours (one lunar night) of continuous thermal and electrical power generation. Although initially targeted to support lunar equatorial landings, the technology is extensible to missions at the lunar poles, other extreme environments in the Solar System or even air-independent applications on Earth (e.g., ocean exploration).

West, William↗
NASA NTRS2019

Physics-Based Modeling and Simulation of Emerging Battery Technologies for Aerospace

Recently there is a growing interest in the aviation sector to reduce air and noise pollution. Electrochemical energy storage devices such as batteries coupled with a distributed electric propulsion system can reduce noise concerns as well as emissions and allow the concept of Urban Air Mobility to come to fruition. The battery performance needs to improve considerably compared from current state-of-art Li-ion battery (about 200Wh/Kg) to realize all-electric passenger jet for short flights (up to 690 miles). NASA is exploring lithium-oxygen battery chemistry to power hybrid (battery powered electrical system) and all-electric aircraft for short distance and long-distance flights. Li-O2 is one of the advanced Li-ion technologies that promise to provide specific energy of more than 750Wh/Kg. For this presentation, we present our work on improving power density of Li-O2 batteries through the use of multiphysics simulations. Next, a path is outlined to port these model to simulate performance for a new battery chemistry for space application, Li-CO2. Li-CO2 uses carbon dioxide as the active material instead of oxygen. Although this technology is in its early development, the offers two benefits: it can be used as a CO2 scrubber, as oxygen is one of the by-products on charging, and as a backup or a standalone battery for various Mars or Venus missions, where the carbon dioxide content in the atmosphere is high and need battery to operate at higher temperatures.

Mehta, Mohit↗
NASA NTRS1991

Civil air transport: A fresh look at power-by-wire and fly-by-light

Power-by-wire (PBW) is a key element under subsonic transport flight systems technology with potential savings of over 10 percent in operating empty weight and in fuel consumption compared to today's transport aircraft. The PBW technology substitutes electrical actuation in place of centralized hydraulics, uses internal starter-motor/generators and eliminates the need for variable engine bleed air to supply cabin comfort. The application of advanced fiber optics to the electrical power system controls, to built-in-test (BIT) equipment, and to fly-by-light (FBL) flight controls provides additional benefits in lightning and high energy radio frequency (HERF) immunity over existing mechanical or even fly-by-wire controls. The program plan is reviewed and a snapshot is given of the key technologies and their benefits to all future aircraft, both civil and military.

Sundberg, Gale R.↗
NASA NTRS2017

A Novel UAV Electric Propulsion Testbed for Diagnostics and Prognostics

This paper presents a novel hardware-in-the-loop (HIL) testbed for systems level diagnostics and prognostics of an electric propulsion system used in UAVs (unmanned aerial vehicle). Referencing the all electric, Edge 540T aircraft used in science and research by NASA Langley Flight Research Center, the HIL testbed includes an identical propulsion system, consisting of motors, speed controllers and batteries. Isolated under a controlled laboratory environment, the propulsion system has been instrumented for advanced diagnostics and prognostics. To produce flight like loading on the system a slave motor is coupled to the motor under test (MUT) and provides variable mechanical resistance, and the capability of introducing nondestructive mechanical wear-like frictional loads on the system. This testbed enables the verification of mathematical models of each component of the propulsion system, the repeatable generation of flight-like loads on the system for fault analysis, test-to-failure scenarios, and the development of advanced system level diagnostics and prognostics methods. The capabilities of the testbed are extended through the integration of a LabVIEW-based client for the Live Virtual Constructive Distributed Environment (LVCDC) Gateway which enables both the publishing of generated data for remotely located observers and prognosers and the synchronization the testbed propulsion system with vehicles in the air. The developed HIL testbed gives researchers easy access to a scientifically relevant portion of the aircraft without the overhead and dangers encountered during actual flight.

Gorospe, George E., Jr.↗
NASA NTRS1990

Civil air transport - A fresh look at power-by-wire and fly-by-light

Power-by-wire (PBW) is a key element under subsonic transport flight systems technology, with potential savings of over 10 percent in gross take off weight and in fuel consumption compared to today's transport aircraft. The PBW technology substitutes electrical actuation in place of centralized hydraulics, uses internal starter-motor/generators, and eliminates the need for variable engine bleed air to supply cabin comfort. The application of advanced fiber optics to the electrical power system controls, to built-in-test (BITE) equipment, and to fly-by-light (FBL) flight controls provides additional benefits in lightning and high-energy radio frequency (HERF) immunity over existing mechanical or even fly-by-wire controls. The program plan is reviewed and a snapshot is given of the key technologies and their benefits to future aircraft, both civil and military.

Sundberg, Gale R.↗
NASA NTRS2000

Reliability Modeling of Microelectromechanical Systems Using Neural Networks

Microelectromechanical systems (MEMS) are a broad and rapidly expanding field that is currently receiving a great deal of attention because of the potential to significantly improve the ability to sense, analyze, and control a variety of processes, such as heating and ventilation systems, automobiles, medicine, aeronautical flight, military surveillance, weather forecasting, and space exploration. MEMS are very small and are a blend of electrical and mechanical components, with electrical and mechanical systems on one chip. This research establishes reliability estimation and prediction for MEMS devices at the conceptual design phase using neural networks. At the conceptual design phase, before devices are built and tested, traditional methods of quantifying reliability are inadequate because the device is not in existence and cannot be tested to establish the reliability distributions. A novel approach using neural networks is created to predict the overall reliability of a MEMS device based on its components and each component's attributes. The methodology begins with collecting attribute data (fabrication process, physical specifications, operating environment, property characteristics, packaging, etc.) and reliability data for many types of microengines. The data are partitioned into training data (the majority) and validation data (the remainder). A neural network is applied to the training data (both attribute and reliability); the attributes become the system inputs and reliability data (cycles to failure), the system output. After the neural network is trained with sufficient data. the validation data are used to verify the neural networks provided accurate reliability estimates. Now, the reliability of a new proposed MEMS device can be estimated by using the appropriate trained neural networks developed in this work.

Perera. J. Sebastian↗
NASA NTRS1990

Civil air transport: A fresh look at power-by-wire and fly-by-light

Power-by-wire (PBW) is a key element under subsonic transport flight systems technology with potential savings of over 10 percent in gross take-off-weight and in fuel consumption compared to today's transport aircraft. The PBW technology substitutes electrical actuation in place of centralized hydraulics, uses internal starter-motor/generators and eliminates the need for variable engine bleed air to supply cabin comfort. The application of advanced fiber optics to the electrical power system controls, to built-in-test (BITE) equipment, and to fly-by-light (FBL) flight controls provides additional benefits in lightning and high energy radio frequency (HERF) immunity over existing mechanical or even fly-by-wire controls. The program plan is reviewed and a snapshot is given of the key technologies and their benefits to all future aircraft, both civil and military.

Sundberg, Gale R.↗
NASA NTRS2018

A Risk Analysis Tool for Estimating the Risk of Electrical Failures Due to Human Induced Defects

Aerospace electrical systems are required to withstand and adequately operate in extremely harsh environments that include, for example, high radiation exposure, temperature extremes, intense vibrational stress and drastic temperature cycling. The nature of aerospace electronics also demands high reliability since, with very few exceptions, there is no chance for hardware servicing or repairs. Common risk mitigation techniques for this type of situation are to perform a Reliability Analysis of the system throughout the development cycle, and to use electrical components that are regarded as “high reliability” because of additional controls and requirements applied in their design, manufacturing and testing. Unfortunately, studies have shown that even though these techniques are used, many systems fail to meet mission requirements well before the predicted lifetimes. This paper presents the analysis of failures of electrical parts, experienced during various stages of system development, at NASA Goddard Space Flight Center, Greenbelt MD, between the years 2001 and 2013. These components were subjected to qualification, screening and testing in which the goal was to ensure that the components would survive the stresses of the mission. The analysis categorizes failures by part type and failure mechanisms. One of the results of the analysis was the realization that a surprising proportion of failures experienced during system integration and testing were caused by human error (i.e. human induced defect). Further analysis included the determination of root failure mechanisms and any influencing factors contributing to these failures. The major causes of these defects were attributed to electrostatic damage (ESD), electrical overstress (EOS), mechanical overstress (MOS), and thermal overstress (TOS). Finally, the study proposes a risk analysis tool which incorporates these major causes for the failures, termed error-producing conditions (EPCs), and a proportionality factor representing the number of each type of failure that has occurred at the facility under study. These factors are quantified and used to communicate the risk of human induced defects for the assembly, integration and testing of space hardware based on the system’s electrical parts list. The new risk identification can trigger risk-mitigating actions more effectively, based on the presence of component categories or other hazardous conditions that have a history of failure due to human error.

Majewicz, Peter J.↗
NASA NTRS2018

Development of High Power Hall Thruster Systems to Enable the NASA Exploration Vision

The next phase of space exploration missions requires high power Solar Electric Propulsion (SEP) systems for large-scale science missions and cargo transportation. Development is underway at Aerojet Rocketdyne on Hall thruster systems that are intended to bracket the needs of future NASA SEP missions in support of space exploration. The Advanced Electric Propulsion System (AEPS) program is developing and qualifying a 13.3kW Hall thruster system to be demonstrated on the Power and Propulsion Element (PPE), which is intended to be the first element of a Lunar Outpost Platform - Gateway (LOP-G). The NextSTEP program is integrating a nested Hall thruster into a 100kW system and testing it for 100 hours. These two programs will provide a path to efficient in-space propulsion that will allow NASA to transfer the large amounts of cargo that is needed to support human missions - first to the moon and then on to Mars. The Advanced Electric Propulsion System (AEPS) program is completing development, qualification and delivery of five flight 13.3kW EP systems to NASA. The flight AEPS system includes a magnetically shielded long-life Hall thruster, Power Processing Unit (PPU) and a Xenon Flow Controller (XFC). The Hall thruster, developed and demonstrated by NASA, operates at input powers up to 12.5kW while providing a specific impulse over an estimated 2800s at an input voltage of 600V. The power processor is designed to accommodate an input voltage range of 95-140V, consistent with operation beyond the orbit of Mars. The integrated system input power is continuously throttleable between 3 and 13.3kW. Component level testing of the EP String has begun with prototype hardware. The NextSTEP program is developing a 100kW Electric Propulsion (EP) system using a nested Hall thruster designed for powers up to 250kW, a modular power processor and a modular mass flow controller. While the program objective is to operate the integrated EP system continuously at 100kW for 100hrs to demonstrate thermal stability and support the development of system life time models, it builds on decades of experience with long-life Hall thrusters and the design is evolvable to a capability of 250kW. Design upgrades that demonstrate the 100kW EP system have been completed and tested. Aerojet Rocketdyne (AR) is excited to support NASA as it extends human reach into deep space and believes that these programs will provide the propulsion to make such missions affordable and sustainable. These systems provide NASA with a range of options to power its deep space transport vehicles. This paper presents the mission requirements for supporting the NASA exploration vision, as well as the status for the high power Hall thruster systems in development.

Jackson, Jerry↗
NASA NTRS2018

Development of High Power Hall Thruster Systems to Enable the NASA Exploration Vision

The next phase of space exploration missions requires high power Solar Electric Propulsion (SEP) systems for large-scale science missions and cargo transportation. Development is underway at Aerojet Rocketdyne on Hall thruster systems that are intended to bracket the needs of future NASA SEP missions in support of space exploration. The Advanced Electric Propulsion System (AEPS) program is developing and qualifying a 13.3kW Hall thruster system to be demonstrated on the Power and Propulsion Element (PPE), which is intended to be the first element of a Lunar Outpost Platform - Gateway (LOP-G). The NextSTEP program is integrating a nested Hall thruster into a 100 kW system and testing it for 100 hours. These two programs will provide a path to efficient in-space propulsion that will allow NASA to transfer the large amounts of cargo that is needed to support human missions - first to the moon and then on to Mars. The Advanced Electric Propulsion System (AEPS) program is completing development, qualification and delivery of five flight 13.3kW EP systems to NASA. The flight AEPS system includes a magnetically shielded long-life Hall thruster, Power Processing Unit (PPU) and a Xenon Flow Controller (XFC). The Hall thruster, developed and demonstrated by NASA, operates at input powers up to 12.5 kW while providing a specific impulse over an estimated 2800s at an input voltage of 600V. The power processor is designed to accommodate an input voltage range of 95-140V, consistent with operation beyond the orbit of Mars. The integrated system input power is continuously throttleable between 3 and 13.3kW. Component level testing of the EP String has begun with prototype hardware. The NextSTEP program is developing a 100kW Electric Propulsion (EP) system using a nested Hall thruster designed for powers up to 250kW, a modular power processor and a modular mass flow controller. While the program objective is to operate the integrated EP system continuously at 100kW for 100 hours to demonstrate thermal stability and support the development of system life time models, it builds on decades of experience with long-life Hall thrusters and the design is evolvable to a capability of 250kW. Design upgrades that demonstrate the 100kW EP system have been completed and tested. Aerojet Rocketdyne is excited to support NASA as it extends human reach into deep space and believes that these programs will provide the propulsion to make such missions affordable and sustainable. These systems provide NASA with a range of options to power its deep space transport vehicles. This paper presents the mission requirements for supporting the NASA exploration vision, as well as the status for the high power Hall thruster systems in development.

Jackson, Jerry↗
NASA NTRS2023

Modeling and Simulation of a Parallel Hybrid Electric Regional Aircraft for the Electrified Powertrain Flight Demonstration (EPFD) Program

This paper presents a parametric modeling and integrated aircraft sizing and synthesis approach for a charge depleting parallel hybrid electric architecture. The developed models are integrated within the baseline thin-haul and regional aircraft. In addition to the physical architecture, different modes of operation enabled by propulsion system electrification are also modeled parametrically. The modes of operation presented in this paper are the peak power shaving, climb power electric boost, in-flight battery recharging, and electric taxi. The sizing of the powertrain and the aircraft are performed within the multidisciplinary analysis and optimization environment, E-PASS. The consideration of the physical system and its operation together provides a holistic approach where the propulsion system and the airframe are designed under an optimized power and energy management strategy. The parametric nature of the work enables the design space exploration for electrification and lays the groundwork for future technology projection and uncertainty quantification studies. The developed capability is generic and can be applied on other aircraft classes. The work is done as part of the Electrified Powertrain Flight Demonstration program.

electrified powertrain flight demonstration↗
NASA NTRS2014

The ISS 2B PVTCS Ammonia Leak: An Operational History

In 2006, the Photovoltaic Thermal Control System (PVTCS) for the International Space Station's 2B power channel began leaking ammonia at a rate of approximately 1.5lbm/year (out of a starting approximately 53lbm system ammonia mass). Initially, the operations strategy was "feed the leak," a strategy successfully put into action via Extra Vehicular Activity during the STS‐134 mission. During this mission the system was topped off with ammonia piped over from a separate thermal control system. This recharge was to have allowed for continued power channel operation into 2014 or 2015, at which point another EVA would have been required. Without these periodic EVAs to refill the 2B coolant system, the channel would eventually leak enough fluid as to risk pump cavitation and system failure, resulting in the loss of the 2B power channel - the most critical of the Space Station's 8 power channels. In mid‐2012, the leak rate increased to approximately 5lbm/year. Once discovered, an EVA was planned and executed within a 5 week timeframe to drastically alter the architecture of the PVTCS via connection to a dormant thermal control system not intended to be utilized as anything other than spare components. The purpose of this rerouting of the TCS was to increase system volume and to isolate the photovoltaic radiator, thought to be the likely leak source. This EVA was successfully executed on November 1st, 2012 and left the 2B PVTCS in a configuration where the system was now being adequately cooled via a totally different radiator than what the system was designed to utilize. Unfortunately, data monitoring over the next several months showed that the isolated radiator was not leaking, and the system itself continued to leak steadily until May 9th, 2013. It was on this day that the ISS crew noticed the visible presence of ammonia crystals escaping from the 2B channel's truss segment, signifying a rapid acceleration of the leak from 5lbm/year to 5lbm/day. Within 48 hours of the crew noticing the leak, an EVA was in progress to replace the coolant pump - the only other replaceable leak source. This paper will explore the management of the 2B PVTCS leak from the operations perspective. It will discuss the methodology of performing the STS‐134 refill, the considerations and contingency plans which went into the architectural overhaul of the system in 2012, and the unprecedented effort which went into the EVA response to the visible leak of May 2013. In particular the paper will focus on the techniques utilized by flight controllers to monitor the system health and to respond to such instances as the rapid May 2013 leak by putting the electrical system in a safe configuration for loss of cooling, and will use recorded telemetry of these events to describe system response to EVA crew and ground actions. It will discuss the innovative design for redundancy of the integrated truss structure's cooling systems which allowed for this leak to be managed with minimal impact to other ISS operations and electrical services, contrasted against the real unintended operations consequences of utilizing the flexibility of the spacecraft's design in this manner. The paper will discuss how the training of the crew and flight controller personnel has adapted to the changing architecture of the power system and the unpredictable nature of the 2B leak.

Vareha, Anthony↗
NASA NTRS2021

Reliability and Safety Assessment of Urban Air Mobility Concept Vehicles

The following work was commissioned by the National Aeronautics and Space Administration (NASA) to guide industry and future regulation related to urban air mobility (UAM). Prior studies compared the relative safety of NASA concept vehicles, Figure ES1, designed for UAM and provided recommendations for industry research, aircraft architectural improvements, and regulatory updates. After the prior study completed, the European Aviation Safety Agency (EASA) released regulatory guidance in the form of a special condition (SC), SC-VTOL-01, for multirotors with distributed propulsion and flight controls (DPFC). The objective of the current work was to develop DPFC architectures that will comply with SC-VTOL-01. Vehicle designs, DPFC architectures, and stability & control (S&C) models were developed to find limitations and trends to guide industry. To guide this task, NASA developed quad, hex, and an octorotor to better define vehicle attributes and trade space. Aircraft for study included electric, hybrid-electric, and turboshaft powerplants and collective and RPM control schemes. Assessments are in terms of the safety level achieved, and/or aircraft component/features needed to meet SC-VTOL-01. The most challenging criteria being the catastrophic failure rate, ≤10-9 catastrophic failures per flight hour, and that no single failures may result in a catastrophic event. A disciplined process was followed, similar to that in Aerospace Recommended Practice (ARP) 4761. A preliminary system safety assessment (PSSA) leveraged prior work as a basis of creating failure rate budgets for system design teams. System designs were updated and iterated upon, working with reliability and safety subject matter experts to develop SC-VTOL-01 compliant designs. Design changes were reflected in updated PSSAs for initial verification of compliance. The DPFC architecture was broken into four system design teams, the (1) flight control system (FCS), (2) drive and power system, (3) thermal management system (TMS), and (4) electrical power and distribution system. The FCS including elements necessary to control the aircraft, drive and power including elements necessary to generate and transmit shaft power, the TMS including elements necessary to maintain temperature limits in all operating environments, and electrical pow-er and distribution including equipment necessary to store and transmit electrical energy. Results found that all aircraft evaluated may have paths to comply with SC-VTOL-01, Figure ES2. However, S&C models showed large power transients that must be addressed and PSSA results show that future work is needed in single load path structures, high voltage power storage and distribution, and in motor/rotor overspeed protection.

Boeing↗
DOE OSTI2021

Google/Makani Energy Kite Modeling (CRADA CRD-18-00569 Final Report)

In support of Makani’s energy kite development program, Makani engaged with NREL to develop, verify, and document a multiphysics engineering model of a megawatt-scale tethered energy kite (named KiteFAST). After the original development of KiteFAST was completed, Makani further engaged NREL to develop a software enhancement of KiteFAST to enable the dynamics modeling of an energy kite tethered to a floating offshore platform (named KiteFAST-OS). Before this project, the capability to model the aero-hydro-servo-elastic dynamics of an airborne wind energy (AWE) system with electricity generation on the flying device (fly gen) and crosswind flight operation did not exist. Such physics-based modeling capability is needed for loads analysis, structural design, and certification of energy kites. Throughout the project, an exclusive license kept KiteFAST and KiteFAST-OS available only to Makani, NREL, and its counterparts. With the ending of Makani, the KiteFAST and KiteFAST-OS software were merged into a single code base (named KiteFAST), the exclusive license was terminated, and the source code, documentation, and example models were released publicly.

17 WIND ENERGY↗