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NASA NTRSDate not supplied

Li-ion Battery Aging with Hybrid Physics-Informed Neural Networks and Fleet-wide Data

In this work, we propose a hybrid model for Li-ion battery discharge and aging prediction that leverages fleet-wide data to predict future capacity drops.The model is built upon an hybrid approach merging physics-based and empirical equations, as well as neural network models in a recurrent neural network cell. The hybrid physics-informed neural network can predict voltage discharge cycles given the loading profile, and estimate the used capacity of the battery under random-loading conditions by tracking aging parameters connected to the residual capacity of the battery. By merging information on the battery aging parameters with existing fleet-wide aging data, the model can predict the future residual capacity of the battery that is being monitored, and therefore enable predictions of voltage discharge curves far ahead in the battery life cycle. We validated the approach using the NASA Prognostics Data Repository Battery data-set, which contains experimental data on Li-ion batteries discharged at random loading conditions in a controlled environment. The approach also allows the identification of discrepancies between the battery aging trend and the trend observed at the fleet level, so that batteries behaving differently from the rest of the fleet can be subject to closer monitoring and further testing to refine predictions.

PINN↗

NASA NTRS1987

DACS II - A distributed thermal/mechanical loads data acquisition and control system

A distributed data acquisition and control system has been developed for the NASA Flight Loads Research Facility. The DACS II system is composed of seven computer systems and four array processors configured as a main computer system, three satellite computer systems, and 13 analog input/output systems interconnected through three independent data networks. Up to three independent heating and loading tests can be run concurrently on different test articles or the entire system can be used on a single large test such as a full scale hypersonic aircraft. Thermal tests can include up to 512 independent adaptive closed loop control channels. The control system can apply up to 20 MW of heating to a test specimen while simultaneously applying independent mechanical loads. Each thermal control loop is capable of heating a structure at rates of up to 150 F per second over a temperature range of -300 to +2500 F. Up to 64 independent mechanical load profiles can be commanded along with thermal control. Up to 1280 analog inputs monitor temperature, load, displacement and strain on the test specimens with real time data displayed on up to 15 terminals as color plots and tabular data displays. System setup and operation is accomplished with interactive menu-driver displays with extensive facilities to assist the users in all phases of system operation.

Zamanzadeh, Behzad↗

NASA NTRS1994

Modeling and life prediction methodology for Titanium Matrix Composites subjected to mission profiles

Titanium matrix composites (TMC) are being evaluated as structural materials for elevated temperature applications in future generation hypersonic vehicles. In such applications, TMC components are subjected to complex thermomechanical loading profiles at various elevated temperatures. Therefore, thermomechanical fatigue (TMF) testing, using a simulated mission profile, is essential for evaluation and development of life prediction methodologies. The objective of the research presented in this paper was to evaluate the TMF response of the (0/90)2s SCS-6/Timetal-21S subjected to a generic hypersonic flight profile and its portions with a temperature ranging from -130 C to 816 C. It was found that the composite modulus, prior to rapid degradation, had consistent values for all the profiles tested. A micromechanics based analysis was used to predict the stress-strain response of the laminate and of the constituents in each ply during thermomechanical loading conditions by using only constituent properties as input. The fiber was modeled as elastic with transverse orthotropic and temperature dependent properties. The matrix was modeled using a thermoviscoplastic constitutive relation. In the analysis, the composite modulus degradation was assumed to result from matrix cracking and was modeled by reducing the matrix modulus. Fatigue lives of the composite subjected to the complex generic hypersonic flight profile were well correlated using the predicted stress in 0 degree fibers.

Mirdamadi, M.↗

NASA NTRS2000

Reliability Constrained Priority Load Shedding for Aerospace Power System Automation

The need for improving load shedding on board the space station is one of the goals of aerospace power system automation. To accelerate the optimum load-shedding functions, several constraints must be involved. These constraints include congestion margin determined by weighted probability contingency, component/system reliability index, generation rescheduling. The impact of different faults and indices for computing reliability were defined before optimization. The optimum load schedule is done based on priority, value and location of loads. An optimization strategy capable of handling discrete decision making, such as Everett optimization, is proposed. We extended Everett method to handle expected congestion margin and reliability index as constraints. To make it effective for real time load dispatch process, a rule-based scheme is presented in the optimization method. It assists in selecting which feeder load to be shed, the location of the load, the value, priority of the load and cost benefit analysis of the load profile is included in the scheme. The scheme is tested using a benchmark NASA system consisting of generators, loads and network.

Momoh, James A.↗

NASA NTRS2015

Advanced Stirling Convertor Dual Convertor Controller Testing at NASA Glenn Research Center in the Radioisotope Power Systems System Integration Laboratory

NASA Glenn Research Center (GRC) developed a non-nuclear representation of a Radioisotope Power System (RPS) consisting of a pair of Advanced Stirling Convertors (ASC), a Dual Convertor Controller (DCC) EM (engineering model) 2 & 3, and associated support equipment, which were tested in the Radioisotope Power Systems System Integration Laboratory (RSIL). The DCC was designed by the Johns Hopkins University/Applied Physics Laboratory (JHU/APL) to actively control a pair of Advanced Stirling Convertors (ASC). The first phase of testing included a Dual Advanced Stirling Convertor Simulator (DASCS) which was developed by JHU/APL and simulates the operation and electrical behavior of a pair of ASC's in real time via a combination of hardware and software. RSIL provides insight into the electrical interactions between a representative radioisotope power generator, its associated control schemes, and realistic electric system loads. The first phase of integration testing included the following spacecraft bus configurations: capacitive, battery, and supercapacitor. A load profile, created based on data from several missions, tested the RPS and RSIL ability to maintain operation during load demands above and below the power provided by the RPS. The integration testing also confirmed the DCC's ability to disconnect from the spacecraft when the bus voltage dipped below 22 V or exceeded 36 V. Once operation was verified with the DASCS, the tests were repeated with actual operating ASC's. The goal of this integration testing was to verify operation of the DCC when connected to a spacecraft and to verify the functionality of the newly designed RSIL. The results of these tests are presented in this paper.

systems integration↗

NASA NTRS2015

Advanced Stirling Convertor Dual Convertor Controller Testing at NASA Glenn Research Center in the Radioisotope Power Systems System Integration Laboratory

NASA Glenn Research Center developed a nonnuclear representation of a Radioisotope Power System (RPS) consisting of a pair of Advanced Stirling Convertors (ASCs), Dual Convertor Controller (DCC) EMs (engineering models) 2 and 3, and associated support equipment, which were tested in the Radioisotope Power Systems System Integration Laboratory (RSIL). The DCC was designed by the Johns Hopkins University Applied Physics Laboratory (JHU/APL) to actively control a pair of ASCs. The first phase of testing included a Dual Advanced Stirling Convertor Simulator (DASCS), which was developed by JHU/APL and simulates the operation and electrical behavior of a pair of ASCs in real time via a combination of hardware and software. RSIL provides insight into the electrical interactions between a representative radioisotope power generator, its associated control schemes, and realistic electric system loads. The first phase of integration testing included the following spacecraft bus configurations: capacitive, battery, and super-capacitor. A load profile, created based on data from several missions, tested the RPS's and RSIL's ability to maintain operation during load demands above and below the power provided by the RPS. The integration testing also confirmed the DCC's ability to disconnect from the spacecraft when the bus voltage dipped below 22 volts or exceeded 36 volts. Once operation was verified with the DASCS, the tests were repeated with actual operating ASCs. The goal of this integration testing was to verify operation of the DCC when connected to a spacecraft and to verify the functionality of the newly designed RSIL. The results of these tests are presented in this paper.

Controllers↗

NASA NTRS2026

Comparison Testing and Analysis of Braided Cords Constructed from Varying Kevlar Fiber Types

Material advancements move quickly, resulting in material obsolescence issues that require parachute engineers to understand how the material properties change with a material upgrade and if there are expected impacts to overall performance of the parachute. The objective of this test was to measure the elongation of different Kevlar types at both the cord-level and at the yarn-level to determine if a change in Kevlar type impacted parachute performance. An Instron machine with a pin-to-pin connection was used to measure load and displacement time histories for various Kevlar cord and yarn samples in a cyclic load profile. Twenty-four cord samples and twenty-four yarn samples were tested; each sample was cycled five times to the relevant maximum load (either to breaking strength or to the 30% peak load expected in flight). In addition to the Instron measurements, a photogrammetric technique was also implemented to measure elongation of the samples using button targets. Modulus test data shows approximately a 10% difference in elongation between Kevlar 29AP and Kevlar 129, and another 10% difference in elongation between Kevlar 29 and Kevlar 29AP. Although this is a measurable change in modulus, the effect on parachute performance is likely small, particularly for non-primary structure elements of relatively small length. This test campaign provides an example of the type of testing and analysis that parachute engineers could do to understand how material upgrades affect performance of their systems.

Parachutes↗

NASA NTRS2026

Comparison Testing and Analysis of Braided Cords Constructed from Varying Kevlar Fiber Types

Modulus Testing↗

NASA NTRS2002

NiH2 Cycle Life Study

Cycle life studies have been performed at Eagle Picher Technologies (EPT), on HST Mantech design cells with various pedigrees of slurry and dry sinter processed electrodes, to evaluate peak load voltage performance during generic load profile testing. These tests provide information for determining voltage and capacity fade (degradation) mechanisms, and their impact on nickel hydrogen cell cycle life. Comparison of peak load voltage fade, as a function of State of Charge and cycle life, with capacity data from HST indicates that the cycle life limiting mechanism is due to impedance growth, and formation of a second discharge plateau. With a second plateau on discharge, capacity from the cell is still available, but at an unacceptable low voltage of 0.8 V per cell (17.6 V battery). Data shows that cell impedance increases with cycle number and depth of discharge, as expected.

Hollandsworth, Roger P.↗

NASA NTRS1990

Influence of linear profile modification and loading conditions on the dynamic tooth load and stress of high contact ratio gears

A computer simulation for the dynamic response of high-contact-ratio spur gear transmissions is presented. High contact ratio gears have the potential to produce lower dynamic tooth loads and minimum root stress but they can be sensitive to tooth profile errors. The analysis presented examines various profile modifications under realistic loading conditions. The effect of these modifications on the dynamic load (force) between mating gear teeth and the dynamic root stress is presented. Since the contact stress is dependent on the dynamic load, minimizing dynamic loads will also minimize contact stresses. It is shown that the combination of profile modification and the applied load (torque) carried by a gear system has a significant influence on gear dynamics. The ideal modification at one value of applied load will not be the best solution for a different load. High-contact-ratio gears were found to require less modification than standard low-contact-ratio gears. High-contact-ratio gears are more adversely affected by excess modification than by under modification. In addition, the optimal profile modification required to minimize the dynamic load (hence the contact stress) on a gear tooth differs from the optimal modification required to minimize the dynamic root (bending) stress. Computer simulation can help find the design tradeoffs to determine the best profile modification to satisfy the conflicting constraints of minimizing both the load and root stress in gears which must operate over a range of applied loads.

Lee, Chinwai↗

NASA NTRS1998

An Adaptable Power System with Software Control Algorithm

A low cost, flexible and modular spacecraft power system design was developed in response to a call for an architecture that could accommodate multiple missions in the small to medium load range. Three upcoming satellites will use this design, with one launch date in 1999 and two in the year 2000. The design consists of modular hardware that can be scaled up or down, without additional cost, to suit missions in the 200 to 600 Watt orbital average load range. The design will be applied to satellite orbits that are circular, polar elliptical and a libration point orbit. Mission unique adaptations are accomplished in software and firmware. In designing this advanced, adaptable power system, the major goals were reduction in weight volume and cost. This power system design represents reductions in weight of 78 percent, volume of 86 percent and cost of 65 percent from previous comparable systems. The efforts to miniaturize the electronics without sacrificing performance has created streamlined power electronics with control functions residing in the system microprocessor. The power system design can handle any battery size up to 50 Amp-hour and any battery technology. The three current implementations will use both nickel cadmium and nickel hydrogen batteries ranging in size from 21 to 50 Amp-hours. Multiple batteries can be used by adding another battery module. Any solar cell technology can be used and various array layouts can be incorporated with no change in Power System Electronics (PSE) hardware. Other features of the design are the standardized interfaces between cards and subsystems and immunity to radiation effects up to 30 krad Total Ionizing Dose (TID) and 35 Mev/cm(exp 2)-kg for Single Event Effects (SEE). The control algorithm for the power system resides in a radiation-hardened microprocessor. A table driven software design allows for flexibility in mission specific requirements. By storing critical power system constants in memory, modifying the system code for other programs is simple. These constants can be altered also by ground command, or in response to an anomolous event. All critical power system functions have backup hardware functions to prevent a software or computer glitch from propagating. A number of battery charge control schemes can be implemented by selecting the proper control terms in the code. The architecture allows the design engineer to tune the system response to various system components and anticipated load profiles without costly alterations. A design trade was made with the size, weight and power dissipation of the electronics versus the performance of the power bus to load variations. Linear, fine control is maintained with a streamlined electronics design. This paper describes the hardware design as well as the software control algorithm. The challenges of closing the system control loop digitally is discussed. Control loop margin and power system performance is presented. Lab measurements are shown and compared to the system response of a hardware model running actual flight software.

Castell, Karen↗

NASA NTRSDate not supplied

Predicting Maximum Temperatures of a Li-ion Battery on a Simulated Flight Profile using a Model-based Prognostics

One of the challenges in using Li-ion packs in aeronautics is their safety, and thermal runaway (TR) is a major concern. The current engineering solutions to prevent a Li-ion pack from a catastrophic TR require additional mass and volume to isolate cells. The excess mass could be reduced by improving detection and, thus, preventing a TR event. One of the possible early warning indicators of a TR is crossing a threshold temperature. We have developed an approach, based on the Unscented Kalman Filter (UKF), to predict the likelihood of reaching the threshold temperature for simulated flight profiles. The current battery prognostics algorithms for aerospace predict state-of-charge (SOC) and end-of-life (EOL) [1]. We extended this two-level algorithm to predict the maximum temperature during discharge. The amount of heat generated in a cell depends on factors such as cell chemistry, cell packaging, total cycles, operating temperature, and abuse history [2]. Our semi-empirical thermal model depends on three phenomenological parameters which account for those factors. In addition, a two-parameter reduced-order model is developed to predict the temperature rise for short bursts of “random-walk” (RW) discharge current sequence, which simulates a flight's current-loading profile. The performance of these models on different datasets and types of current loading will be presented. To predict the maximal temperatures for future cycles we must estimate the evolution of thermal parameters as the batteries age. It is found that the parameters of the 3-parametric thermal model cannot be estimated only from the RW data. To address the issue, we will present two alternative approaches: i) expanding the datasets to include discharge profiles beyond RWs; ii) model reduction to a two-parametric model. The two approaches will be illustrated by an application to the cycling data from a commercial LG 18650 cell. References: 1. M. Daigle, C.S. Kulkarni, End-of-discharge and End-of-life Prediction in Lithium-ion Batteries with Electrochemistry-based Aging Models, in: AIAA Infotech @ Aerospace, American Institute of Aeronautics and Astronautics, San Diego, California, USA, 2016. 2. M. Börner, et. al, Correlation of aging and thermal stability of commercial 18650-type lithium ion batteries, Journal of Power Sources. 342 (2017) 382–392.

Thermal runaway↗

NASA NTRS2004

An Empirical Cumulus Parameterization Scheme for a Global Spectral Model

Realistic vertical heating and drying profiles in a cumulus scheme is important for obtaining accurate weather forecasts. A new empirical cumulus parameterization scheme based on a procedure to improve the vertical distribution of heating and moistening over the tropics is developed. The empirical cumulus parameterization scheme (ECPS) utilizes profiles of Tropical Rainfall Measuring Mission (TRMM) based heating and moistening derived from the European Centre for Medium- Range Weather Forecasts (ECMWF) analysis. A dimension reduction technique through rotated principal component analysis (RPCA) is performed on the vertical profiles of heating (Q1) and drying (Q2) over the convective regions of the tropics, to obtain the dominant modes of variability. Analysis suggests that most of the variance associated with the observed profiles can be explained by retaining the first three modes. The ECPS then applies a statistical approach in which Q1 and Q2 are expressed as a linear combination of the first three dominant principal components which distinctly explain variance in the troposphere as a function of the prevalent large-scale dynamics. The principal component (PC) score which quantifies the contribution of each PC to the corresponding loading profile is estimated through a multiple screening regression method which yields the PC score as a function of the large-scale variables. The profiles of Q1 and Q2 thus obtained are found to match well with the observed profiles. The impact of the ECPS is investigated in a series of short range (1-3 day) prediction experiments using the Florida State University global spectral model (FSUGSM, T126L14). Comparisons between short range ECPS forecasts and those with the modified Kuo scheme show a very marked improvement in the skill in ECPS forecasts. This improvement in the forecast skill with ECPS emphasizes the importance of incorporating realistic vertical distributions of heating and drying in the model cumulus scheme. This also suggests that in the absence of explicit models for convection, the proposed statistical scheme improves the modeling of the vertical distribution of heating and moistening in areas of deep convection.

Rajendran, K.↗

NASA NTRS2012

Evaluating Three Active Thermal Architectures for Exploration Missions

Mass and cost are typically the two biggest challenges facing space craft designers. Active thermal control systems for crewed space-craft are typically among the more massive and costly systems on the vehicle. A study was completed evaluating three different thermal control system architectures to evaluate overall performance, mass and cost for a typical exploration mission profile. The architectures that were evaluated were 1 - a two-loop system using an internal liquid loop interfacing with an external liquid loop and flow loop with flow through radiators; 2 - a-single loop architecture with flow through radiators utilizing a regenerative heat exchanger and heater; and 3 - a single-loop architecture with heat pipe radiators. Environmental conditions, calculated for a given lunar exploration mission, and mission heat load profiles, generated based on previous Orion time lines, were evalauated through the phases of the on orbit mission. Performance for each of the architectures was evaluated along with the resultant mass of each system. Recommendations include adding a thermal topping system to lunar missions due to the extreme hot environments encountered in near-lunar approaches.

Cross, Cynthia D.↗

NASA NTRS1998

Simulation of Crack Propagation in Engine Rotating Components under Variable Amplitude Loading

The crack propagation life of tested specimens has been repeatedly shown to strongly depend on the loading history. Overloads and extended stress holds at temperature can either retard or accelerate the crack growth rate. Therefore, to accurately predict the crack propagation life of an actual component, it is essential to approximate the true loading history. In military rotorcraft engine applications, the loading profile (stress amplitudes, temperature, and number of excursions) can vary significantly depending on the type of mission flown. To accurately assess the durability of a fleet of engines, the crack propagation life distribution of a specific component should account for the variability in the missions performed (proportion of missions flown and sequence). In this report, analytical and experimental studies are described that calibrate/validate the crack propagation prediction capability ]or a disk alloy under variable amplitude loading. A crack closure based model was adopted to analytically predict the load interaction effects. Furthermore, a methodology has been developed to realistically simulate the actual mission mix loading on a fleet of engines over their lifetime. A sequence of missions is randomly selected and the number of repeats of each mission in the sequence is determined assuming a Poisson distributed random variable with a given mean occurrence rate. Multiple realizations of random mission histories are generated in this manner and are used to produce stress, temperature, and time points for fracture mechanics calculations. The result is a cumulative distribution of crack propagation lives for a given, life limiting, component location. This information can be used to determine a safe retirement life or inspection interval for the given location.

Bonacuse, P. J.↗

NASA NTRS1999

Simulation of Crack Propagation in Engine Rotating Components Under Variable Amplitude Loading

The crack propagation life of tested specimens has been repeatedly shown to strongly depend on the loading history. Overloads and extended stress holds at temperature can either retard or accelerate the crack growth rate. Therefore, to accurately predict the crack propagation life of an actual component, it is essential to approximate the true loading history. In military rotorcraft engine applications, the loading profile (stress amplitudes, temperature, and number of excursions) can vary significantly depending on the type of mission flown. To accurately assess the durability of a fleet of engines, the crack propagation life distribution of a specific component should account for the variability in the missions performed (proportion of missions flown and sequence). In this report, analytical and experimental studies are described that calibrate/validate the crack propagation prediction capability for a disk alloy under variable amplitude loading. A crack closure based model was adopted to analytically predict the load interaction effects. Furthermore, a methodology has been developed to realistically simulate the actual mission mix loading on a fleet of engines over their lifetime. A sequence of missions is randomly selected and the number of repeats of each mission in the sequence is determined assuming a Poisson distributed random variable with a given mean occurrence rate. Multiple realizations of random mission histories are generated in this manner and are used to produce stress, temperature, and time points for fracture mechanics calculations. The result is a cumulative distribution of crack propagation lives for a given, life limiting, component location. This information can be used to determine a safe retirement life or inspection interval for the given location.

P J Bonacuse↗

NASA NTRSDate not supplied

Numerical Investigation of the Effect of Trailing Edge Thickness of Simulated CMC Blades on Loss Profiles

To understand such an aerodynamic characteristic of the CMC blades, we numerically simulated three different CMC blades with varying trailing edge thicknesses. We utilized both Large Eddy simulation (LES) with the dynamic subgrid closure as well a recently implemented Reynolds-averaged Navier-Stokes (RANS) coupled with an intermittency function-based transition model. We considered two sets of experimental test points with different Reynolds numbers at a given high-free stream turbulence range (Tu=~10-13%). The LES was previously validated by comparing with experimental data. We validated the RANS transition model against the ERCOFTAC experimental zero-pressure gradient smooth flat plate boundary layer at low- and high freestream turbulence offered by the experimental data of the Stripf’s for a turbine blade. We confirmed that the predicted skin friction and the heat flux agree well with the data. Finally, we computed the predicted pressure loading profiles, total loss distributions, and integrated loss of the three different CMC blades and made comparisons against the experimental data acquired at NASA Glenn Transonic Turbine Blade Cascade Rig. Our paper will present the aforementioned results and discuss the ability of LES and RANS model of our choice to predict salient aerodynamic characteristics of the blades with the various trailing edge thickness effects.

LES↗

NASA NTRSDate not supplied

Numerical Investigation of the Effect of Trailing Edge Thickness of Simulated CMC Blades on Loss Profiles

LES↗