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Search indexed NASA NTRS and DOE OSTI research on propulsion, heat transfer, battery materials and energy systems. Follow report and document links to the original sources.

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At least 235 records · Page 13
All sourcesNASA NTRSDOE OSTI
NASA NTRS1956

A Theory for Stability and Buzz Pulsation Amplitude in Ram Jets and an Experimental Investigation Including Scale Effects

From a theory developed on a quasi-one-dimensional-flow basis, it is found that the stability of the ram jet is dependent upon the instantaneous values of mass flow and total pressure recovery of the supersonic diffuser and immediate neighboring subsonic diffuser. Conditions for stable and unstable flow are presented. The theory developed in the report is in agreement with the experimental data of NACA-TN-3506 and NACA-RM-L50K30. A simple theory for predicting the approximate amplitude of small pressure pulsation in terms of mass-flow decrement from minimum-stable mass flow is developed and found to agree with experiments. Cold-flow tests at a Mach number of 1.94 of ram-jet models having scale factors of 3.15:1 and Reynolds number ratios of 4.75:1 with several supersonic diffuser configurations showed only small variations in performance between geometrically similar models. The predominant variation in steady-flow performance resulted from the larger boundary layer in the combustion chamber of the low Reynolds number models. The conditions at which buzz originated were nearly the same for the same supersonic diffuser (cowling-position angle) configurations in both large and small diameter models. There was no appreciable variation in stability limits of any of the models when the combustion-chamber length was increased by a factor of three. The unsteady-flow performance and wave patterns were also similar when considered on a reduced-frequency basis determined from the relative lengths of the model. The negligible effect of Reynolds number on stability of the off-design configurations was not anticipated in view of the importance of boundary layer to stability, and this result should not be construed to be generally applicable. (author)

Trimpi, Robert L↗
NASA NTRS1995

He II Liquid/Vapor Phase Separator for Large Dynamic Range Operation

A phase separator, which separates helium vapor from liquid superfluid helium (He II), is an indispensable device for space cryogenics. The most recent approach to the Space Infrared Telescope Facility (SIRTF) uses a new design concept in which only the detector package is cold at launch, the remainder of the telescope being subsequently cooled to operating temperature on orbit. Therefore, a large dynamic operational range is required of the cryogen system. This is a report of initial laboratory test results with candidate porous plugs as phase separators. Mass flow rates and pressure and temperature differences across a porous plug were measured in this experiment. Relatively large mass flow rates were observed even at small pressure differences. In the high mass flow rate region, a hysteresis was observed with increases and decreases of the pressure difference. A linear theory is proposed and compared with experimental data to explain several phenomena observed in this system.

Phase Separator, Space Cryogenics↗
NASA NTRS2016

Pressure Loss Predictions of the Reactor Simulator Subsystem at NASA Glenn Research Center

Testing of the Fission Power System (FPS) Technology Demonstration Unit (TDU) is being conducted at NASA Glenn Research Center. The TDU consists of three subsystems: the reactor simulator (RxSim), the Stirling Power Conversion Unit (PCU), and the heat exchanger manifold (HXM). An annular linear induction pump (ALIP) is used to drive the working fluid. A preliminary version of the TDU system (which excludes the PCU for now) is referred to as the "RxSim subsystem" and was used to conduct flow tests in Vacuum Facility 6 (VF 6). In parallel, a computational model of the RxSim subsystem was created based on the computer-aided-design (CAD) model and was used to predict loop pressure losses over a range of mass flows. This was done to assess the ability of the pump to meet the design intent mass flow demand. Measured data indicates that the pump can produce 2.333 kg/sec of flow, which is enough to supply the RxSim subsystem with a nominal flow of 1.75 kg/sec. Computational predictions indicated that the pump could provide 2.157 kg/sec (using the Spalart-Allmaras (S‒A) turbulence model) and 2.223 kg/sec (using the k- turbulence model). The computational error of the predictions for the available mass flow is ‒0.176 kg/sec (with the S-A turbulence model) and -0.110 kg/sec (with the k- turbulence model) when compared to measured data.

Mass Flow↗
NASA NTRS2004

Effect of Detonation through a Turbine Stage

Pulse detonation engines (PDE) have been investigated as a more efficient means of propulsion due to its constant volume combustion rather than the more often used constant pressure combustion of other propulsion systems. It has been proposed that a hybrid PDE-gas turbine engine would be a feasible means of improving the efficiency of the typical constant pressure combustion gas turbine cycle. In this proposed system, multiple pulse detonation tubes would replace the conventional combustor. Also, some of the compressor stages may be removed due to the pressure rise gained across the detonation wave. The benefits of higher thermal efficiency and reduced compressor size may come at a cost. The first question that arises is the unsteadiness in the flow created by the pulse detonation tubes. A constant pressure combustor has the advantage of supplying a steady and large mass flow rate. The use of the pulse detonation tubes will create an unsteady mass flow which will have currently unknown effects on the turbine located downstream of the combustor. Using multiple pulse detonation tubes will hopefully improve the unsteadiness. The interaction between the turbine and the shock waves exiting the tubes will also have an unknown effect. Noise levels are also a concern with this hybrid system. These unknown effects are being investigated using TURBO, an unsteady turbomachinery flow simulation code developed at Mississippi State University. A baseline case corresponding to a system using a constant pressure combustor with the same mass flow rate achieved with the pulse detonation hybrid system will be investigated first.

Ellis, Matthew T.↗
NASA NTRS2002

Tip-Clearance Vortex Characterized With Three-Dimensional Digital Particle Image Velocimetry

An optical measurement technique known as Three-Dimensional Digital Particle Image Velocimetry (3-D DPIV) was used to characterize the tip clearance flow in NASA Glenn Research Center's low-speed axial compressor. 3-D DPIV is a technique in which a stereoscopic imaging system consisting of two cross-correlation cameras is used to record particles entrained in a flow as a laser light sheet is pulsed at two instances in time. Although 3-D DPIV has been used elsewhere, this is the first time it has been used to measure compressor tip clearance flows. In-house modifications of the DPIV system include the use of effective seeding technology and a novel system to perform a priori calibrations at all five measurement planes, greatly reducing facility run time. Computational fluid dynamics predictions, which are used to guide design changes toward improving the efficiency and operating range of turbomachinery, can be verified and improved by comparison with 3-D DPIV measurements of the actual tip clearance flow. This measurement campaign dealt with the characterization of the tip clearance vortex in the first stage of a four-stage axial compressor. The tip clearance vortex is formed in compressors operating with a clearance gap between the moving rotor blade tips and the stationary casing when a leakage flow, forced from the pressure side of the blade over the blade tip, forms a vortical structure on the suction side of the blade. 3-D DPIV is ideally suited to measure the clearance vortex for two reasons: (1) this technique captures the entire blade passage flow at one instant in time, so that wandering of the vortex during the measurement does not smear out velocity gradients in the flow field, and (2) the spanwise component of velocity changes sign across the vortex core, providing a more accurate measurement of the vortex location than was available with previous two-dimensional measurement approaches. These two attributes of the data will enable computational fluid dynamics researchers to validate their predictions to a level of accuracy not previously attainable. In an effort to understand the effects of this flow phenomenon on the operation of the low speed axial compressor, data were acquired at two mass flow coefficients: 0.395 (design operating point) and 0.35 (operating point just above stall). In order to identify the spatial extent, location, and magnitude of the tip clearance vortex for each mass flow condition, data were acquired at five equally spaced spanwise locations, from 90- to 100-percent span. The data presented shows a cross section of the tip clearance vortex at 94-percent span and a mass flow coefficient of 0.395, where the color contours represent the radial velocity component. The blue (flow towards the hub) and red (flow towards the casing) regions illustrate the extent of the tip clearance vortex. The interface between the blue and red regions delineates the location of the vortex core.

John, W. Trevor↗
NASA NTRS1988

Particle-laden weakly swirling free jets: Measurements and predictions

A theoretical and experimental investigation of particle-laden, weakly swirling, turbulent free jets was conducted. Glass particles, having a Sauter mean diameter of 39 microns, with a standard deviation of 15 microns, were used. A single loading ratio (the mass flow rate of particles per unit mass flow rate of air) of 0.2 was used in the experiments. Measurements are reported for three swirl numbers, ranging from 0 to 0.33. The measurements included mean and fluctuating velocities of both phases, and particle mass flux distributions. Measurements were also completed for single-phase non-swirling and swirling jets, as baselines. Measurements were compared with predictions from three types of multiphase flow analysis, as follows: (1) locally homogeneous flow (LHF) where slip between the phases was neglected; (2) deterministic separated flow (DSF), where slip was considered but effects of turbulence/particle interactions were neglected; and (3) stochastic separated flow (SSF), where effects of both interphase slip and turbulence/particle interactions were considered using random sampling for turbulence properties in conjunction with random-walk computations for particle motion. Single-phase weakly swirling jets were considered first. Predictions using a standard k-epsilon turbulence model, as well as two versions modified to account for effects of streamline curvature, were compared with measurements. Predictions using a streamline curvature modification based on the flux Richardson number gave better agreement with measurements for the single-phase swirling jets than the standard k-epsilon model. For the particle-laden jets, the LHF and DSF models did not provide very satisfactory predictions. The LHF model generally overestimated the rate of decay of particle mean axial and angular velocities with streamwise distance, and predicted particle mass fluxes also showed poor agreement with measurements, due to the assumption of no-slip between phases. The DSF model also performed quite poorly for predictions of particle mass flux because turbulent dispersion of the particles was neglected. The SSF model, which accounts for both particle inertia and turbulent dispersion of the particles, yielded reasonably good predictions throughout the flow field for the particle-laden jets.

Bulzan, Daniel L.↗
NASA NTRS1952

Experimental Investigation of a Two-dimensional Split-wing Ram-jet Inlet at Mach Number of 3.85

Performance characteristics of a two-dimensional isentropic diffuser have been experimentally determined at a Mach number of 3.85. At zero angle of attack, a maximum total-pressure recovery of 0.41 was obtained with a supercritical mass-flow ratio of 0.95. As a consequence of the twin-duct arrangement of the diffuser, a large discontinuity in pressure recovery and mass flow with a characteristic hysteresis was encountered between critical and subcritical operation. An asymmetric shock pattern with large-scale separation and flow reversal in one of the passages occurred at reduced mass flows. Pressure and force data presented for an angle-of-attack range from zero to 4 degrees.

DIFFUSERS, SUPERSONIC↗
NASA NTRS2012

Experimental Studies of Low-Pressure Turbine Flows and Flow Control. Streamwise Pressure Profiles and Velocity Profiles

This report summarizes research performed in support of the NASA Glenn Research Center (GRC) Low-Pressure Turbine (LPT) Flow Physics Program. The work was performed experimentally at the U.S. Naval Academy faculties. The geometry corresponded to "Pak B" LPT airfoil. The test section simulated LPT flow in a passage. Three experimental studies were performed: (a) Boundary layer measurements for ten baseline cases under high and low freestream turbulence conditions at five Reynolds numbers of 25,000, 50,000, 100,000, 200,000, and 300,000, based on passage exit velocity and suction surface wetted length; (b) Passive flow control studies with three thicknesses of two-dimensional bars, and two heights of three-dimensional circular cylinders with different spanwise separations, at same flow conditions as the 10 baseline cases; (c) Active flow control with oscillating synthetic (zero net mass flow) vortex generator jets, for one case with low freestream turbulence and a low Reynolds number of 25,000. The Passive flow control was successful at controlling the separation problem at low Reynolds numbers, with varying degrees of success from case to case and varying levels of impact at higher Reynolds numbers. The active flow control successfully eliminated the large separation problem for the low Reynolds number case. Very detailed data was acquired using hot-wire anemometry, including single and two velocity components, integral boundary layer quantities, turbulence statistics and spectra, turbulent shear stresses and their spectra, and intermittency, documenting transition, separation and reattachment. Models were constructed to correlate the results. The report includes a summary of the work performed and reprints of the publications describing the various studies.This report summarizes research performed in support of the NASA Glenn Research Center (GRC) Low-Pressure Turbine (LPT) Flow Physics Program. The work was performed experimentally at the U.S. Naval Academy faculties. The geometry corresponded to "Pak B" LPT airfoil. The test section simulated LPT flow in a passage. Three experimental studies were performed: (a) Boundary layer measurements for ten baseline cases under high and low freestream turbulence conditions at five Reynolds numbers of 25,000, 50,000, 100,000, 200,000, and 300,000, based on passage exit velocity and suction surface wetted length; (b) Passive flow control studies with three thicknesses of two-dimensional bars, and two heights of three-dimensional circular cylinders with different spanwise separations, at same flow conditions as the 10 baseline cases; (c) Active flow control with oscillating synthetic (zero net mass flow) vortex generator jets, for one case with low freestream turbulence and a low Reynolds number of 25,000. The Passive flow control was successful at controlling the separation problem at low Reynolds numbers, with varying degrees of success from case to case and varying levels of impact at higher Reynolds numbers. The active flow control successfully eliminated the large separation problem for the low Reynolds number case. Very detailed data was acquired using hot-wire anemometry, including single and two velocity components, integral boundary layer quantities, turbulence statistics and spectra, turbulent shear stresses and their spectra, and intermittency, documenting transition, separation and reattachment. Models were constructed to correlate the results. The report includes a summary of the work performed and reprints of the publications describing the various studies. The folders in this supplement contain processed data in ASCII format. Streamwise pressure profiles and velocity profiles are included. The velocity profiles were acquired using single sensor and cross sensor hot-wire probes which were traversed from the wall to the freestream at various streamwise locations. In some of the flow control cases (3D Trips and Jets) profiles were acquired at multiple spanwise locations.

Volino, Ralph↗
NASA NTRS1992

High power arcjet

In this period a new mass flow controller was brought into the gas supply system, so that the upper limit for the mass flow rate could be increased up to 500 mg/s with hydrogen. A maximum specific impulse of 1500 s could be achieved with the high powered arcjet (HIPARC) at an efficiency of slightly better than 20 percent. Different nozzle throat diameters had been tested. The 100 kilo-watt input power limit was reached with the 4 mm nozzle throat diameter at a mass flow rate of 400 mg/s. Tests were carried out with different cathode gaps and with three different cathodes. In addition measurements of pressure and gas temperature were taken in the feed line in order to determine the pressure drop in the propellant injectors.

Goelz, T. M.↗
NASA NTRS2004

1998 Calibration of the Mach 4.7 and Mach 6 Arc-Heated Scramjet Test Facility Nozzles

A calibration of the Arc-Heated Scramjet Test Facility (AHSTF) Mach 4.7 and Mach 6 nozzles was performed in 1998. For each nozzle, three different typical facility operating test points were selected for calibration. Each survey consisted of measurements, at 340 separate locations across the 11 inch square nozzle exit plane, of pitot pressure, static pressure, and total temperature. Measurement density was higher (4/inch) in the boundary layer near the nozzle wall than in the core nozzle flow (1/inch). The results generated for each of these calibration surveys were contour plots at the nozzle exit plane of the measured and calculated flow properties which completely defined the thermodynamic state of the nozzle exit flow. An area integration of the mass flux at the nozzle exit for each survey was compared to the AHSTF mass flow meter results to provide an indication of the overall quality of the calibration performed. The percent difference between the integrated nozzle exit mass flow and the flow meter ranged from 0.0 to 1.3 percent for the six surveys. Finally, a comparison of this 1998 calibration was made with the 1986 calibration. Differences of less than 10 percent were found within the nozzle core flow while in the boundary layer differences on the order of 20 percent were quite common.

Witte, David W.↗
NASA NTRS1994

A Low-Erosion Starting Technique for High-Performance Arcjets

The NASA arcjet program is currently sponsoring development of high specific impulse thrusters for next generation geosynchronous communications satellites (2 kW-class) and low-power arcjets for power limited spacecraft (approx. 0.5 kW-class). Performance goals in both of these efforts will require up to 1000 starts at propellant mass flow rates significantly below those used in state-of-the-art arcjet thruster systems (i.e., high specific power levels). Reductions in mass flow rate can lead to damaging modes of operation, particularly at thruster ignition. During the starting sequence, the gas dynamic force due to low propellant flow is often insufficient to rapidly push the arc anode attachment to its steady-state position in the diverging section of the nozzle. This paper describes the development and demonstration of a technique which provides for non-damaging starts at low steady-state flow rates. The technique employs a brief propellant pressure pulse at ignition to increase gas dynamic forces during the critical ignition/transition phase of operation. Starting characteristics obtained using both pressure-pulsed and conventional starting techniques were compared across a wide range of propellant flow rates. The pressure-pulsed starting technique provided reliable starts at mass flow rates down to 21 mg/s, typically required for 700 s specific impulse level operation of 2 kW thrusters. Following the comparison, a 600 start test was performed across a wide flow rate range. Post-test inspection showed minimal erosion of critical arcjet anode/nozzle surfaces.

Sankovic, John M.↗
NASA NTRS1991

Estimation of additive forces and moments for supersonic inlets

A technique for estimating the additive forces and moments associated with supersonic, external compression inlets as a function of mass flow ratio has been developed. The technique makes use of a low order supersonic paneling method for calculating minimum additive forces at maximum mass flow conditions. A linear relationship between the minimum additive forces and the maximum values for fully blocked flow is employed to obtain the additive forces at a specified mass flow ratio. The method is applicable to two-dimensional inlets at zero or nonzero angle of attack, and to axisymmetric inlets at zero angle of attack. Comparisons with limited available additive drag data indicate fair to good agreement.

Perkins, Stanley C., Jr.↗
NASA NTRS2018

Design of Distributed Mail-Slot Propulsion System on a Hybrid Wingbody Aircraft

This paper deals with designing a thrust distribution strategy when a Turboelectric Distributed Propulsion (TeDP) system of 16 embedded propulsors is installed on an aerodynamically optimized hybrid wing-body configuration. This HWB previously designed to satisfy conditions of trim, longitudinally static stability and specific cargo space is employed as the baseline configuration for the current study of seeking an optimal propulsion/power system. According to the nature of the entrance flow condition for each distributed propulsion passage in hybrid wing-body aircraft, the ingested boundary layer thickness differs and results in different propulsive reaction. An optimal distribution of thrust and power output is determined by how the system utilizes the propulsive characteristics of each passage. The design space and the number of design variables are selected and described accordingly. An actuator disk model is employed to model thrust generation and shaft power from the propulsor. To carry out the optimization of the propulsion/power system on a computationally expensive CFD model, a Kriging method in conjunction with a Genetic Algorithm (GA) is applied. Throughout the design process, the propulsion performances of the sampled propulsion/power system are analyzed and compared to those of a clean flow engine. The performance metrics includes mass flow rate, fan pressure ratio besides the thrust and shaft power. Minimization of total shaft power from the distributed engine is performed at multiple thrust levels. The benefit of boundary layer ingestion propulsion system is quantified via comparison of thrust equivalent, shaft power and mass flow equivalent clean flow engines with CFD based system design.

Hybrid Wingbody↗
NASA NTRS1987

Microprocessor-Based Valved Controller

New controller simpler, more precise, and lighter than predecessors. Mass-flow controller compensates for changing supply pressure and temperature such as occurs when gas-supply tank becomes depleted. By periodically updating calculation of mass-flow rate, controller determines correct new position for valve and keeps mass-flow rate nearly constant.

Norman, Arnold M., Jr.↗
DOE OSTI2022

Development of an MC&A toolbox for liquid-fueled molten salt reactors with online reprocessing (Final Report)

A critical barrier to the deployment of MSRs is the absence of a well-defined nuclear material control and accounting (MC&A) approach, a vital prerequisite to meet NRC licensing requirements as well as facilitating future international exports. Liquid-fueled MSR variants (especially those incorporating online refueling or reprocessing) present a unique set of challenges to traditional nuclear material control and accountancy. Unlike solid-fueled light-water reactors, traditional item-counting methods cannot be applied as an accountancy strategy. Rather, MC&A approaches to MSR variants (including both uranium and thorium fueled designs) are more analogous to bulk material handling facilities (e.g., enrichment and reprocessing); yet further complicating matters is the fact that fuel medium is also highly radioactive. Meanwhile, the space of MSRs covers a broad range of design parameters, including thermal and fast spectra designs; operation in actinide breeder or burner modes, choice of actinide fuel used (e.g., 235 U, 232 Th / 233 U, denatured 233 U, etc.), pool or loop-type configuration, and even different salt chemistry. Each of these design choices introduce significant challenges to MC&A approaches within MSR facilities. We propose to bridge this gap for liquid-fueled MSRs by developing a modular, component-based test framework for evaluating viable process monitoring and MC&A techniques specifically suited to liquid-fueled MSR system variants employing online refueling or reprocessing. This test platform will consist of a toolbox of independent process modules representing discrete physical units (such as the reactor core, off-gas processing, decay tanks, and actinide separation units), each with its own self-contained physics responsive to the input mass flow, along with appropriate measurement models that can be coupled to key flow points. These dynamic physical signatures thus afford the ability to test the viability and efficacy of potential accountancy techniques under the full range of reactor operating conditions. As process modules are connected via mass flows, the result is a reconfigurable, generic MSR mass flow model capable of serving as an MC&A test platform for a broad spectrum of possible MSR configurations. The resulting MSR MC&A toolbox will enable robust assessment of accountancy strategies for this unique facility type, including analysis of physical feedbacks arising both from depletion of the fuel over time as well as from potential off-normal events, including those introduced by equipment failures (e.g., a pump failure) as well as by malicious action (i.e., attempts to divert material). The proposed toolbox both addresses a critical needs area for the MPACT analysis toolkit while leveraging existing MPACT-sponsored tools, especially with respect to simulation of measurement and accountancy techniques for advanced fuel cycle facilities. Beyond enabling new analysis capabilities for MSR systems, the design of this toolbox will be to such to maximize compatibility with existing MPACT tools, such to enhance existing facility MC&A analysis capabilities.

11 NUCLEAR FUEL CYCLE AND FUEL MATERIALS↗
NASA NTRS2015

Simplified Methodology to Estimate the Maximum Liquid Helium (LHe) Cryostat Pressure from a Vacuum Jacket Failure

The aircraft-based Stratospheric Observatory for Infrared Astronomy (SOFIA) is a platform for multiple infrared astronomical observation experiments. These experiments carry sensors cooled to liquid helium temperatures. The liquid helium supply is contained in large (i.e., 10 liters or more) vacuum-insulated dewars. Should the dewar vacuum insulation fail, the inrushing air will condense and freeze on the dewar wall, resulting in a large heat flux on the dewar's contents. The heat flux results in a rise in pressure and the actuation of the dewar pressure relief system. A previous NASA Engineering and Safety Center (NESC) assessment provided recommendations for the wall heat flux that would be expected from a loss of vacuum and detailed an appropriate method to use in calculating the maximum pressure that would occur in a loss of vacuum event. This method involved building a detailed supercritical helium compressible flow thermal/fluid model of the vent stack and exercising the model over the appropriate range of parameters. The experimenters designing science instruments for SOFIA are not experts in compressible supercritical flows and do not generally have access to the thermal/fluid modeling packages that are required to build detailed models of the vent stacks. Therefore, the SOFIA Program engaged the NESC to develop a simplified methodology to estimate the maximum pressure in a liquid helium dewar after the loss of vacuum insulation. The method would allow the university-based science instrument development teams to conservatively determine the cryostat's vent neck sizing during preliminary design of new SOFIA Science Instruments. This report details the development of the simplified method, the method itself, and the limits of its applicability. The simplified methodology provides an estimate of the dewar pressure after a loss of vacuum insulation that can be used for the initial design of the liquid helium dewar vent stacks. However, since it is not an exact tool, final verification of the dewar pressure vessel design requires a complete, detailed real fluid compressible flow model of the vent stack. The wall heat flux resulting from a loss of vacuum insulation increases the dewar pressure, which actuates the pressure relief mechanism and results in high-speed flow through the dewar vent stack. At high pressures, the flow can be choked at the vent stack inlet, at the exit, or at an intermediate transition or restriction. During previous SOFIA analyses, it was observed that there was generally a readily identifiable section of the vent stack that would limit the flow – e.g., a small diameter entrance or an orifice. It was also found that when the supercritical helium was approximated as an ideal gas at the dewar condition, the calculated mass flow rate based on choking at the limiting entrance or transition was less than the mass flow rate calculated using the detailed real fluid model2. Using this lower mass flow rate would yield a conservative prediction of the dewar’s wall heat flux capability. The simplified method of the current work was developed by building on this observation.

Ungar, Eugene K.↗
DOE OSTI2024

Off-design operation and performance of pumped thermal energy storage

In this article, we describe off-design models and control strategies for a Pumped Thermal Energy Storage (PTES) system that uses liquid thermal energy storage: specifically molten salt for hot storage and methanol for cold storage. Off-design conditions arise when load-following, or due to variations in storage tank temperatures or ambient temperatures. We propose a control strategy that uses inventory control to manage the mass flow rate in the thermodynamic cycles, which facilitates load following. We also propose a control strategy for the storage fluid mass flow rates, which are varied to ensure the molten salt is maintained at its design temperature. This maximizes efficiency and minimizes problems with salt freezing or degradation. The cold storage fluid mass flow rate is varied so that the cold tanks have the same state-of-charge as the hot tanks. This leads to variations in cold fluid temperature, but these variations are shown to be acceptably small (e.g. 7.5% increase), and this control method is shown to be simpler and more efficient than an alternative strategy where tanks become unbalanced. The ambient temperature and storage tank temperatures are moved ±50 °C from the design values and the impact on power, duration, and tank temperatures is quantified. Results demonstrate that the proposed control strategy is stable and self-correcting - that is, storage temperatures converge on stable values after two-to-three charge-discharge cycles. When inputs return to design values, the system returns to its design point after two charge-discharge cycles. We also demonstrate that inventory control enables delivery of the target power output even when off-design conditions exist that would normally reduce the power output.

25 ENERGY STORAGE↗