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NASA NTRS1993

Ram acceleration from a two phase detonative system

A concept for ram acceleration is presented here, which uses a combination of a gas core and a layer of solid explosive or propellant to generate high thrust densities. The concept can be either self-synchronized or externally synchronized, and may be reusable. It has the potential to achieve very high acceleration rates, higher exit velocities and to lower the tube length requirements. Preliminary numerical simulations are presented and discussed, which show the characteristics of the flow fields. Stable conditions can be achieved for low mass loadings of solid explosive, and relatively slow combustion. Accurate knowledge of the thermo-chemical properties and the equations of state of the gas and solid components is essential for further tuning of the concept.

Cambier, Jean-Luc↗

NASA NTRS1995

Magnetic Annual Arc (MAARC)

Electromagnetic propulsion systems offer high specific impulses (greater than 1500) and high thrust-density as compared to conventional chemical rocket propulsion system. Understanding and simulating the physical processes in electromagnetic propulsion devices (as well as testing with refined diagnostic equipment) hold the key to improved performance and longevity of these devices. This in turn is an essential requirement for a practical application in future space endevors.

Thomas, Herbert↗

NASA NTRS2002

The FRC Acceleration Space Thruster (FAST) Experiment

The objective of the FRC (Field Reversed Configuration) Acceleration Space Thruster (FAST) Experiment is to investigate the use of a repetitive FRC source as a thruster, specifically for an NEP (nuclear electric propulsion) system. The Field Reversed Configuration is a plasmoid with a closed poloidal field line structure, and has been extensively studied as a fusion reactor core. An FRC thruster works by repetitively producing FRCs and accelerating them to high velocity. An FRC thruster should be capable of I(sub sp)'s in the range of 5,000 - 25,000 seconds and efficiencies in the range of 60 - 80 %. In addition, they can have thrust densities as high as 10(exp 6) N/m2, and as they are inductively formed, they do not suffer from electrode erosion. The jet-power should be scalable from the low to the high power regime. The FAST experiment consists of a theta-pinch formation chamber, followed by an acceleration stage. Initially, we will produce and accelerate single FRCs. The initial focus of the experiment will be on the ionization, formation and acceleration of a single plasmoid, so as to determine the likely efficiency and I(sub sp). Subsequently, we will modify the device for repetitive burst-mode operation (5-10 shots). A variety of diagnostics are or will be available for this work, including a HeNe interferometer, high-speed cameras, and a Thomson-scattering system. The status of the experiment will be described.

Martin, Adam↗

NASA NTRS2003

A Plasmoid Thruster for Space Propulsion

There are a number of possible advantages to using accelerated plasmoids for in-space propulsion. A plasmoid is a compact plasma structure with an integral magnetic field. They have been studied extensively in controlled fusion research and are classified according to the relative strength of the poloidal and toroidal magnetic field (BP and Bt, respectively). An Object with B P t >> 1 is classified as a Field Reverse Configuration (FRC); if B, = Bt, it is called a Spheromak. The plasmoid thruster operates by producing FRC-like plasmoids, and subsequently ejecting them from the device at high velocity. The plasmoid is formed inside of a single turn conical theta-pinch coil. As this process is inductive, there are no electrodes. Similar experiments have yielded plasmoid velocities of at least 50 km/s (l), and calculations indicate that velocities in excess of 100 km/s should be possible. This concept should be capable of producing Isp s in the range of 5,000 - 10,000 s with thrust densities of order 10(exp 5) N/sq m. The current experiment is designed to produce jet powers in the range of 5-10 kW, although the concept should be scalable to several MW s. The plasmoids mass and velocity will be measured with a variety of diagnostics, including internal and external B-dot probes, flux loops, Langmuir probes, high-speed cameras, and a laser interferometer. Also of key importance will be measurements of the efficiency and mass utilization. Simulations of the plasmoid thruster using MOQUI, a time dependent MHD code, will be carried out concurrently with experimental testing.

Koelfgen, Syri J.↗

NASA NTRS2003

A Plasmoid Thruster for Space Propulsion

There are a number of possible advantages to using accelerated plasmoids for in-space propulsion. A plasmoid is a compact plasma structure with an integral magnetic field. They have been studied extensively in controlled fusion research and are classified according to the relative strength of the poloidal and toroidal magnetic field (B(sub p), and B(sub t), respectively). An object with B(sub p), / B(sub t) much greater than 1 is classified as a Field Reversed Configuration (FRC); if B(sub p) approximately equal to B(sub t), it is called a Spheromak. The plasmoid thruster operates by producing FRC-like plasmoids and subsequently ejecting them from the device at a high velocity. The plasmoid is formed inside of a single-turn conical theta-pinch coil. As this process is inductive, there are no electrodes. Similar experiments have yielded plasmoid velocities of at least 50 km/s, and calculations indicate that velocities in excess of 100 km/s should be possible. This concept should be capable of producing Isp's in the range of 5,000 - 15,000 s with thrust densities on the order of 10(exp 5) N per square meters. The current experiment is designed to produce jet powers in the range of 5 - 10 kW, although the concept should be scalable to several MW's. The plasmoid mass and velocity will be measured with a variety of diagnostics, including internal and external B-dot probes, flux loops, Langmuir probes, high-speed cameras and a laser interferometer. Also of key importance will be measurements of the efficiency and mass utilization. Simulations of the plasmoid thruster using MOQUI, a time-dependent MHD code, will be carried out concurrently with experimental testing.

Koelfgen, Syri J.↗

NASA NTRS2004

Plasma Propulsion Research at NASA Marshall Space Flight Center

The Propulsion Research Center at NASA Marshall Space Flight Center is pursuing a range of research efforts aimed at identifying and developing new technologies for primary spacecraft propulsion. Efficient high-power electric propulsion (Ep) thrusters are a particular area of emphasis; these would enable the relatively rapid transit of large payloads about the solar system for unmanned or manned science and exploration. Such a mission would make heavy demands on the propulsion system, which may be required to run reliably for several years at a specific impulse approaching 10,OOO s with an efficiency of turning electrical power into jet power of at least 70%. The transit time to a destination scales approximately inversely with the cube root of the specific power, which is the ratio of jet power to power-plant mass. Consequently, reducing a trip time by half requires roughly an eight-fold increase in specific power. Given a renewed NASA commitment to space nuclear power, developing efficient EP thrusters with high jet power (> 100 kW) would seem to provide the most direct means of significantly increasing the specific power and hence reducing trip times. In particular, electromagnetic devices, with their high inherent thrust densities, should be better suited to high power applications than thrusters which depend exclusively on electrostatic forces for propellant acceleration.

Sheehy, Jeffrey A.↗

NASA NTRS2004

Progress on the Plasmoid Thruster Experiment (PTX)

A plasmoid is a compact plasma structure with an integral magnetic field, that may be categorized according to the relative strength of the poloidal and toroidal magnetic field (B(sub p) and B(sub t), respectively). An object with B(sub p)/B(sub t), much much more than 1 is called a Field Reverse Configuration (FRC); if B(sub p) approximately equal to B(sub t), it is called a Spheromak. The thruster operates by repetitively producing plasmoids that are accelerated and ejected at high velocity. As this process is inductive, there are no electrodes. Also, the magnetic structure of the plasmoid should suppress thermal and mass losses to the wall, and improve detachment of the plasma exhaust from the thruster. This concept should be capable of producing an Isp in the range of 5,000 - 10,000 s with thrust densities of order 10(exp 5) N per square meters. The current experiment is designed to produce jet powers in the range of 5-10 kW, although the concept should be scalable into the MW range. In PTX, the plasmoid is formed inside of a single turn conical theta-pinch coil (17.58 cone angle). The coil is driven by a 640 nF, 35 kV capacitor bank, which rings at a frequency of 500 kHz. Previous experiments on PTX were conducted with a static-fill of propellant gas (6% H2 in He), and demonstrated reliable ionization over a pressure range of 40 - 200 mTorr. We are now adding a fast gas-puff valve to load the propellant, and a ringing pre-ionization circuit (f = 5 Mhz) to better control the plasmoid formation. An alternate coil (8.58 cone angle) will also be used, so as to investigate the effect of coil shape on performance. In addition, a variety of propellants will be used, including hydrogen, nitrogen, and argon. The plasmoid mass and velocity will be measured with a variety of diagnostics, including external B-dot probes and flux loops, a high-speed framing camera, and a HeNe laser interferometer. Internal B-dot probes and a quadruple Langmuir probe will provide additional measurements of the plasmoid shape and structure, as well as density, and will be described in a companion paper. The experimental results will be compared to calculations made using a 0-D coil-gun model and also a 2-D time dependent MHD code.

Martin, Adam↗

NASA NTRS2004

The Plasmoid Thruster Experiment (PTX)

A plasmoid is a compact plasma structure with an integral magnetic field, that may be categorized according to the relative strength of the poloidal and toroidal magnetic fields. If B(sup p)/B(sub t) is much greater than 1, it is an FRC; if B(sub p) approximately equals B(sub t), it is a Spheromak. A plasmoid thruster would operate by repetitively producing plasmoids that are accelerated to high velocity. The process is inductive, and the magnetic structure of the plasmoid suppresses thermal and mass losses, and improves detachment of the exhaust. This concept should be capable of producing an Isp in the range of 5,000 - 10,OOO seconds, with high thrust density. PTX is a device designed to study this concept. The plasmoid is formed inside of a single turn conical theta-pinch coil, driven by a 560 nF, 35 kV capacitor bank. Experiments conducted with a static-fill of propellant gas (6% H2 in He) demonstrated reliable ionization over a pressure range of 40 - 200 mTorr. A fast gas-puff valve to inject propellant has since been added, and a ringing pre-ionization circuit to independently control ionization has been tested. Hydrogen, deuterium, argon, and an N2/H2 mixture have been tried as propellants. Measurements of the plasmoid shape, mass, and velocity, using a variety of diagnostics will be presented,

Martin, Adam↗

NASA NTRS2005

The Plasmoid Thruster Experiment (PTX)

A plasmoid is a compact plasma structure with an integral magnetic field, that may be categorized according to the relative strength of the poloidal and toroidal magnetic fields. If B(sub p/B(sub t) much greater than 1 it is an FRC; if B(sub p) approximately equal to B(sub t), it is a Spheromak. A plasmoid thruster would operate by repetitively producing plasmoids that are accelerated to high velocity. The process is inductive, and the magnetic structure of the plasmoid suppresses thermal and mass losses, and improves detachment of the exhaust, This concept should be capable of producing an Isp in the range of 5,000 - l0,000 seconds, with high thrust density. PTX is a device designed to study this concept. The plasmoid is formed inside of a single turn conical theta-pinch coil, driven by a 560 nF, 35 kV capacitor bank. Experiments conducted with a static-fill of propellant gas (6% H2 in He) demonstrated liable ionization over a pressure range of 40 - 200 mTorr. A fast gas-puff valve to inject propellant has since been added, and a ringing preionization circuit to independently control ionization has been tested, hydrogen, deuterium, argon, and an N2 / H2 mixture have been tried as propellants. Measurements of the plasmoid shape, mass, and velocity, using a variety of diagnostics will be presented.

Martin, Adam↗

NASA NTRS2001

Laser High-Cycle Thermal Fatigue of Pulse Detonation Engine Combustor Materials Tested

Pulse detonation engines (PDE's) have received increasing attention for future aerospace propulsion applications. Because the PDE is designed for a high-frequency, intermittent detonation combustion process, extremely high gas temperatures and pressures can be realized under the nearly constant-volume combustion environment. The PDE's can potentially achieve higher thermodynamic cycle efficiency and thrust density in comparison to traditional constant-pressure combustion gas turbine engines (ref. 1). However, the development of these engines requires robust design of the engine components that must endure harsh detonation environments. In particular, the detonation combustor chamber, which is designed to sustain and confine the detonation combustion process, will experience high pressure and temperature pulses with very short durations (refs. 2 and 3). Therefore, it is of great importance to evaluate PDE combustor materials and components under simulated engine temperatures and stress conditions in the laboratory. In this study, a high-cycle thermal fatigue test rig was established at the NASA Glenn Research Center using a 1.5-kW CO2 laser. The high-power laser, operating in the pulsed mode, can be controlled at various pulse energy levels and waveform distributions. The enhanced laser pulses can be used to mimic the time-dependent temperature and pressure waves encountered in a pulsed detonation engine. Under the enhanced laser pulse condition, a maximum 7.5-kW peak power with a duration of approximately 0.1 to 0.2 msec (a spike) can be achieved, followed by a plateau region that has about one-fifth of the maximum power level with several milliseconds duration. The laser thermal fatigue rig has also been developed to adopt flat and rotating tubular specimen configurations for the simulated engine tests. More sophisticated laser optic systems can be used to simulate the spatial distributions of the temperature and shock waves in the engine. Pulse laser high-cycle thermal fatigue behavior has been investigated on a flat Haynes 188 alloy specimen, under the test condition of 30-Hz cycle frequency (33-msec pulse period and 10-msec pulse width including a 0.2-msec pulse spike; ref. 4). Temperature distributions were calculated with one-dimensional finite difference models. The calculations show that that the 0.2-msec pulse spike can cause an additional 40 C temperature fluctuation with an interaction depth of 0.08 mm near the specimen surface region. This temperature swing will be superimposed onto the temperature swing of 80 C that is induced by the 10-msec laser pulse near the 0.53-mm-deep surface interaction region.

Zhu, Dong-Ming↗

NASA NTRS2005

Progress on the Plasmoid Thruster Experiment (PTX)

A plasmoid, also called a compact toroid, is a compact plasma structure with an integral magnetic field, that may be categorized according to the relative strength of the poloidal and toroidal magnetic field (Bp and Bt, respectively). An object with Bp/Bt much greater than 1 is called a Field Reverse Configuration (FRC); if Bp = Bt, it is called a Spheromak. The plasmoid thruster is a pulsed inductive device which operates by repetitively producing plasmoids that are accelerated and ejected at high velocity. As the process is inductive, this thruster avoids the problem of electrode erosion. Also, the magnetic structure of the plasmoid should suppress thermal and mass losses to the wall, and improve detachment of the plasma exhaust from the thruster. This concept should be capable of producing an Isp of 5,000 seconds and greater, with thrust densities of order 10(exp 5) N/sq m. The plasmoid thruster consists chiefly of a conical theta-pinch coil. Propellant is introduced onto a bias magnetic field, produced by an auxiliary coil, and is then pre-ionized, freezing in the magnetic field. The theta-pinch coil is then energized producing a field aligned anti- parallel to the bias field. The reversed field reconnects with the bias field to form the plasmoid. The magnetic pressure of the reversed field accelerates the plasmoid out of the thruster . A series of experiments have been conducted on the PTX device, which consisted of a single turn conical theta-pinch coil, driven by a 560 nF, 35 kV capacitor bank, which rang at a frequency of 500 kHz, and served all three functions required for formation: pre-ionization, bias field loading, and field reversal. Initial ionization was found to occur in an annular region at the exit plane of the coil, and was found to be reproducible with a variety of gases, including H2, D2, Ar, and an H2/N2 mixture (75% / 25%). A fast gas valve for injecting propellant has been tested, as well as a ringing pre-ionization circuit (operating at 5 Mhz) to better control the plasmoid formation. Velocities as high as 20 km/s have been measured for a plasmoid propagating into a static back-fill of neutral gas. Magnetic field and density measurements indicate that the plasmoid retains its structure well into the down-stream exhaust region, and that it is capable of pushing a dense cloud of neutral gas in front of it. Further experiments will focus on the effects of coil geometry and bias-field strength on the performance of the device. A high-power solid-state switching system is being developed in order to replace the spark-gap switch.

Martin, Adam↗

NASA NTRS2007

High-Heat-Flux Cyclic Durability of Thermal and Environmental Barrier Coatings

Advanced ceramic thermal and environmental barrier coatings will play an increasingly important role in future gas turbine engines because of their ability to protect the engine components and further raise engine temperatures. For the supersonic vehicles currently envisioned in the NASA fundamental aeronautics program, advanced gas turbine engines will be used to provide high power density thrust during the extended supersonic flight of the aircraft, while meeting stringent low emission requirements. Advanced ceramic coating systems are critical to the performance, life and durability of the hot-section components of the engine systems. In this work, the laser and burner rig based high-heat-flux testing approaches were developed to investigate the coating cyclic response and failure mechanisms under simulated supersonic long-duration cruise mission. The accelerated coating cracking and delamination mechanism under the engine high-heat-flux, and extended supersonic cruise time conditions will be addressed. A coating life prediction framework may be realized by examining the crack initiation and propagation in conjunction with environmental degradation under high-heat-flux test conditions.

Zhu, Dongming↗

NASA NTRS2008

High Velocity Linear Induction Launcher with Exit-Edge Compensation for Testing of Aerospace Components

Advances in ultra high speed linear induction electromagnetic launchers over the past decade have focused on magnetic compensation of the exit and entry-edge transient flux wave to produce efficient and compact linear electric machinery. The paper discusses two approaches to edge compensation in long-stator induction catapults with typical end speeds of 150 to 1,500 m/s. In classical linear induction machines, the exit-edge effect is manifest as two auxiliary traveling waves that produce a magnetic drag on the projectile and a loss of magnetic flux over the main surface of the machine. In the new design for the Stator Compensated Induction Machine (SCIM) high velocity launcher, the exit-edge effect is nulled by a dual wavelength machine or alternately the airgap flux is peaked at a location prior to the exit edge. A four (4) stage LIM catapult is presently being constructed for 180 m/s end speed operation using double-sided longitudinal flux machines. Advanced exit and entry edge compensation is being used to maximize system efficiency, and minimize stray heating of the reaction armature. Each stage will output approximately 60 kN of force and produce over 500 G s of acceleration on the armature. The advantage of this design is there is no ablation to the projectile and no sliding contacts, allowing repeated firing of the launcher without maintenance of any sort. The paper shows results of a parametric study for 500 m/s and 1,500 m/s linear induction launchers incorporating two of the latest compensation techniques for an air-core stator primary and an iron-core primary winding. Typical thrust densities for these machines are in the range of 150 kN/sq.m. to 225 kN/sq.m. and these compete favorably with permanent magnet linear synchronous machines. The operational advantages of the high speed SCIM launcher are shown by eliminating the need for pole-angle position sensors as would be required by synchronous systems. The stator power factor is also improved.

Kuznetsov, Stephen↗

NASA NTRS2015

Testing of an Arcjet Thruster with Capability of Direct-Drive Operation

Electric thrusters typically require a power processing unit (PPU) to convert the spacecraft provided power to the voltage-current that a thruster needs for operation. Testing has been initiated to study whether an arcjet thruster can be operated directly with the power produced by solar arrays without any additional conversion. Elimination of the PPU significantly reduces system-level complexity of the propulsion system, and lowers developmental cost and risk. The work aims to identify and address technical questions related to power conditioning and noise suppression in the system and heating of the thruster in long-duration operation. The apparatus under investigation has a target power level from 400-1,000 W. However, the proposed direct-drive arcjet is potentially a highly scalable concept, applicable to solar-electric spacecraft with up to 100's of kW and beyond. A direct-drive electric propulsion system would be comprised of a thruster that operates with the power supplied directly from the power source (typically solar arrays) with no further power conditioning needed between those two components. Arcjet thrusters are electric propulsion devices, with the power supplied as a high current at low voltage; of all the different types of electric thruster, they are best suited for direct drive from solar arrays. One advantage of an arcjet over Hall or gridded ion thrusters is that for comparable power the arcjet is a much smaller device and can provide more thrust and orders of magnitude higher thrust density (approximately 1-10 N/sq m), albeit at lower I(sub sp) (approximately 800-1000 s). In addition, arcjets are capable of operating on a wide range of propellant options, having been demonstrated on H2, ammonia, N2, Ar, Kr, Xe, while present SOA Hall and ion thrusters are primarily limited to Xe propellant. Direct-drive is often discussed in terms of Hall thrusters, but they require 250-300 V for operation, which is difficult even with high-voltage solar arrays. The arcjet requires under 100 V, which is more in-line with what is easily possible with a solar array. Direct-drive of an electric propulsion system confers the advantage of reducing or eliminating the power processing unit (PPU) that is typically needed to convert the spacecraft-provided power to the voltage and current needed for thruster operation. Since the PPU is typically the most expensive piece of an electric thruster system, from both a fabrication and qualification standpoint, its elimination offers the potential for major reductions in system cost and risk. The design of the arcjet built for this effort was based on previous low power (1 kW class) arcjets. It has a precision machined 99.95% pure tungsten anode which also serves as the nozzle. The anode constrictor region is 1 mm (0.040-in) diameter and 1 mm (0.040-in) long. The cathode is a tungsten welding electrode doped with LaO2; its tip was precision ground to a 30◦ angle ending in a blunt end. The two electrodes are separated by a boron-nitride insulator which also serves as the propellant injection manifold; it ends in six small holes which introduce the propellant gas in the diverging section of the nozzle, directly adjacent to the cathode. The electrodes and insulator are housed in a stainless-steel outer-body, with a Macor insulator at the mid-plane to provide thermal isolation between the front and back halves of the device. The gas seals were made using Grafoil gaskets. Figure 1A shows the assembled thruster in the vacuum chamber; figure 1B shows the thruster in operation on argon at a flow rate of 676 sccm (20 mg/s). Initial testing was conducted in a 3.5-ft diameter vacuum chamber; the ultimate pressure reached during quasi-steady operation of the thruster was about 330 millitorr. The thruster was powered with a high-current, 0-100A, 15 kW power supply. The discharge was initiated with a high-voltage (approximately 10 kV) spark initiator that was isolated from the supply by a stack of diodes. The testing indicated that an operating point exists within the I-V characteristics that is compatible with direct-drive solar-electric operation; for a flow rate of 20 mg/s (argon) the arc could be sustained at a voltage of about 20 V and a current of 25 A (500W).

Martin, Adam K.↗

NASA NTRS2015

Robust, Flexible and Lightweight Dielectric Barrier Discharge Actuators Using Nanofoams/Aerogels

Robust, flexible, lightweight, low profile enhanced performance dielectric barrier discharge actuators (plasma actuators) based on aerogels/nanofoams with controlled pore size and size distribution as well as pore shape. The plasma actuators offer high body force as well as high force to weight ratios (thrust density). The flexibility and mechanical robustness of the actuators allows them to be shaped to conform to the surface to which they are applied. Carbon nanotube (CNT) based electrodes serve to further decrease the weight and profile of the actuators while maintaining flexibility while insulating nano-inclusions in the matrix enable tailoring of the mechanical properties. Such actuators are required for flow control in aeronautics and moving machinery such as wind turbines, noise abatement in landing gear and rotary wing aircraft and other applications.

Sauti, Godfrey↗

NASA NTRS1973

Full scale hover test of a 25 foot tilt rotor

The tilt rotor underwent a hover performance test on the Aero Propulsion Laboratory whirl stand at Wright-Patterson Air Force Base. The maximum thrust over density ratio measured at the design tip speed of 740 feet per second was 10,016 pounds. This occurred when the power over density ratio was 1721 horsepower. At the hover overspeed rpm, the thrust and power, over density ratio, were 11,008 pounds and 1866 horsepower. During the test, the maximum measured thrust coefficient was 0.177, and the rotor figure of merit exceeded 0.81. Measured lifting efficiency was 8.35 pounds per horsepower at the thrust a 13,000-pound aircraft would require for hover at sea level on a standard day. No effect of compressibility on performance is discernible in the test results (the range of tip Mach numbers tested was 0.55 to 0.71).

Helf, S.↗

NASA NTRS1999

High-Energy Space Propulsion Based on Magnetized Target Fusion

A conceptual study is made to explore the feasibility of applying magnetized target fusion (MTF) to space propulsion for omniplanetary travel. Plasma-jet driven MTF not only is highly amenable to space propulsion, but also has a number of very attractive features for this application: 1) The pulsed fusion scheme provides in situ a very dense hydrogenous liner capable of moderating the neutrons, converting more than 97% of the neutron energy into charged particle energy of the fusion plasma available for propulsion. 2) The fusion yield per pulse can be maintained at an attractively low level (< 1 GJ) despite a respectable gain in excess of 70. A compact, low-weight engine is the result. An engine with a jet power of 25 GW, a thrust of 66 kN, and a specific impulse of 77,000 s, can be achieved with an overall engine mass of about 41 metric tons, with a specific power density of 605 kW/kg, and a specific thrust density of 1.6 N/kg. The engine is rep-rated at 40 Hz to provide this power and thrust level. At a practical rep-rate limit of 200 Hz, the engine can deliver 128 GW jet power and 340 kN of thrust, at specific power and thrust density of 1,141 kW/kg and 3 N/kg respectively. 3) It is possible to operate the magnetic nozzle as a magnetic flux compression generator in this scheme, while attaining a high nozzle efficiency of 80% in converting the spherically radial momentum of the fusion plasma to an axial impulse. 4) A small fraction of the electrical energy generated from the flux compression is used directly to recharge the capacitor bank and other energy storage equipment, without the use of a highvoltage DC power supply. A separate electrical generator is not necessary. 5) Due to the simplicity of the electrical circuit and the components, involving mainly inductors, capacitors, and plasma guns, which are connected directly to each other without any intermediate equipment, a high rep-rate (with a maximum of 200 Hz) appears practicable. 6) All fusion related components are within the current state of the art for pulsed power technology. Experimental facilities with the required pulsed power capabilities already exist. 7) The scheme does not require prefabricated fuel target and liner hardware in any esoteric form or state. All necessary fuel and liner material are introduced into the engine in the form of ordinary matter in gaseous state at room temperature, greatly simplifying their handling on board. They are delivered into the fusion reaction chamber in a completely standoff manner.

Thio, Y. C. F.↗

DOE OSTI2025

Freestream turbulence effects on unsteady wind turbine loads and wakes: An IDDES study

We investigate numerically the effects of freestream turbulence on the unsteady aerodynamics and wakes of the National Renewable Energy Laboratory Phase VI wind turbine rotor for increasing wind speed. Turbulence is modeled using the Improved Delayed Detached-Eddy Simulation (IDDES) method. As a first step, a detailed mesh resolution study is conducted with the decaying freestream turbulence model at turbulence intensity of 0.5%. Our blade-resolved IDDES simulations show that grid-independent average torque and thrust results can be achieved with relatively coarse meshes, whereas dramatically higher mesh resolution is required for grid-independent results for power spectral densities of thrust force, especially in the deep-stall regime. Comparing the loads with the Shear-Stress Transport model demonstrates the superiority of IDDES in predicting massively separated flows. The aerodynamic performance and wake predictions with the decaying freestream turbulence model are compared with the synthetic freestream turbulence model. Both models predict nearly the same loads, spectral energy content, and wake characteristics. The properties of both the near- and far-wake regions are then examined. Furthermore, we show that separated boundary layers accelerate turbulent mixing and entrainment of the external flow, which results in faster wake recovery. The effect of increasing turbulence intensity to 6% is investigated using the synthetic freestream turbulence model. In contrast with the fully attached boundary layer, higher freestream turbulence in deep stall does not significantly affect the loads and vortex-shedding characteristics. However, the turbulent mixing in the wake is enhanced, which further hastens the recovery of the self-similar velocity profile. In general, increasing the wind speed at high turbulence intensity shifts the recovery farther upstream and increases the wake width.

17 WIND ENERGY↗