James A. Nabity

Ultrasoft microwire neural electrodes improve chronic tissue integration

Chronically implanted neural multi-electrode arrays (MEA) are an essential technology for recording electrical signals from neurons and/or modulating neural activity through stimulation. However, current MEAs, regardless of the type, elicit an inflammatory response that ultimately leads to device failure. Traditionally, rigid materials like tungsten and silicon have been employed to interface with the relatively soft neural tissue. The large stiffness mismatch is thought to exacerbate the inflammatory response. In order to minimize the disparity between the device and the brain, we fabricated novel ultrasoft electrodes consisting of elastomers and conducting polymers with mechanical properties much more similar to those of brain tissue than previous neural implants. In this study, these ultrasoft microelectrodes were inserted and released using a stainless steel shuttle with polyethyleneglycol (PEG) glue. The implanted microwires showed functionality in acute neural stimulation. When implanted for 1 or 8weeks, the novel soft implants demonstrated significantly reduced inflammatory tissue response at week 8 compared to tungsten wires of similar dimension and surface chemistry. Furthermore, a higher degree of cell body distortion was found next to the tungsten implants compared to the polymer implants. Our results support the use of these novel ultrasoft electrodes for long term neural implants. STATEMENT OF SIGNIFICANCE One critical challenge to the translation of neural recording/stimulation electrode technology to clinically viable devices for brain computer interface (BCI) or deep brain stimulation (DBS) applications is the chronic degradation of device performance due to the inflammatory tissue reaction. While many hypothesize that soft and flexible devices elicit reduced inflammatory tissue responses, there has yet to be a rigorous comparison between soft and stiff implants. We have developed an ultra-soft microelectrode with Youngs modulus lower than 1MPa, closely mimicking the brain tissue modulus. Here, we present a rigorous histological comparison of this novel ultrasoft electrode and conventional stiff electrode with the same size, shape and surface chemistry, implanted in rat brains for 1-week and 8-weeks. Significant improvement was observed for ultrasoft electrodes, including inflammatory tissue reaction, electrode-tissue integration as well as mechanical disturbance to nearby neurons. A full spectrum of new techniques were developed in this study, from insertion shuttle to in situ sectioning of the microelectrode to automated cell shape analysis, all of which should contribute new methods to the field. Finally, we showed the electrical functionality of the ultrasoft electrode, demonstrating the potential of flexible neural implant devices for future research and clinical use.

A Freezable Heat Exchanger for Space Suit Radiator Systems

During an ExtraVehicular Activity (EVA), both the heat generated by the astronaut s metabolism and that produced by the Portable Life Support System (PLSS) must be rejected to space. The heat sources include the heat of adsorption of metabolic CO2, the heat of condensation of water, the heat removed from the body by the liquid cooling garment and the load from the electrical components. Although the sublimator hardware to reject this load weighs only 1.58 kg (3.48 lbm), an additional 3.6 kg (8 lbm) of water are loaded into the unit, most of which is sublimated and lost to space, thus becoming the single largest expendable during an eight-hour EVA. Using a radiator to reject heat from the astronaut during an EVA can reduce the amount of expendable water consumed in the sublimator. Radiators have no moving parts and are thus highly reliable. Past freezable radiators have been too heavy, but the weight can be greatly reduced by placing a small and freeze tolerant heat exchanger between the astronaut and radiator, instead of making the very large radiator freeze tolerant. Therefore, the key technological innovation to improve space suit radiator performance was the development of a lightweight and freezable heat exchanger that accommodates the variable heat load generated by the astronaut. Herein, we present the heat transfer performance of a newly designed heat exchanger that endured several freeze / thaw cycles without any apparent damage. The heat exchanger was also able to continuously turn down or turn up the heat rejection to follow the variable load.

Additives to Increase Fuel Heat Sink Capacity

Recently, NASA has increased its emphasis on reducing the cost of reaching low Earth orbit . Studies indicate that using a rocket based combined cycle (RBCC) engine could help achieve significant cost reductions. In the third stage of a RBCC launch, the air-breathing vehicle will travel at speeds from Mach 5 to Mach 10 prior to exiting the atmosphere. At these speeds, the heat loads in the combustor are so high that sensible heating of the fuel alone will not be sufficient to meet the cooling demand. However, up to 50% additional heat sink capacity can be extracted from the fuel if it undergoes endothermic, thermal cracking reactions prior to combustion. Unfortunately, the very high temperatures required to achieve necessary cracking rates reduce the allowable stress in the heat exchanger, increasing its weight and reducing its efficiency. Thus, the overall objective of this work is to maximize the heat sink available from hydrocarbon fuels such as JP-7 by adding a chemical initiator, which significantly increases the rates of the thermal cracking reactions. In previous work, TDA developed additives that produced significant increases in the rate of thermal cracking of normal paraffins. However, no studies had been done to measure the effectiveness of these additives with real aviation fuels, such as JP -7, which are complex mixtures of bra nched and cyclo paraffins in addition to normal paraffins. Therefore the goal of this project was to measure the effect of our chemical initiator on the heat sink capacity of JP-7. To accomplish this, we first constructed a test section that could be use d to measure fuel heat sink directly. We then measured the heat sink of JP -7 with and without initiator and, for comparison, conducted similar measurements with a normal paraffin, n-heptane. The results of this project clearly demonstrate that the initiator is very effective at increasing the fuel heat sink capacity of JP-7. In the temperature range between 400 and 550oC, we obtained a heat sink value of 3 03 Btu/lb without initiator. When we added initiator, we obtained a value of 373 Btu/lb over the same temperature range, an improvement of 23%. We also found that the effects of the initiator were greater with a normal paraffin fuel. With nheptane, we obtained a heat sink value of 282 Btu/lb as the fuel is heated from 400 and 575oC without initiator. With initiator, we obtained a value of 420 Btu/lb, an increase of 49%. Finally, a kinetic model suggests that the temperatures required to obtain 50 and 75% n-heptane cracking levels with initiator are close to 100oC lower than the temperatures required without initiator.

Space Suit Radiator Performance in Lunar and Mars Environments

During an ExtraVehicular Activity (EVA), both the heat generated by the astronauts metabolism and that produced by the Portable Life Support System (PLSS) must be rejected to space. The heat sources include the heat of adsorption of metabolic CO2, the heat of condensation of water, the heat removed from the body by the liquid cooling garment and the load from the electrical components. Although the sublimator hardware to reject this load weighs only 3.48 lbs, an additional eight pounds of water are loaded into the unit of which about six to eight are sublimated and lost; this is the single largest expendable during an eight-hour EVA. Using a radiator to reject heat from the Astronaut during an EVA, we can significantly reduce the amount of expendable water consumed by the sublimator. Last year we reported on the design and initial operational assessment tests of our novel radiator designated the Radiator And Freeze Tolerant heat eXchanger (RAFT-X). Herein, we report on tests conducted in the NASA Johnson Space Center Chamber E Thermal Vacuum Test Facility. Up to 800 Btu/h of heat were rejected in lunar and Mars environments with temperatures as cold as 150 F. Tilting the radiator did not cause an observable loss in performance. The RAFT-X endured freeze/thaw cycles and in fact, the heat exchanger was completely frozen three times without any apparent damage to the unit. We were also able to operate the heat exchanger in a partially frozen configuration to throttle the heat rejection rate from 530 Btu/h at low water flow rate down to 300 Btu/h. Finally, the deliberate loss of a single loop heat pipe only degraded the heat rejection performance by about 2 to 5%.

Modeling a Freezable Water-Based Heat Exchanger for Use in Spacecraft Thermal Control

A spacecraft thermal control system typically uses water coolant loops to keep the cabin comfortable. However, the relatively high freeze point of water limits its use, because freezing water within the loop may damage the thermal control system. This paper describes a novel freeze-tolerant heat exchanger and the thermal model developed to correlate data from experiments. The radial transport of heat within the heat exchanger was modeled as a thermal circuit. The heat flows from the water into the thermally conductive shell and fins. One flow channel within the heat exchanger was lined with insulation to keep it free of ice. When the heat sink temperature is below the freeze point of water, a layer of ice builds up in the other flow channels to greatly increase the thermal resistance to heat transport, since the thermal conductivity of ice is only about 1/80th that of Al-6061. In addition, the wetted surface area is reduced by a factor of six when only the insulated channel can still flow water (all of th...

A MEMS Fuel Atomizer for Advanced Engines

Future aircraft and weapons need highly efficient propulsive power plants to achieve their performance goals. Higher performance engines such as the pulse detonation engine (PDE) are needed. The pulse detonation engine is an exciting airbreathing propulsion cycle with great potential for improved range and thrust. Most research has been performed with gaseous fuels, so demonstrating rapid deflagration-to-detonation using storable, liquid hydrocarbon fuels is essential to successful development and application of this engine. Unfortunately, liquid fuels require atomization unless pre-vaporized, which introduces the challenge of detonating a liquid spray. Recent research shows that droplet Sauter-mean diameters as small as 3 µm may be required for successful detonations. Unfortunately, currently available atomizers produce a droplet distribution that has an unacceptably large number of big droplets (up to 200µm) even though the Sautermean diameter can be as small as 30-40µm. Thus, new atomization technologies are needed to produce these small droplets. Using microelectromechanical (MEMS) technology, atomizers can be built with the micron scale features needed to obtain very fine droplet atomization. In this paper we describe the design, fabrication and packaging challenges encountered in building a MEMS fuel atomizer for use in advanced engines.

A Self-Regulating Freezable Heat Exchanger for Use in Spacecraft Thermal Control

A spacecraft thermal control system must keep the vehicle, avionics and atmosphere (if crewed) within a defined temperature range. Water coolant loops are typically used to transport heat to or from the cabin of a crewed spacecraft via heat exchangers to the heat sink systems that reject the heat to space. Water is non-toxic and good for heat transport, but it has a high freeze point. Thus, there is concern that the water loop can freeze and damage the thermal control system unless a low freeze point intermediate fluid loop is included. Incorporating a freeze-tolerant water/ice heat exchanger can eliminate this risk and offers a novel approach to spacecraft thermal control, since parts of the heat exchanger can be selectively frozen to passively increase the turndown of the heat rejection rate. In addition, it has the potential to simplify the thermal control system (for example, a secondary loop between the coolant water loop and the radiator may no longer be needed) and thereby reduce its size and mass.

Simulation of an Electrostatically Driven Microinjector

This paper presents results of analytical and numerical simulations of an electrostatically actuated microfuel injector. The electrostatic-structural coupling, efficiency of valves, and net flow rate of the pump are calculated using analytical relations and compared with numerical simulations. One way to increase flow rate is to increase the stroke volume, which may lead to large deflection of the diaphragm. Thus, the membrane load deflection relationship is no longer linear and when coupled to the inherently nonlinear fluid behavior leads to a complex problem. We used numerical methods to solve for the fluid-structure interaction and the large deflection of the membrane, and studied the effect of actuation frequency on the net flow rate.

Studies of MEMS Colloid Thrusters

In this paper we have carried out numerical calculations and electrospray experiments to understand the underlying physics of colloid thruster operation and support the design of micro fabricated colloid thruster arrays. The calculations show that the Hagen-Poiseuille analytic solution predicts the steady flow behavior of the pressure feed system with reasonable accuracy. Our electrostatic simulations illustrate the importance of extraction electrode design to minimize the startup voltage, to minimize emitter-to-emitter interactions, to maximize thrust and to control dispersion in droplet velocities and therefore, the thrust vector. In addition, the experiments reveal that external wetting of the micro emitter electrode can prevent formation of the electrospray.

Electrostatic Modeling of Colloid Droplet Motion

emitters. Colloid thruster technology continues to be attractive for spacecraft propulsion, since the specific impulse can be many times greater than even the best bi-propellant chemical rocket [1]. However, the technology was abandoned in the 1970’s due to low charge-to-mass ratios, which leads to excessively high voltage requirements and a large inert mass fraction. The recent advent of conductive propellants coupled with MEMS technology has created renewed interest in colloid thrusters for microand nano-satellite applications. Colloid thruster arrays can produce relatively large thrust levels, while maintaining the ability to deliver a small impulse bit. However, difficulties in fabrication and assembly, the need for high voltage addressing and thruster operability issues have hindered full development of the concept at the micro-scale.