A phase-diverse-phase-retrieval iterative-transform algorithm has been developed that enables high- spatial frequency, high-dynamic-range wavefront sensing. The algorithm has the advantages of both the iterative-transform (high-spatial frequency) and parametric (high dynamic range) phase-recovery techniques. This is achieved by incorporating feedback during phase-recovery to guide the phase-retrieval process.

The innovation consists of a product and process for producing the product which is capable of transfering gas at spacecraft potential to high voltage for use as a propellant for ion thrusters. The device consists of an alumina tube filled with a porous medium that are bonded to itself and the ceramic tube. The isolator is capped at each end to allow spacecraft potential gas (such as xenon) to enter the isolator from propellant tanks and flow control systems and to flow through the isolator to leave a high voltage (hundreds to thousands of volts) without conducting current so that the gas can then be used as propellant for electric propulsion devices such as ion thrusters. The isolator has surfaces suitably configured and a sputter shielding to greatly reduce the probability of developing conductive coatings which would cause performance degradation or potentially thruster failure.

Design of electric field antenna of several meters length for plasma wave instrument which can withstand photon intensities from the Sun EF = 400 watts/cm2 when the spacecraft comes within 3 solar radii of the Sun's photosphere. This design will withstand temperatures T ≈ 2000 °C and minimize the heat input to the spacecraft via radiation and thermal conduction. The antenna will be mounted at the top of the spacecraft bus just below the secondary heat shield of the spacecraft's thermal protection system (TPS). During launch, it will be stowed in a downward position on spring-loaded hinges attached to the top deck of the spacecraft bus and be folded along its axis at several hinge points. When deployed by an appropriate release mechanism, it will extend outward along the cylindrical radius of the spacecraft (see Figure 1) and unfold at the other spring-loaded hinge points with total cylindrical radius of about 3 meters when fully deployed. The first section of the antenna, primarily confined within shadow, may be composed of a special C–C material with heat conduction along antenna axis of K = 5 W/mK to minimize heat conducted to spacecraft bus. Radiators may be mounted at hinge point where it attaches to spacecraft bus to further reduce heat conduction to spacecraft. The other C–C sections could have a high thermal conductivity so there would be no temperature gradients along antenna axis. The cross-section of the antenna, as shown in Figure 2, will have oblique triangular shape to minimize temperature of side of antenna exposed to direct sunlight. The other side is tilted by 17° to ensure that this side is not exposed to direct sunlight. The inside of the antenna is filled with yttria stabilized zirconia (YSZ) ceramic material with low thermal conductivity and high mechanical strength, (see attachment 9.A below for further details)