This paper discusses the major thermal shield subsystem development activities in support of the Solar Probe study being conducted at JPL. The Solar Probe spacecraft will travel to within 4 solar radii of the sun's center to perform fundamental experiments in space physics. Exposure to 2900 earth suns at perihelion requires the spacecraft to be protected within the shadow envelope of a protective shield. In addition, the mass loss rate off of the shield at elevated temperature must comply with plasma instrument requirements and has become the driver of the shield design. This paper will focus on the analytical design work to size the shield and control the shield mass loss rate for the various spacecraft options under study, the application of carbon-carbon materials for shield components, development and preparation of carbon-carbon samples for materials testing, and a materials testing program for carbon-carbon and tungsten alloys to investigate thermal/optical properties, mass loss (carbon-carbon only), material integrity, and high velocity impact behavior.

Solar Probe thermal shield design and testing

The Solar Probe spacecraft will travel to within 4 solar radii of the sun's center while performing a variety of fundamental experiments in space physics. Exposure to 2900 earth suns (400 W/sq cm) at perihelion imposes severe thermal and material demands on a solar shield system designed to protect the payload that will reside within the shield's shadow envelope or umbra. The design of the shield subsystem is a thermal/materials challenge requiring new technology development. While currently in the preproject study phase, anticipating a 1995 project start, shield preliminary design efforts are currently underway. This paper documents the current status of the mission concept, the materials issues, the configuration concept for the shield subsystem, the current configuration studies performed to date, and the required material testing to provide a database to support a design effort required to develop the shield subsystem.

A thermal shield concept for the Solar Probe mission

This paper describes the preliminary design, development, and initial testing of prototype carbon-carbon components of the thermal protection system of the NASA STARPROBE spacecraft. The thermal protection system is designed to limit the spacecraft scientific instrument package to a maximum temperature of 55 C during the time of closest approach to the earth's sun when the peak radiative heating rate reaches 393 w/sq cm. The thermal protection system is comprised of a carbon-carbon thin shell primary shield and two carbon-carbon sandwich construction secondary shields which block reradiation from the primary shield to the spacecraft. These shields are joined and stiffened by carbon-carbon structural members and attached to the spacecraft with eight thin-walled carbon-carbon struts. Prototype parts of the principal components have been fabricated and will be tested in a solar radiative flux environment simulating peak mission values in late 1984.

Lightweight carbon-carbon thermal protection system for STARPROBE

A preliminary design of the thermal protection system for the NASA Solar Probe spacecraft is presented. As presently conceived, the spacecraft will be launched by the Space Shuttle on a Jovian swing-by trajectory and at perihelion approach to three solar radii of the surface of the Earth's sun. The system design satisfies maximum envelope, structural integrity, equipotential, and mass loss/contamination requirements by employing lightweight carbon-carbon emissive shields. The primary shield is a thin shell, 15.5-deg half-angle cone which absorbs direct solar flux at up to 10-deg off-nadir spacecraft pointing angles. Secondary shields of sandwich construction and low thickness-direction thermal conductivity are used to reduce the primary shield infrared radiation to the spacecraft payload.

R. B. Dirling

Papers

Preliminary design of the thermal protection system for solar probe