James V. McAdams

The MESSENGER mission to Mercury: spacecraft and mission design

Abstract A Mercury orbiter mission is challenging from thermal and mass perspectives. The Mercury Surface, Space Environment, Geochemistry, and Ranging (MESSENGER) mission overcomes these challenges while avoiding esoteric technologies by using an innovative approach with commonly available materials, minimal moving parts, and maximum heritage. This approach yields a spacecraft with good margins in all categories and low technical risk. The key concepts are a ceramic-cloth sunshade, an integrated lightweight structure and high- performance propulsion system, and a solar array incorporating optical solar reflectors (OSRs). The sunshade maintains the spacecraft at room temperature. The integrated structure and propulsion system provides ample mass margin. The solar array with OSRs, which has already undergone significant testing, provides thermal margin even if the panels are inadvertently pointed directly at the Sun at 0.3 AU. 0.3 AU .

Trajectory Design and Maneuver Strategy for the MESSENGER Mission to Mercury

Destined to become the first spacecraft to orbit the planet Mercury, the MESSENGER spacecraft was launched on 3 August 2004. The 6.6-year ballistic trajectory to Mercury will utilize six gravity-assist flybys of Earth (one), Venus (two), and Mercury (three). With three trajectory correction maneuvers completed by mid-December 2005, many more maneuvers will be necessary during the journey to Mercury and the subsequent 1-year duration Mercury orbit phase. The spacecrafts design and operational capability will enable real-time monitoring of every course-correction maneuver. A complex mission plan will provide multiple opportunities to obtain observational data that will help fulfill the missions scientific objectives. Soon after entering Mercury orbit in mid-March 2011, the initial primary science orbit will have an 80-deg orbit inclination relative to Mercurys equator, 200-km periapsis altitude, 60°N subspacecraft periapsis latitude, and a 12-h orbit period. With science goals requiring infrequent orbit-phase trajectory adjustments, pairs of orbit-correction maneuvers occur at about the same time every Mercury year, or every 88 days. For the first time, the spacecrafts orbit design at Mercury accounts for the best available Mercury gravity model, small solar pressure perturbations due to changes in the solar array tilt angle, and an improved strategy for performing orbit correction maneuvers.

A realistic interstellar explorer

For more than 20 years, an “Interstellar Precursor Mission” to ∼1000 AU within the working lifetime of the initiators (<50 years) has been discussed as a high priority for multiple scientific objectives. During the last two years there has been renewed interest in actually sending a probe to another star system - a “grand challenge” for NASA - and the idea of a precursor mission has been renewed as a beginning step to achieve this goal. We revisit an old idea for implementing such a mission. The probe is launched initially to Jupiter and then falls to the Sun where a large propulsive ΔV maneuver propels it on a high-energy ballistic escape trajectory from the solar system. The implementation requires a low-mass, highly-integrated spacecraft to make use of moderate (Delta-class) expendable launch vehicles. We provide a first-order cut at many of the engineering realities associated with such a probe. We identify a mission concept that can link the required science, desired instruments, spacecraft engineeri...

Mission to the sun: The Solar Pioneer

Abstract The Solar Pioneer is a mission concept developed by the Johns Hopkins University Applied Physics Laboratory (JHU/APL) to do exploratory science in the inner heliosphere and outer solar corona. The concept is derived from the Near Earth Asteroid Rendezvous (NEAR) spacecraft now being built by JHU/APL for NASA and draws upon that design to the maximum extent possible. The primary purpose, to conduct an exploratory in situ basic science investigation, has been focused to emphasize in situ particles and fields measurements of the outer solar corona and inner solar wind region. The purpose of the Solar Pioneer is to deliver a payload of scientific instruments to approximately 4 solar radii. The basic characteristics of the JHU/APL baseline concept provide for delivery of 30 to 50 kg of scientific payload to 4 R s while maintaining a high data rate. Life-cycle cost control is obtained by: (1) focusing on prime science objectives, (2) applying advanced technology where it makes sense, and (3) capitalizing on developed subsystems derived from the NEAR spacecraft. By implementing a focused strategy from science goals through all stages of program management, such a mission can be carried out for less than one quarter of previous Solar Probe mission cost estimates.

Io Volcano Observer's (IVO) integrated approach to optimizing system design for radiation challenges

One of the major challenges for a mission to the Jovian system is the radiation tolerance of the spacecraft (S/C) and the payload. Moreover, being able to achieve science observations with high signal to noise ratios (SNR), while passing through the high flux radiation zones, requires additional ingenuity on the part of the instrument provider. Consequently, the radiation mitigation is closely intertwined with the payload, spacecraft and trajectory design, and requires a systems-level approach. This paper presents a design for the Io Volcano Observer (IVO), a Discovery mission concept that makes multiple close encounters with Io while orbiting Jupiter. The mission aims to answer key outstanding questions about Io, especially the nature of its intense active volcanism and the internal processes that drive it. The payload includes narrow-angle and wide-angle cameras (NAC and WAC), dual fluxgate magnetometers (FGM), a thermal mapper (ThM), dual ion and neutral mass spectrometers (INMS), and dual plasma ion analyzers (PIA). The radiation mitigation is implemented by drawing upon experiences from designs and studies for missions such as the Radiation Belt Storm Probes (RBSP) and Jupiter Europa Orbiter (JEO). At the core of the radiation mitigation is IVOs inclined and highly elliptical orbit, which leads to rapid passes through the most intense radiation near Io, minimizing the total ionizing dose (177 krads behind 100 mils of Aluminum with radiation design margin (RDM) of 2 after 7 encounters). The payload and the spacecraft are designed specifically to accommodate the fast flyby velocities (e.g. the spacecraft is radioisotope powered, remaining small and agile without any flexible appendages). The science instruments, which collect the majority of the high-priority data when close to Io and thus near the peak flux, also have to mitigate transient noise in their detectors. The cameras use a combination of shielding and CMOS detectors with extremely fast readout to minimize noise. INMS microchannel plate detectors and PIA channel electron multipliers require additional shielding. The FGM is not sensitive to noise induced by energetic particles and the ThM microbolometer detector is nearly insensitive. Detailed SNR calculations are presented. To facilitate targeting agility, all of the spacecraft components are shielded separately since this approach is more mass efficient than using a radiation vault. IVO uses proven radiation-hardened parts (rated at 100 krad behind equivalent shielding of 280 mils of Aluminum with RDM of 2) and is expected to have ample mass margin to increase shielding if needed.

Maneuver Design Strategy Enables Precise Targeting of the First MESSENGER Mercury Flyby

On 14 January 2008 MESSENGER (MErcury Surface, Space ENvironment, GEochemistry, and Ranging) completed the first Mercury flyby since Mariner 10 did so in March 1975. This paper will describe how post-launch modifications to MESSENGER’s maneuver design process contributed to highly accurate delivery near the encounter B-plane aim point for the second Venus and first Mercury flybys. Maneuver performance for completed trajectory-correction maneuvers and requirements for upcoming maneuvers and planetary flybys will further demonstrate maneuver design and trajectory optimization improvements made since launch. Also included are the effects of all six planetary gravity assists on the spacecraft trajectory.

Contingency Plans for MESSENGER's Mercury Orbit Insertion Maneuver

MESSENGER (MErcury Surface, Space ENvironment, GEochemistry, and Ranging) will be the first spacecraft to orbit the planet Mercury when it begins its one-year Mercury orbital mission phase next year. On 18 March 2011 MESSENGER will perform the critical 862 m/s Mercury orbit insertion (MOI) maneuver. This paper summarizes strategies for recovering MESSENGER’s science mission in the event of an aborted or anomalous MOI maneuver. If 70% or more of the MOI burn is completed, MESSENGER will be captured into a high Mercury orbit. One or two maneuvers would then be required to achieve the planned 82.5inclination, 12.0-hour orbit, and all science objectives can be met for most of these cases. If less than 70% of the MOI burn is completed, MESSENGER would remain in a heliocentric orbit, and a recovery maneuver must occur either 10 to 14 days after the 18 March 2011 MOI attempt, or approximately one Mercury year (87.97 days) later in June 2011. For these heliocentric trajectories, solutions were found by which the spacecraft returns to Mercury after either one Mercury year or multiple (Earth) years (subsequent to completing one more or one less revolution of the Sun than Mercury). None of the successful return solutions exceeds the 7-year maximum preferred return time to Mercury.