This paper describes the design decisions taken and the mass properties tracking and testing flow chosen for the Van Allen Probes spacecraft and their deployable systems to achieve the coning angle requirements. Topics include a list of major requirements, a brief description of the error budget, a description of the tracking process of the spacecraft mass properties prior to test, a description of the spin balance and mass properties testing of the spacecraft core and deployable systems, and a presentation of the final mass properties and coning angle calculations of the fully deployed observatories. Launched August 30, 2012, the observed on-orbit, fully deployed configuration coning angles met the requirements, validating the spin balance and mass properties tracking, testing, and calculation methods chosen for the Van Allen Probes mission.

APL has developed a novel miniaturized release mechanism that can be used in CubeSats, NanoSats and miniature space-borne science instruments. This miniature device is inexpensive, reusable, powerefficient, and doesn’t consume any flight parts. The principle of operation for the release mechanism is to use two parts that have complementary coefficients of thermal expansion (CTE). Requirements, key challenges, performance results and lessons learned are presented. Results from a total of 37 test actuations of the design under various environmental conditions demonstrated a robust device that performed reliably in a flight-like environment.

Recent advances in miniaturization techniques and materials have enabled progressively smaller satellites to provide significant mission capability. The cost of launch, however, continues to be a barrier to executing low-cost missions. Capitalizing on significant, otherwise unused launch vehicle (LV) volume and lift mass capability, many developers have found inexpensive launch accommodation as either secondary payloads, or as truly “opportunistic” piggyback/tertiary rideshares. Examples are CubeSats and EELV Secondary Payload Adapter (ESPA) payloads. The success of rideshare missions has given launch integrators confidence that rideshare payloads will not adversely impact primary payloads. As a result of these rideshare successes, there is growing global interest to increase small satellite launch accommodation to quantities well in excess of ten free-flyer deployments per mission. JHU/APL assessed existing capabilities for manifesting small satellites as rideshare payloads on a variety of launch vehicles. Through a combination of top-down analysis and bottoms-up design activities, it was determined that there was an un-served niche between 5-10 kg 3U/6U CubeSats and 180 kg ESPA-class small satellites; accommodation for a space vehicle of 30-50 kg mass and approximately 88,000 cm3 volume was needed. To address this need, JHU/APL has developed a unique flexible adapter system and a corresponding Multi-Mission Nanosatellite (MMN) space vehicle (SV) concept, that can readily integrate with multiple LVs, and more optimally utilizes available mass and volume provisions. In this paper we will discuss the analysis findings, associated requirements, space vehicle provisions and technical details of the adapter system design.

The MErcury Surface, Space ENvironment, GEochemistry, and Ranging (MESSENGER) spacecraft was designed and constructed to withstand the harsh environments associated with achieving and operating in Mercury orbit. The system can be divided into eight subsystems: structures and mechanisms (e.g., the composite core structure, aluminum launch vehicle adapter, and deployables), propulsion (e.g., the state-of-the-art titanium fuel tanks, thruster modules, and associated plumbing), thermal (e.g., the ceramic-cloth sunshade, heaters, and radiators), power (e.g., solar arrays, battery, and controlling electronics), avionics (e.g., the processors, solid-state recorder, and data handling electronics), software (e.g., processor-supported code that performs commanding, data handling, and spacecraft control), guidance and control (e.g., attitude sensors including star cameras and Sun sensors integrated with controllers including reaction wheels), radio frequency telecommunications (e.g., the spacecraft antenna suites and supporting electronics), and payload (e.g., the science instruments and supporting processors). This system architecture went through an extensive (nearly four-year) development and testing effort that provided the team with confidence that all mission goals will be achieved.

This paper presents an example of a new class of planetary exploration missions that is been enabled by the combination of the three technologies of advanced radioisotope power systems, electric propulsion, and expendable launch vehicles. These PARIS (Planetary Access with Radioisotope Ion-drive System) missions are optimized for rendezvous with outer solar system bodies in shallow gravity wells. They are low-thrust missions that are launched to a high C-3 and use their electric propulsion systems to slow them to enable orbit insertion or landing on the target body. The PARIS spacecraft can be powered by traditional Radioisotope Thermoelectric Generators (RTGs), but win benefit greatly from the improved power-to-mass ratio of Stirling radioisotope generators (SRGs) that results from their high conversion efficiency. These New-Frontiers class missions can carry a significant science payload to the Jovian Trojan asteroids. The Trojans are very primitive bodies located near the Jovian L4 and L5 Lagrange points. The PARIS to Hektor mission can reach the asteroids in less than 5 years, orbit 624 Hektor, the largest of the Jovian Trojans, and go on to orbit at least one other nearby object. There are estimated to be more than 10(5) Jovian Trojans greater than 1 km in diameter. The PARIS to Hektor spacecraft has a candidate payload that includes wide-field and narrow-field cameras, a UV-Vis-IR spectrograph, gamma-ray and neutron spectrometers, and plasma and energetic particle spectrometers. The power system generates about 900 W and the launch mass is slightly less than 1000 kg. The trip time is 5 years if "classic" GPHS RTGs are used for power. Next generation Stirling radioisotope generators (SRGs), with a demonstrated thermal conversion efficiency of > 30% and an estimated specific power of > 8W/kg would reduce the travel time to about,4 years.

The Sentinels mission is a key component of NASA's Living With a Star (LWS) program. The Sentinels Science and Technology Definition Team (STDT) has completed a study to define the science objectives, measurement requirements and observational strategies, and mission design for the Sentinels mission. The Inner Heliospheric Sentinels (IHS) are one of the three flight elements (the others are the Near Earth Sentinel and Far Side Sentinel) that make up the Sentinels mission. The four spin-stabilized IHS spacecraft are in elliptical heliocentric orbit with perihelia at ~0.25 AU and aphelia at ~0.75 AU. This orbit presents unique spacecraft thermal control and power challenges. This study has demonstrated mission feasibility by developing a spacecraft design concept using conventional technologies that satisfy the science and mission requirements defined by the Sentinels STDT. Numerous trade studies were performed to optimize the spacecraft design. The trade studies and their results are discussed.

NASA's MESSENGER mission, part of its Discovery program, is the first mission to return to the planet Mercury since the Mariner 10 flybys in 1974 and 1975. The spacecraft incorporates many innovative features, including a sunshade made of ceramic cloth for protection from the Sun, a pair of electronically steerable phased-array antennas, and specially hardened solar panels. A suite of seven miniaturized science instruments, along with the antennas, will globally characterize the planet's composition, structure, atmosphere, and charged particle environment. MESSENGER was launched on August 3, 2004, and performed its single Earth flyby on August 2, 2005. The spacecraft will make two flybys of Venus and three of Mercury prior to orbiting the planet for one Earth-year beginning in March 2011. Highlights of a busy first year of flight operations include initial testing of all spacecraft systems and instruments, execution of six trajectory control maneuvers, and instrument observations of the Earth and Moon surrounding the August flyby

This paper presents the contrasts and similarities between two interplanetary spacecraft structure and propulsion system designs. Pro,sram risk philosophy, schedule and mass requirements all played a significant role in shaping the final configuration of the very different NEAR and MESSENGER structure and propulsion systems. The configuration differences also greatly influenced the integration and testing plans for the two programs.

he Midcourse Space Experiment (MSX) spacecraft represents the largest structure ever assembled at the Applied Physics Laboratory. The design of the structure involves the application of conventional aluminum framework/honeycomb panel construction with composite materials technology to satisfy both the launch vehicle induced loads and the dimensional stability requirements of the optical sensors and instruments. The 1.59-m 2 by 4.11-m-long structure is divided into three discrete elements consisting of (1) an electronics section, which provides an interface to the Delta II launch vehicle and a mounting platform for instrument and spacecraft electronics; (2) a thermally stable graphite/epoxy composite truss structure for mounting the Spatial Infrared Imaging Telescope III (SPIRIT III) instrument and providing an interface between the instrument and electronics sections; and (3) a temperature-controlled instrument section with embedded heat pipes for mounting delicate sensors and instruments. The instrument section also provides a thermally stable mounting structure for the optical bench and the beacon receiver bench. This article describes the features of the MSX structural design, the structural analysis, and testing efforts associated with qualifying the design. A description of the efforts involved in the development of the graphite/epoxy composite optical bench and beacon receiver bench is also presented. T