The Deep Atmosphere Venus Investigation of Noble gases, Chemistry, and Imaging (DAVINCI) mission was selected in June 2021 as part of the NASA Discovery Program to explore Venus through remote sensing, in-situ chemistry measurements, and near-surface imaging. This mission will investigate the evolution of Venus' atmosphere using an architecture designed to optimize science-relevant measurements within the atmosphere and on targeted flybys. Currently slated to launch in June 2029, DAVINCI's Carrier-Relay-Imaging Spacecraft (CRIS) will conduct Flybys of Venus in January and November of 2030 where it will acquire remote sensing of the upper Venus atmosphere and surface. In June 2031 during a third Flyby, the Descent Sphere (DS) containing a suite of instruments will transect the atmosphere while acquiring chemistry and environmental measurements, as well as descent imaging of a highland region. The focus of this paper is the technical mission design, concept of operations and flight system design including the DS and CRIS, with mission overview and science summarized in Garvin and others (2020, 2021). The DAVINCI concept has been in development at NASA Goddard Space Flight Center (GSFC) since 2011 with substantial hardware fabrication and test. The project is managed by NASA GSFC, which provides the DS and two key instruments, and Lockheed Martin will provide CRIS and flight operations. Other instruments and major components are provided by NASA Jet Propulsion Laboratory, Johns Hopkins Applied Physics Laboratory, Malin Space Science Systems, KinetX, and other partners. Together with other planned Venus missions, DAVINCI will usher in a new phase of Venus scientific understanding via the first in situ measurements since 1985.

The Deep Atmosphere Venus Investigation of Noble Gases, Chemistry, and Imaging Plus (DAVINCI+) mission is one of four finalists now in Phase A study as part of the ongoing Discovery Program competition [1-5]. If selected for flight, DAVINCI+ will be the first mission to Venus to incorporate flybys, a descent probe, and an orbital phase into one unified architecture – at the cost of a Discovery mission. The result will be a transformative new understanding of the atmosphere, surface, and evolutionary path of Venus as a once-habitable planet [6] (and model exoplanet) that is now host to a certainly unhabitable surface environment.

This paper describes a method for modeling microwave propagation and signal attenuation along an electromagnetic ray linking a spacecraft with a descent probe within the atmosphere of Venus. Attenuation in the Venusian atmosphere is governed by energy absorption and defocusing effects along the path of propagation. Empirical models of the relevant atmospheric properties affecting these are used to determine the curved ray path by way of an efficient numerical integration procedure, coupled with a fast converging iterative algorithm. Emphasis is placed on S-band frequencies, but the method and formulations are applicable up to about 10GHz.

On September 8, 2016, the third NASA New Frontiers mission launched on an Atlas V 411. The Origins, Spectral Interpretation, Resource Identification, Security-Regolith Explorer (OSIRIS-REx) will rendezvous with asteroid Bennu in 2018, collect a sample in 2020, and return that sample to Earth in September 2023. The development team has overcome a number of challenges in order to design and build a system that will make contact with an unexplored, airless, low-gravity body. This paper will provide an overview of the mission, then focus in on the system-level challenges and some of the key system-level processes. Some of the lessons here are unique to the type of mission, like discussion of operating at a largely-unknown, low-gravity object. Other lessons, particularly from the build phase, have broad implications. The OSIRIS-REx risk management process was particularly effective in achieving an on-time and under-budget development effort. The systematic requirements management and verification and the system validation also helped identify numerous potential problems. The final assessment of the OSIRIS-REx performance will need to wait until the sample is returned in 2023, but this post-launch assessment will capture some of the key systems-engineering lessons from the development team.

This paper addresses the technical aspects of the sample return system for the upcoming Origins, Spectral Interpretation, Resource Identification, and Security-Regolith Explorer (OSIRIS-REx) asteroid sample return mission. The overall mission design and current implementation are presented as an overview to establish a context for the technical description of the reentry and landing segment of the mission.

In September of 2016, the OSIRIS-REx (Origins, Spectral Interpretation, Resource Identification, Security, Regolith EXplorer) spacecraft will depart for asteroid (101955) Bennu, and when it does, humanity will turn an important corner in the exploration of the Solar System. After arriving at the asteroid in the Fall of 2018, it will undertake a program of observations designed to select a site suitable for retrieving a sample that will be returned to the Earth in 2023. The third mission in NASA's New Frontiers program, OSIRIS-REx will obtain a minimum of 60 g of a primitive asteroid's surface, the largest sample of extra-terrestrial material returned to the Earth since the end of the Apollo lunar missions (Figure 1). OSIRIS-REx will also return a separate sample of the fine-grained surface material that is <;1 mm in diameter.

Schedule pressure is common in the commercial world, where late delivery of a product means delayed income and loss of profit.12 Research spacecraft developed by NASA, on the other hand, tend to be driven by the high cost of launch vehicles and the public scrutiny of failure—the primary driver is ensuring proper operation in space for a system that cannot be retrieved for repair. The Lunar Reconnaissance Orbiter (LRO) development faced both schedule pressure and high visibility. The team had to balance the strong push to meet a launch date against the need to ensure that this first mission for Exploration succeeded. This paper will provide an overview of the mission from concept through its first year of operation and explore some of the challenges the systems engineering team faced taking a mission from preliminary design review to pre-ship review in 3 years.

Launched June 18, 2009, with its primary mission scheduled to end September 2010, NASA's Lunar Reconnaissance Orbiter will be the first observatory ever to spend an entire year orbiting and observing the Moon at a low altitude of just 50 km. The spacecraft carries a wide variety of scientific instruments and will provide an extraordinary opportunity to study the lunar landscape at resolutions and over time scales never achieved before. This paper is intended as a companion to the series of papers released simultaneously in this journal detailing LRO's instruments and their planned measurements. The paper describes the design and key performance drivers of the LRO spacecraft and overall mission design. It presents a comprehensive description of the operation of the various systems that comprise the spacecraft and illustrates how these systems enable achievement of the mission requirements.

This paper describes the approach taken for the reliability analyses, including probabilistic risk assessments (PRAs), of the Lunar Reconnaissance Orbiter (LRO) that have been performed concurrently with the designs to identify problem areas and ensure corrective actions were taken in a timely manner. The LRO is a large system with more than 14,000 piece-parts and over 120 purchased or contractor-built components. These analyses formulate the probability of failure for each of the components of the system. PRA then integrates these analytical techniques and results to assess the potential for failure and to help find ways to reduce the mission risks. By utilizing the traditional and innovative methods to perform reliability analyses, such as reliability predictions, failure modes and effects analysis (FMEA), and fault tree analysis (FTA), the LRO project office, with close collaboration with the reliability team, has improved the probability of mission success, from 0.72 to 0.81, within a tight schedule and limited budget. As more data from the manufacturers and GSFC designs become available, the reliability and system teams will continue to update the analyses to assess risks throughout the entire lifecycle of the mission

The Global Precipitation Measurement (GPM) program is an international partnership led by the National Aeronautics and Space Administration (NASA) and the Japan Aerospace Exploration Agency (JAXA). GPM will improve climate, weather, and hydrometeorological forecasts through more frequent and more accurate measurement of precipitation across the globe. This paper describes the concept and the preparations for ground validation within the GPM program. Ground validation (GV) plays a critical role in the program by investigating and quantitatively assessing the errors within the satellite retrievals. These quantitative estimates of retrieval errors will assist the scientific community by bounding the errors within their research products. The two fundamental requirements of the GPM ground validation program are: (1) error characterization of the precipitation retrievals and (2) continual improvement of the satellite retrieval algorithms. These two driving requirements determine the measurements, instrumentation, and location for ground observations. This paper describes GV plans for estimating the systematic and random components of retrieval error and for characterizing the spatial and temporal structure of the error. This paper describes the GPM program for algorithm improvement in which error models are developed and experimentally explored to uncover the physical causes of errors within the retrievals. GPM will ensure that information gained through ground validation is applied to future improvements in the space-borne retrieval algorithms. This paper discusses the potential locations for validation measurement and research, the anticipated contributions of GPM's international partners, and the interaction of ground validation with other GPM program elements.

In this paper, we have dealt with the validation and error characterization for the Global Precipitation Measurement (GPM) research initiative. The GPM is a three-year on-orbit duration program with a five-year duration goal. The Core satellite launch is scheduled tentatively for Fall 2008. Presently, GPM is in formulation stage in which the team is developing concepts and requirements prior to design work. As a part of formulation, ground validation is developing its requirements with a top level schedule requirement of commencing GV operations two years prior to the Core satellite launch. The rationale is that GV will benefit from a two-year head start in preparation for the Core observations. The requirements for GV are being developed in collaboration with, and vetted by, the precipitation science community.

This paper introduces plans for ground validation (GV) for the Global Precipitation Measurement. At NASA's request, a Ground Validation Working Group, formed from the meteorological and hydrological communities, is recommending plans to guide the GV program. Ground validation efforts will commence as early as 2003 with the Spring 2003 Pilot Experiment and described herein. The Pilot Experiment is focused on mitigating engineering and scientific risk to the GPM program and, in particular, to the ground validation program.

This paper summarizes the GPM-Core coverage trade space analysis. The goal of this analysis was to determine the GPM-Core sensitivity to changes in altitude and inclination for the three onboard instruments: the radiometer, the KU band radar and the KA band radar. This study will enable a better choice of the nominal GPM-Core orbit as well as the optimal size of the maintenance box (+/-1 km, +/-5 km..). For this work, we used two different figures-of-merit: (1) the time required to cover 100% of the +/-65 deg latitude band and (2) the coverage obtained for a given propagation time (7 days and 30 days). The first figure-of-merit is used for the radiometer as it has a sensor cone half-angle between 3 to 5 times bigger than the radars. Thus, we anticipate that for this instrument the period of the orbit (i.e. altitude) will be the main driver and that the 100% coverage value will be reached within less than a week. The second figure-of-merit is used for the radar instruments as they have small sensor cone half-angle and will, in some cases, never reach the 100% coverage threshold point.

The Global Precipitation Measurement (GPM) is an international effort led by the National Aeronautics and Space Administration (NASA) of the U.S.A. and the National Space Development Agency of Japan (NASDA) for the purpose of improving research into the global water and energy cycle. GPM will improve climate, weather, and hydrological forecasts through more frequent and more accurate measurement of precipitation world-wide. Comprised of U.S. domestic and international partners, GPM will incorporate and assimilate data streams from many spacecraft with varied orbital characteristics and instrument capabilities. Two of the satellites will be provided directly by GPM, the core satellite and a constellation member. The core satellite, at the heart of GPM, is scheduled for launch in November 2007. The core will carry a conical scanning microwave radiometer, the GPM Microwave Imager (GMI), and a two-frequency cross-track-scanning radar, the Dual-frequency Precipitation Radar (DPR). The passive microwave channels and the two radar frequencies of the core are carefully chosen for investigating the varying character of precipitation over ocean and land, and from the tropics to the high-latitudes. The DPR will enable microphysical characterization and three-dimensional profiling of precipitation. The GPM-provided constellation spacecraft will carry a GMI radiometer identical to that on the core spacecraft. This paper presents calibration plans for the GPM, including on-board instrument calibration, external calibration methods, and the role of ground validation. Particular emphasis is on plans for inter-satellite calibration of the GPM constellation. With its Unique instrument capabilities, the core spacecraft will serve as a calibration transfer standard to the GPM constellation. In particular the Dual-frequency Precipitation Radar aboard the core will check the accuracy of retrievals from the GMI radiometer and will enable improvement of the radiometer retrievals. Observational intersections of the core with the constellation spacecraft are essential in applying this technique to the member satellites. Information from core spacecraft retrievals during intersection events will be transferred to the constellation radiometer instruments in the form of improved calibration and, with experience, improved radiometric algorithms. In preparation for the transfer standard technique, comparisons using the Tropical Rainfall Measuring Mission (TRMM) with sun-synchronous radiometers have been conducted. Ongoing research involves study of critical variables in the inter-comparison, such as correlation with spatial-temporal separation of intersection events, frequency of intersection events, variable azimuth look angles, and variable resolution cells for the various sensors.

Following a formal and competitive process, the Wide-Field Infrared Explorer (WIRE) instrument was selected in 1994 to become a participant in NASA/Goddard Space Flight Center's Small Explorer Program (SMEX). The WIRE instrument will be used to conduct a deep infrared, extra galactic science survey 500 times more sensitive than the Infrared Astronomy Satellite (IRAS) Faint Source Catalog. The WIRE instrument includes a cryogenically cooled (<7.5 K), 128/spl times/128 pixel, long wave, infrared detector embedded within an optical assembly which is cooled by a two-stage, solid-hydrogen cryostat. The WIRE spacecraft is being configured to be placed into a 470/spl times/540 km sun-synchronous orbit using an Orbital Sciences Corporation Pegasus XL launch vehicle. Total spacecraft mass will be less than 275 kg. Orbit average power consumed will be less than 175 W. The WIRE instrument will be delivered to NASA/Goddard Space Flight Center in early 1998 to undergo integration and test with the WIRE spacecraft bus and subsequently will be launched in the fall of 1998. The SMEX program uses a strict design-to-cost approach to manage and to contain overall mission cost. This paper will present the major instrument, operational and cost requirements driving the spacecraft systems design for the mechanical, structural, thermal, attitude control, command and data handling, power and electrical systems. It will document major system trade studies results and the subsequent spacecraft design to meet mission requirements.