Risk Evaluation in the Pre-Phase A Conceptual Design of Spacecraft

Typically, the most important decisions in the design of a spacecraft are made in the earliest stages of its conceptual design the Pre-Phase A stages. It is in these stages that the greatest number of design alternatives is considered, and the greatest number of alternatives is rejected. The focus of Pre-Phase A conceptual development is on the evaluation and comparison of whole concepts and the larger-scale systems comprising those concepts. This comparison typically uses general Figures of Merit (FOMs) to quantify the comparative benefits of designs and alternative design features. Along with mass, performance, and cost, risk should be one of the major FOMs in evaluating design decisions during the conceptual design phases. However, risk is often given inadequate consideration in conceptual design practice. The reasons frequently given for this lack of attention to risk include: inadequate mission definition, lack of rigorous design requirements in early concept phases, lack of fidelity in risk assessment methods, and under-evaluation of risk as a viable FOM for design evaluation. In this paper, the role of risk evaluation in early conceptual design is discussed. The various requirements of a viable risk evaluation tool at the Pre-Phase A level are considered in light of the needs of a typical spacecraft design study. A technique for risk identification and evaluation is presented. The application of the risk identification and evaluation approach to the conceptual design process is discussed. Finally, a computational tool for risk profiling is presented and applied to assess the risk for an existing Pre-Phase A proposal. The resulting profile is compared to the risks identified for the proposal by other means

Benefits of Using a Mars Forward Strategy for Lunar Surface Systems

This paper identifies potential risk reduction, cost savings and programmatic procurement benefits of a Mars Forward Lunar Surface System architecture that provides commonality or evolutionary development paths for lunar surface system elements applicable to Mars surface systems. The objective of this paper is to identify the potential benefits for incorporating a Mars Forward development strategy into the planned Project Constellation Lunar Surface System Architecture. The benefits include cost savings, technology readiness, and design validation of systems that would be applicable to lunar and Mars surface systems. The paper presents a survey of previous lunar and Mars surface systems design concepts and provides an assessment of previous conclusions concerning those systems in light of the current Project Constellation Exploration Architectures. The operational requirements for current Project Constellation lunar and Mars surface system elements are compared and evaluated to identify the potential risk reduction strategies that build on lunar surface systems to reduce the technical and programmatic risks for Mars exploration. Risk reduction for rapidly evolving technologies is achieved through systematic evolution of technologies and components based on Moore's Law superimposed on the typical NASA systems engineering project development "V-cycle" described in NASA NPR 7120.5. Risk reduction for established or slowly evolving technologies is achieved through a process called the Mars-Ready Platform strategy in which incremental improvements lead from the initial lunar surface system components to Mars-Ready technologies. The potential programmatic benefits of the Mars Forward strategy are provided in terms of the transition from the lunar exploration campaign to the Mars exploration campaign. By utilizing a sequential combined procurement strategy for lunar and Mars exploration surface systems, the overall budget wedges for exploration systems are reduced and the costly technological development gap between the lunar and Mars programs can be eliminated. This provides a sustained level of technological competitiveness as well as maintaining a stable engineering and manufacturing capability throughout the entire duration of Project Constellation

Tailoring Systems Engineering Processes in a Conceptual Design Environment: A Case Study at NASA Marshall Spaceflight Center's ACO

This paper provides an overview of Systems Engineering as it is applied in a conceptual design space systems department at the National Aeronautics and Space Administration (NASA) Marshall Spaceflight Center (MSFC) Advanced Concepts Office (ACO). Engineering work performed in the NASA MFSC's ACO is targeted toward the Exploratory Research and Concepts Development life cycle stages, as defined in the International Council on Systems Engineering (INCOSE) System Engineering Handbook. This paper addresses three ACO Systems Engineering tools that correspond to three INCOSE Technical Processes: Stakeholder Requirements Definition, Requirements Analysis, and Integration, as well as one Project Process Risk Management. These processes are used to facilitate, streamline, and manage systems engineering processes tailored for the earliest two life cycle stages, which is the environment in which ACO engineers work. The role of systems engineers and systems engineering as performed in ACO is explored in this paper. The need for tailoring Systems Engineering processes, tools, and products in the ever-changing engineering services ACO provides to its customers is addressed

Reduction of Martian Sample Return Mission Launch Mass with Solar Sail Propulsion

Solar sails have the potential to provide mass and cost savings for spacecraft traveling within the inner solar system. Companies like L'Garde have demonstrated sail manufacturability and various in-space deployment methods. The purpose of this study was to evaluate a current Mars sample return architecture and to determine how cost and mass would be reduced by incorporating a solar sail propulsion system. The team validated the design proposed by L'Garde, and scaled the design based on a trajectory analysis. Using the solar sail design reduced the required mass, eliminating one of the three launches required in the original architecture

Z-Pinch Pulsed Plasma Propulsion Technology Development

Fusion-based propulsion can enable fast interplanetary transportation. Magneto-inertial fusion (MIF) is an approach which has been shown to potentially lead to a low cost, small reactor for fusion break even. The Z-Pinch/dense plasma focus method is an MIF concept in which a column of gas is compressed to thermonuclear conditions by an axial current (I approximates 100 MA). Recent advancements in experiments and the theoretical understanding of this concept suggest favorable scaling of fusion power output yield as I(sup 4). This document presents a conceptual design of a Z-Pinch fusion propulsion system and a vehicle for human exploration. The purpose of this study is to apply Z-Pinch fusion principles to the design of a propulsion system for an interplanetary spacecraft. This study took four steps in service of that objective; these steps are identified below. 1. Z-Pinch Modeling and Analysis: There is a wealth of literature characterizing Z-Pinch physics and existing Z-Pinch physics models. In order to be useful in engineering analysis, simplified Z-Pinch fusion thermodynamic models are required to give propulsion engineers the quantity of plasma, plasma temperature, rate of expansion, etc. The study team developed these models in this study. 2. Propulsion Modeling and Analysis: While the Z-Pinch models characterize the fusion process itself, propulsion models calculate the parameters that characterize the propulsion system (thrust, specific impulse, etc.) The study team developed a Z-Pinch propulsion model and used it to determine the best values for pulse rate, amount of propellant per pulse, and mixture ratio of the D-T and liner materials as well as the resulting thrust and specific impulse of the system. 3. Mission Analysis: Several potential missions were studied. Trajectory analysis using data from the propulsion model was used to determine the duration of the propulsion burns, the amount of propellant expended to complete each mission considered. 4. Vehicle Design: To understand the applicability of Z-Pinch propulsion to interplanetary travel, it is necessary to design a concept vehicle that uses it -- the propulsion system significantly impacts the design of the electrical, thermal control, avionics and structural subsystems of a vehicle. The study team developed a conceptual design of an interplanetary vehicle that transports crew and cargo to Mars and back and can be reused for other missions. Several aspects of this vehicle are based on a previous crewed fusion vehicle study -- the Human Outer Planet Exploration (HOPE) Magnetized Target Fusion (MTF) vehicle. Portions of the vehicle design were used outright and others were modified from the MTF design in order to maintain comparability

Deep Space Habitat Configurations Based On International Space Station Systems

A Deep Space Habitat (DSH) is the crew habitation module designed for long duration missions. Although humans have lived in space for many years, there has never been a habitat beyond low-Earth-orbit. As part of the Advanced Exploration Systems (AES) Habitation Project, a study was conducted to develop weightless habitat configurations using systems based on International Space Station (ISS) designs. Two mission sizes are described for a 4-crew 60-day mission, and a 4-crew 500-day mission using standard Node, Lab, and Multi-Purpose Logistics Module (MPLM) sized elements, and ISS derived habitation systems. These durations were selected to explore the lower and upper bound for the exploration missions under consideration including a range of excursions within the Earth-Moon vicinity, near earth asteroids, and Mars orbit. Current methods for sizing the mass and volume for habitats are based on mathematical models that assume the construction of a new single volume habitat. In contrast to that approach, this study explored the use of ISS designs based on existing hardware where available and construction of new hardware based on ISS designs where appropriate. Findings included a very robust design that could be reused if the DSH were assembled and based at the ISS and a transportation system were provided for its return after each mission. Mass estimates were found to be higher than mathematical models due primarily to the use of multiple ISS modules instead of one new large module, but the maturity of the designs using flight qualified systems have potential for improved cost, schedule, and risk benefits

Z-Pinch Magneto-Inertial Fusion Propulsion Engine Design Concept

Fusion-based nuclear propulsion has the potential to enable fast interplanetary transportation. Due to the great distances between the planets of our solar system and the harmful radiation environment of interplanetary space, high specific impulse (Isp) propulsion in vehicles with high payload mass fractions must be developed to provide practical and safe vehicles for human spaceflight missions. Magneto-Inertial Fusion (MIF) is an approach which has been shown to potentially lead to a low cost, small fusion reactor/engine assembly (1). The Z-Pinch dense plasma focus method is an MIF concept in which a column of gas is compressed to thermonuclear conditions by an estimated axial current of approximately 100 MA. Recent advancements in experiments and the theoretical understanding of this concept suggest favorable scaling of fusion power output yield as I(sup 4) (2). The magnetic field resulting from the large current compresses the plasma to fusion conditions, and this is repeated over short timescales (10(exp -6) sec). This plasma formation is widely used in the field of Nuclear Weapons Effects (NWE) testing in the defense industry, as well as in fusion energy research. There is a wealth of literature characterizing Z-Pinch physics and existing models (3-5). In order to be useful in engineering analysis, a simplified Z-Pinch fusion thermodynamic model was developed to determine the quantity of plasma, plasma temperature, rate of expansion, energy production, etc. to calculate the parameters that characterize a propulsion system. The amount of nuclear fuel per pulse, mixture ratio of the D-T and nozzle liner propellant, and assumptions about the efficiency of the engine, enabled the sizing of the propulsion system and resulted in an estimate of the thrust and Isp of a Z-Pinch fusion propulsion system for the concept vehicle. MIF requires a magnetic nozzle to contain and direct the nuclear pulses, as well as a robust structure and radiation shielding. The structure, configuration, and materials of the nozzle must meet many severe requirements. The configuration would focus, in a conical manner, the Deuterium-Tritium (D-T) fuel and Lithium-6/7 liner fluid to meet at a specific point that acts as a cathode so the Li-6 can serve as a current return path to complete the circuit. In addition to serving as a current return path, the Li liner also serves as a radiation shield. The advantage to this configuration is the reaction between neutrons and Li-6 results in the production of additional Tritium, thus adding further fuel to the fusion reaction and boosting the energy output. To understand the applicability of Z-Pinch propulsion to interplanetary travel, it is necessary to design a concept vehicle that uses it. The propulsion system significantly impacts the design of the electrical, thermal control, avionics, radiation shielding, and structural subsystems of a vehicle. The design reference mission is the transport of crew and cargo to Mars and back, with the intention that the vehicle be reused for other missions. Several aspects of this vehicle are based on a previous crewed fusion vehicle study called Human Outer Planet Exploration (HOPE), which employed a Magnetized Target Fusion (MTF) propulsion concept. Analysis of this propulsion system concludes that a 40-fold increase of Isp over chemical propulsion is predicted. This along with a greater than 30% predicted payload mass fraction certainly warrants further development of enabling technologies. The vehicle is designed for multiple interplanetary missions and conceivably may be suited for an automated one-way interstellar voyage

AXTAR: Mission Design Concept

The Advanced X-ray Timing Array (AXTAR) is a mission concept for X-ray timing of compact objects that combines very large collecting area, broadband spectral coverage, high time resolution, highly flexible scheduling, and an ability to respond promptly to time-critical targets of opportunity. It is optimized for submillisecond timing of bright Galactic X-ray sources in order to study phenomena at the natural time scales of neutron star surfaces and black hole event horizons, thus probing the physics of ultradense matter, strongly curved spacetimes, and intense magnetic fields. AXTAR's main instrument, the Large Area Timing Array (LATA) is a collimated instrument with 2-50 keV coverage and over 3 square meters effective area. The LATA is made up of an array of supermodules that house 2-mm thick silicon pixel detectors. AXTAR will provide a significant improvement in effective area (a factor of 7 at 4 keV and a factor of 36 at 30 keV) over the RXTE PCA. AXTAR will also carry a sensitive Sky Monitor (SM) that acts as a trigger for pointed observations of X-ray transients in addition to providing high duty cycle monitoring of the X-ray sky. We review the science goals and technical concept for AXTAR and present results from a preliminary mission design study.Comment: 19 pages, 10 figures, to be published in Space Telescopes and Instrumentation 2010: Ultraviolet to Gamma Ray, Proceedings of SPIE Volume 773