Donovan Mathias

Numerical Simulation of Bolide Entry with Ground Footprint Prediction

As they decelerate through the atmosphere, meteors deposit mass, momentum and energy into the surrounding air at tremendous rates. Trauma from the entry of such bolides produces strong blast waves that can propagate hundreds of kilometers and cause substantial terrestrial damage even when no ground impact occurs. We present a new simulation technique for airburst blast prediction using a fully-conservative, Cartesian mesh, finite-volume solver and investigate the ability of this method to model far- field propagation over hundreds of kilometers. The work develops mathematical models for the deposition of mass, momentum and energy into the atmosphere and presents verification and validation through canonical problems and the comparison of surface overpressures, and blast arrival times with actual results in the literature for known bolides. The discussion also examines the effects of various approximations to the physics of bolide entry that can substantially decrease the computational expense of these simulations. We present parametric studies to quantify the influence of entry-angle, burst-height and other parameters on the ground footprint of the airburst, and these values are related to predictions from analytic and handbook-methods.

Simulation assisted risk assessment applied to launch vehicle conceptual design

A simulation-based risk assessment approach is presented and is applied to the analysis of abort during the ascent phase of a space exploration mission. The approach utilizes groupings of launch vehicle failures, referred to as failure bins, which are mapped to corresponding failure environments. Physical models are used to characterize the failure environments in terms of the risk due to blast overpressure, resulting debris field, and the thermal radiation due to a fireball. The resulting risk to the crew is dynamically modeled by combining the likelihood of each failure, the severity of the failure environments as a function of initiator and time of the failure, the robustness of the crew module, and the warning time available due to early detection. The approach is shown to support the launch vehicle design process by characterizing the risk drivers and identifying regions where failure detection would significantly reduce the risk to the crew.

Simulating plasma production from hypervelocity impacts

Hypervelocity particles, such as meteoroids and space debris, routinely impact spacecraft and are energetic enough to vaporize and ionize themselves and as well as a portion of the target material. The resulting plasma rapidly expands into the surrounding vacuum. While plasma measurements from hypervelocity impacts have been made using ground-based technologies such as light gas guns and Van de Graaff dust accelerators, some of the basic plasma properties vary significantly between experiments. There have been both ground-based and in-situ measurements of radio frequency (RF) emission from hypervelocity impacts, but the physical mechanism responsible and the possible connection to the impact-produced plasma are not well understood. Under certain conditions, the impact-produced plasma can have deleterious effects on spacecraft electronics by providing a new current path, triggering an electrostatic discharge, causing electromagnetic interference, or generating an electromagnetic pulse. Multi-physics simula...

Simulation-Assisted Risk Assessment

A probabilistic risk assessment (PRA) approach has been developed and applied to the risk analysis of capsule abort during ascent. The PRA is used to assist in the identification of modeling and simulation applications that can significantly impact the understanding of crew risk during this potentially dangerous maneuver. The PRA approach is also being used to identify the appropriate level of fidelity for the modeling of those critical failure modes. The Apollo launch escape system (LES) was chosen as a test problem for application of this approach. Failure modes that have been modeled and/or simulated to date include explosive overpressure-based failure, explosive fragment-based failure, land landing failures (range limits exceeded either near launch or Mode III trajectories ending on the African continent), capsule-booster re-contact during separation, and failure due to plume-induced instability. These failure modes have been investigated using analysis tools in a variety of technical disciplines at various levels of fidelity. The current paper focuses on the roles and impacts of the higher-fidelity methods on this process and, by association, the roles and impacts of the high performance computing resources of the Columbia supercomputer system at NASA Ames Research Center.

A risk evaluation approach for safety in aerospace preliminary design

The preliminary design phase of any program is key to its eventual successful development. The more advanced a design the more this tends to be true. For this reason the preliminary design phase is particularly important in the design of aerospace systems. Errors in preliminary design tend to be fundamental and tend to cause programs to be abandoned, or to be changed fundamentally, and at great cost later in the design development. In the past aerospace system designers have used the tools of systems engineering to enable the development of designs that were more likely to be functionally adequate. However to do so has meant the application significant resources to the review and investigation of proposed design alternatives. This labor-intensive process can no longer be afforded in the current design environment. The realization has led to the development of an approach that attempts to focus the tools of systems engineering on the risk drivers in the design. One of the most important factors in the development of successful designs is adequately addressing the safety and reliability risk. All too often these important features of the developed design are left to afterthoughts as the design gives sway to the more traditional performance focus. Thus even when a successful functional design is forthcoming significant resources are often required to reduce its reliability and safety risk to an acceptable level. This builds upon the experience base of the integrated shuttle risk assessment and its expansions and applications to the evaluation of newly proposed launcher designs. The approach used the shuttle developed PRA models and associated data sets as functional analogs for new launcher functions. The concept is that associated models would characterize the function of any launcher developed for those functions on the shuttle. Once this functional decomposition and reconstruction has been accomplished a proposed new design is compared on a function-by-function basis and specific design enhancements that have significant promise of reducing the functional risk over the shuttle are highlighted. The potential for enhancement is then incorporated into those functions by suitable modification of the shuttle models and or the associated quantification data sets representing those design features addressed by the new design. The level of risk reduction potential is then estimated by those component failure modes and mechanisms identified for the shuttle function and eliminated in the new design. In addition heritage data that would support the claims of risk reduction for those failure modes and mechanisms that remain albeit at a reduced level of risk are applied.

Blast overpressure modeling enhancements for application to risk-informed design of human space flight launch abort systems

This paper describes recent enhancements to the engineering-level analysis tools used by the simulation assisted risk assessment (SARA) project (Ref. 1) at NASA Ames Research Center in evaluating the blast overpressure risk to the crew. The primary enhancements to the model include incorporation of vapor cloud explosion (VCE) curve fits for propellant explosions, development of an improved model for the effects of vehicle velocity on blast propagation, improvement in the representation of blast/vehicle interaction effects, and incorporation of pressure vs. impulse (P-I) failure criteria to better represent structural failure modes. High-fidelity computational fluid dynamics (CFD) simulations, using the Overflow2 (Ref. 2) code, played a crucial role in the development of some of these enhancements. A subset of the high-fidelity results is presented.

Explosion Hazard from a Propellant-Tank Breach in Liquid Hydrogen-Oxygen Rockets

An engineering risk assessment of the conditions for massive explosions of cryogenic liquid hydrogen-oxygen rockets during launch accidents is presented. The assessment is based on the analysis of the data of purposeful rupture experiments with liquid oxygen and hydrogen tanks and on an interpretation of these data via analytical semiquantitative estimates and numerical simulations of simplified models for the whole range of the physical phenomena governing the outcome of a propellant-tank breach. The following sequence of events is reconstructed: rupture of fuel tanks, escape of the fluids from the ruptured tanks, liquid film boiling, fragmentation of liquid flow, formation of aerosol oxygen and hydrogen clouds, mixing of the clouds, droplet evaporation, self-ignition of the aerosol clouds, and aerosol combustion. The power of the explosion is determined by a small fraction of the escaped cryogens that become well mixed within the aerosol cloud during the delay time between rupture and ignition. Several ...

Comparative analysis of static and dynamic probabilistic risk assessment

This study examines three different methodologies for producing loss-of-mission (LOM) and loss-of-crew (LOC) risks estimates for probabilistic risk assessments (PRA) of crewed spacecraft. The three bottom-up, component-based PRA approaches examined are a traditional static fault tree, a dynamic Monte Carlo simulation, and a fault tree hybrid that incorporates some dynamic elements. These approaches were used to model the reaction control system thruster pod of a generic crewed spacecraft and mission, and a comparative analysis of the methods is presented. The methodologies are assessed in terms of the process of modeling a system, the actionable information produced for the design team, and the overall fidelity of the quantitative risk evaluation generated. The system modeling process is compared in terms of the effort required to generate the initial model, update the model in response to design changes, and support mass-versus-risk trade studies. The results are compared by examining the top-level LOM/LOC estimates and the relative risk driver rankings at the failure mode level. The fidelity of each modeling methodology is discussed in terms of its capability to handle real-world system dynamics such as cold-sparing, changes in mission operations due to loss of redundancy, and common cause failure modes. The paper also discusses the applicability of each methodology to different phases of system development and shows that a single methodology may not be suitable for all of the many purposes of a spacecraft PRA. The fault tree hybrid approach is shown to be best suited to the needs of early assessments during conceptual design phases. As the design begins to mature, the level of detail represented in the risk model must go beyond redundancy and nominal mission operations to include dynamic, time- and state-dependent system responses as well as diverse system capabilities. This is best accomplished using the dynamic simulation approach, since these phenomena are not easily captured by static methods. Ultimately, once the design has been finalized and the goal of the PRA is to provide design validation and requirement verification, more traditional, static fault tree approaches may become as appropriate as the simulation method.

Integrated risk sensitivity study for Lunar Surface Systems

This paper illustrates an innovative approach to assessing the reliability of conceptual Lunar Surface Systems architectures using an integrated analysis model. The integrated model represents systems, dependencies, and interactions to develop risk-based reliability requirements that balance functional characteristics, needs, demands, and constraints to achieve availability goals. The model utilizes “availability” metrics based on first-order descriptions of the architecture to begin providing reliability impacts even before much design detail exists. Sensitivity analyses are performed to identify key risk parameters and find “knees” in the curve for establishment of system architecture- and element-level requirements.

Susceptibility of spacecraft to impact-induced electromagnetic pulses

Spacecraft are routinely bombarded with interplanetary dust particles, called meteoroids, and defunct objects of human origin, called orbital debris. Collectively we refer to these particles as hypervelocity impactors. Meteoroids have impact speeds up to 72 km/s and orbital debris have impact speeds <;11 km/s in low Earth orbit. Most hypervelocity impactors possess enough energy to ionize and vaporize themselves as well as a significant portion of the spacecraft material upon impact, forming a plasma that rapidly expands into the surrounding vacuum. A plasma is a gas of charged particles whose dynamics are dominated by electromagnetic forces. Under certain conditions, the expansion of the impact-induced plasma can trigger electrostatic discharges and electromagnetic pulses that can disable or destroy spacecraft electronics, and in the worst cases, result in complete loss of mission. A number of spacecraft have experienced unexplained electrical anomalies correlated with impact events or meteoroid showers. The associated electrical effects and potential for damage to satellite electronics through these processes have not been previously investigated. This paper describes multi-physics simulations of particle impacts on spacecraft using a combination of computational continuum dynamics and electromagnetic particle-in-cell methods. These simulations incorporate elasticity and plasticity of the solid target, phase change and plasma formation, strongly coupled plasma physics due to the high density and low temperature of the plasma, a fully kinetic description of the plasma, and free space electromagnetic radiation. By simulating a series of hypervelocity impacts, we determine properties (e.g., temperature, expansion speed, and charge state) of the plasma plume for impact speeds from 10 km /s to 72 km/s, and particle masses from one femtogram to one microgram. These plasma properties yield the amplitude, frequency, directionality, and the spatial and temporal decay of impact-induced electromagnetic pulses. In this paper, these simulation results are used to assess the susceptibility of spacecraft components to electrical damage from meteoroid and orbital debris impacts. The results show that electromagnetic pulses are only produced for impact speeds greater than 18 km/s and that the pulses are capable of causing significant damage via current and voltage spikes for impact speeds greater than 50 km/s. The particle mass does not affect these speed thresholds. The model predicts that the electric and magnetic field limits to which spacecraft electronics are currently designed are far below the fields produced by the fastest meteoroid impacts. While electronics are normally shielded in a Faraday cage, this also provides insufficient mitigation at the expected frequencies of the radiation from electromagnetic pulses. Results indicate the fields decay rapidly from the point of impact, meaning that a susceptible component must be physically nearby to be threatened. Understanding key parameters of impact plasma plumes and associated electromagnetic pulses will aid in designing more robust and reliable spacecraft that are well protected in the space environment.