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At NASA, there exists no standardized design or testing protocol for spacecraft fire suppression systems (either handheld or total flooding designs). An extinguisher's efficacy in safely suppressing any reasonable or conceivable fire is the primary benchmark. That concept, however, leads to the question of what a reasonable or conceivable fire is. While there exists the temptation to over-size' the fire extinguisher, weight and volume considerations on spacecraft will always (justifiably) push for the minimum size extinguisher required. This paper attempts to address the question of extinguisher size by examining how large a fire a crew member could successfully survive and extinguish in the confines of a spacecraft. The hazards to the crew and equipment during an accidental fire include excessive pressure rise resulting in a catastrophic rupture of the vehicle skin, excessive temperatures that burn or incapacitate the crew (due to hyperthermia), carbon dioxide build-up or other accumulation of other combustion products (e.g. carbon monoxide). Estimates of these quantities are determined as a function of fire size and mass of material burned. This then becomes the basis for determining the maximum size of a target fire for future fire extinguisher testing.

The combustion of flammable material in a sealed chamber invariably leads to an initial pressure rise in the volume. The pressure rise is due to the increase in the total number of gaseous moles (condensed fuel plus chamber oxygen combining to form gaseous carbon dioxide and water vapor) and, most importantly, the temperature rise of the gas in the chamber. Though the rise in temperature and pressure would reduce with time after flame extinguishment due to the absorption of heat by the walls and contents of the sealed spacecraft, the initial pressure rise from a fire, if large enough, could lead to a vehicle over-pressure and the release of gas through the pressure relief valve. This paper presents a simple lumped-parameter model of the pressure rise in a sealed chamber resulting from the heat release during combustion. The transient model considers the increase in gaseous moles due to combustion, and heat transfer to the chamber walls by convection and radiation and to the fuel-sample holder by conduction, as a function of the burning rate of the material. The results of the model are compared to the pressure rise in an experimental chamber during flame spread tests as well as to the pressure falloff after flame extinguishment. The experiments involve flame spread over thin solid fuel samples. Estimates of the heat release rate profiles for input to the model come from the assumed stoichiometric burning of the fuel along with the observed flame spread behavior. The sensitivity of the model to predict maximum chamber pressure is determined with respect to the uncertainties in input parameters. Model predictions are also presented for the pressure profile anticipated in the Fire Safety-1 experiment, a material flammability and fire safety experiment proposed for the European Space Agency (ESA) Automated Transfer Vehicle (ATV). Computations are done for a range of scenarios including various initial pressures and sample sizes. Based on these results, various mitigation approaches are suggested to prevent vehicle over-pressurization and help guide the definition of the space experiment.

Prior work has revealed the unique aspects of re suppression in reduced gravity and presented the results of trade studies for both the Crew Exploration Vehicle (CEV ) and lunar landers. The trade studies for both of these spacecraft showed that ne water mist (FWM) systems are the best solution; extremely e ective, non-toxic, easy clean-up and minimal interference with spacecraft life support systems. More recent work has shown the initial characterization of a prototype potential FWM system as well as its performance in extinguishing prototype res. There exists, however, no standardized design or testing protocol for spacecraft re suppression systems (either handheld or total ooding designs), let alone a speci cation for a FWM system. While an extinguisher’s e cacy in safely suppressing any reasonable or conceivable re is the primary benchmark, it is not the only performance metric. There exist a range of criteria related to safety and reliability that must be met by any portable extinguisher. In this paper, we present the important performance speci cations of FWM systems with data from prototype and commercial systems as exemplars.

Prior work has revealed the unique aspects of fire suppression in reduced gravity and presented the results of trade studies for both the Crew Exploration Vehicle (CEV) and Lunar landers. The trade studies for both spacecraft showed that water mist systems are the best solution; extremely effective, non-toxic, easy clean-up and minimal interference with spacecraft life support systems. The only downside to water mist is the relatively low Technology Readiness Level (TRL). In the event that the TRL for water mist does not accelerate to a suitable level, the trade analyses showed that water-based foam agents represent an acceptable alternative. There exists no standardized design or testing protocol for spacecraft fire suppression systems (either handheld or total flooding designs). This paper discusses the design of handheld extinguishers, the design of standardized acceptance tests for candidate extinguishers and the results of tests for both water-based foam and water-mist extinguishers.

Using numerical simulations, this work demonstrates a concept called back-end ignition for lighting-off and pre-heating a catalytic monolith in a power generation system. In this concept, a downstream heat source (e.g. a flame) or resistive heating in the downstream portion of the monolith initiates a localized catalytic reaction which subsequently propagates upstream and heats the entire monolith. The simulations used a transient numerical model of a single catalytic channel which characterizes the behavior of the entire monolith. The model treats both the gas and solid phases and includes detailed homogeneous and heterogeneous reactions. An important parameter in the model for back-end ignition is upstream heat conduction along the solid. The simulations used both dry and wet CO chemistry as a model fuel for the proof-of-concept calculations; the presence of water vapor can trigger homogenous reactions, provided that gas-phase temperatures are adequately high and there is sufficient fuel remaining after surface reactions. With sufficiently high inlet equivalence ratio, back-end ignition occurs using the thermophysical properties of both a ceramic and metal monolith (coated with platinum in both cases), with the heat-up times significantly faster for the metal monolith. For lower equivalence ratios, back-end ignition occurs without upstream propagation. Once light-off and propagation occur, the inlet equivalence ratio could be reduced significantly while still maintaining an ignited monolith as demonstrated by calculations using complete monolith heating.

Less-intrusive, fast-responding, and full-field temperature measurements have long been a desired tool for the research community. Recently, the emission of a silicon-carbide (SiC) fiber placed in a flowing hot (or reacting) gas has been used to measure the temperature profile along the length of the fiber. The relationship between the gas and fiber temperature comes from an energy balance on the fiber. In the present work, we compared single point flame temperature measurements using thin-filament pyrometry (TFP) and thermocouples. The data was from vertically traversing a thermocouple and a SiC fiber through a methanol/air diffusion flame of a porous-metal wick burner. The results showed that the gas temperature using the TFP technique agreed with the thermocouple measurements (25.4 m diameter wire) within 3.5% for temperatures above 1200 K. Additionally, we imaged the entire SiC fiber (with a spatial resolution of 0.14 mm) while it was in the flame using a high resolution CCD camera. The intensity level along the fiber length is a function of the temperature. This results in a one-dimensional temperature profiles at various heights above the burner wick. This temperature measurement technique, while having a precision of less than 1 K, showed data scatter as high as 38 K. Finally, we discuss the major sources of uncertainty in gas temperature measurement using TFP.

Combustion tests were carried out on the Mir Space Station. Flat sheets of paper, polyethyleneinsulated wires, cylindrical cellulosic samples, and candles were burned in microgravity. The test parameters included sample size, fuel preheating levels, and lowspeed air velocity. Data were collected mainly through video cameras, audio recordings of crew observations, and 35 mm still pictures. For many of these tests, thermocouples permitted the recording of temperatures of the gas phase flame and/or solid fuel. After the flight, the flame images and temperature data were compared to numerical simulations. Several unique phenomena were observed and the results have implications for spacecraft fire safety. These include the influence of airflow, fuel melting and bubbling, and fuel-vapor generation, and condensation after the flame extinguished.