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Search indexed NASA NTRS and DOE OSTI research on propulsion, heat transfer, battery materials and energy systems. Follow report and document links to the original sources.

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At least 343 records · Page 19
All sourcesNASA NTRSDOE OSTI
NASA NTRS1990

Rapid near-optimal trajectory generation and guidance law development for single-stage-to-orbit airbreathing vehicles

General problems associated with on-board trajectory optimization, propulsion system cycle selection, and with the synthesis of guidance laws were addressed for an ascent to low-earth-orbit of an air-breathing single-stage-to-orbit vehicle. The NASA Generic Hypersonic Aerodynamic Model Example and the Langley Accelerator aerodynamic sets were acquired and implemented. Work related to the development of purely analytic aerodynamic models was also performed at a low level. A generic model of a multi-mode propulsion system was developed that includes turbojet, ramjet, scramjet, and rocket engine cycles. Provisions were made in the dynamic model for a component of thrust normal to the flight path. Computational results, which characterize the nonlinear sensitivity of scramjet performance to changes in vehicle angle of attack, were obtained and incorporated into the engine model. Additional trajectory constraints were introduced: maximum dynamic pressure; maximum aerodynamic heating rate per unit area; angle of attack and lift limits; and limits on acceleration both along and normal to the flight path. The remainder of the effort focused on required modifications to a previously derived algorithm when the model complexity cited above was added. In particular, analytic switching conditions were derived which, under appropriate assumptions, govern optimal transition from one propulsion mode to another for two cases: the case in which engine cycle operations can overlap, and the case in which engine cycle operations are mutually exclusive. The resulting guidance algorithm was implemented in software and exercised extensively. It was found that the approximations associated with the assumed time scale separation employed in this work are reasonable except over the Mach range from roughly 5 to 8. This phenomenon is due to the very large thrust capability of scramjets in this Mach regime when sized to meet the requirement for ascent to orbit. By accounting for flight path angle and flight path angle rate in construction of the flight path over this Mach range, the resulting algorithm provides the means for rapid near-optimal trajectory generation and propulsion cycle selection over the entire Mach range from take-off to orbit.

Calise, A. J.↗
NASA NTRS1998

Detailed Experimental Data for CFD Code Validation

Recent interest in low cost, reliable access to space has generated increased interest in advanced technology approaches to space transportation systems. A key to the success of such programs lies in the development of advanced propulsion systems capable of achieving the performance and operations goals required for the next generation of space vehicles. One extremely promising approach involves the combination of rocket and air-breathing engines into a rocket-based combined-cycle engine (RBCC). A key element of that engine is the rocket ejector that is utilized in the zero to Mach two operating regime. Studies of RBCC engine concepts are not new and studies dating back thirty years are well documented in the literature. However, studies focused on the rocket ejector mode of the RBCC cycle are lacking. The present investigation utilizes an integrated experimental and computation fluid dynamics (CFD) approach to examine critical rocket ejector performance issues. In particular, the development of a predictive methodology capable of performance prediction is a key objective in order to analyze thermal choking and its control, primary/secondary pressure matching considerations, and effects of nozzle expansion ratio. To achieve this objective, the present study emphasizes obtaining new data using advanced optical diagnostics such as Schlieren photography and Raman spectroscopy, and CFD techniques to investigate mixing in the rocket ejector mode. A new research facility for the study of the rocket ejector mode has been developed. In the last milestone report, the operational capabilities of this research facility were described. In this milestone report, the experimentally obtained measurements for CFD code validation are presented and discussed.

Santoro, Robert J.↗
NASA NTRS1997

Rocket Ejector Studies for Application to RBCC Engines: An Integrated Experimental/CFD Approach

Recent interest in low cost, reliable access to space has generated increased interest in advanced technology approaches to space transportation systems. A key to the success of such programs lies in the development of advanced propulsion systems capable of achieving the performance and operations goals required for the next generation of space vehicles. One extremely promising approach involves the combination of rocket and air- breathing engines into a rocket-based combined-cycle engine (RBCC). A key element of that engine is the rocket ejector which is utilized in the zero to Mach two operating regime. Studies of RBCC engine concepts are not new and studies dating back thirty years are well documented in the literature. However, studies focused on the rocket ejector mode of the RBCC cycle are lacking. The present investigation utilizes an integrated experimental and computation fluid dynamics (CFD) approach to examine critical rocket ejector performance issues. In particular, the development of a predictive methodology capable of performance prediction is a key objective in order to analyze thermal choking and its control, primary/secondary pressure matching considerations, and effects of nozzle expansion ratio. To achieve this objective, the present study emphasizes obtaining new data using advanced optical diagnostics such as Raman spectroscopy and CFD techniques to investigate mixing in the rocket ejector mode. A new research facility for the study of the rocket ejector mode is described along with the diagnostic approaches to be used. The CFD modeling approach is also described along with preliminary CFD predictions obtained to date.

Pal, S.↗
NASA NTRS2002

Carbon Nanotubes Growth by CVD on Graphite Fibers

Due to the superior electrical and mechanical properties of carbon nanotubes (CNT), synthesizing CNT on various substances for electronics devices and reinforced composites have been engaged in many efforts for applications. This presentation will illustrate CNT synthesized on graphite fibers by thermal CVD. On the fiber surface, iron nanoparticles as catalysts for CNT growth are coated. The growth temperature ranges from 600 to 1000 C and the pressure ranges from 100 Torr to one atmosphere. Methane and hydrogen gases with methane content of 10% to 100% are used for the CNT synthesis. At high growth temperatures (greater than or equal to 900 C), the rapid inter-diffusion of the transition metal iron on the graphite surface results in the rough fiber surface without any CNT grown on it. When the growth temperature is relative low (650-800 C), CNT with catalytic particles on the nanotube top ends are fabricated on the graphite surface. (Methane and hydrogen gases with methane content of 10% to 100% are used for the CNT synthesis.) (By measuring the samples) Using micro Raman spectroscopy in the breath mode region, single-walled or multi-walled CNT (MWCNT), depending on growth concentrations, are found. Morphology, length and diameter of these MWCNT are determined by scanning electron microscopy and Raman spectroscopy. The detailed results of syntheses and characterizations will be discussed in the presentation.

Zhu, Shen↗
NASA NTRS1993

Navy GTE seal development activity

Under the auspices of the Integrated High Performance Turbine Engine Technology Initiative, the Naval Air Warfare Center conducts advanced development programs for demonstration in the next generation of air-breathing propulsion systems. Among the target technologies are gas path and lube oil seals. Two development efforts currently being managed by NAWCAD are the High Performance Compressor Discharge Film-Riding Face Seal and the Subsonic Core High Speed Air/Oil Seal. The High Performance Compressor Discharge Film-Riding Face Seal Program aims at reducing parasitic leakage through application of a film-riding face sea concept to the compressor discharge location of a Phase 2 IHPTET engine. An order-of-magnitude leakage reduction relative to current labyrinth seal configurations is expected. Performance goals for these seals are (1) 1200 F air temperature, (2) 800 feet-per-second surface velocity, and (3) 600 SPI differential pressure. The two designs chosen for fabrication and rig test are a spiral groove and a Rayleigh step seal. Rig testing is currently underway. The Subsonic Core High Speed Air/Oil Seal Program is developing shaft-to-ground seals for next-generation propulsion systems that will minimize leakage and provide full life. Significantly higher rotor speeds and temperatures will be experienced. Technologies being exploited include, hydrodynamic lift assist features, ultra light weight designs, and improved cooling schemes. Parametric testing has been completed; a final seal design is entering the endurance test phase.

Grala, Carl P.↗
NASA NTRS1991

A 99 percent purity molecular sieve oxygen generator

Molecular sieve oxygen generating systems (MSOGS) have become the accepted method for the production of breathable oxygen on military aircraft. These systems separate oxygen for aircraft engine bleed air by application of pressure swing adsorption (PSA) technology. Oxygen is concentrated by preferential adsorption in nitrogen in a zeolite molecular sieve. However, the inability of current zeolite molecular sieves to discriminate between oxygen and argon results in an oxygen purity limitations of 93-95 percent (both oxygen and argon concentrate). The goal was to develop a new PSA process capable of exceeding the present oxygen purity limitations. A novel molecular sieve oxygen concentrator was developed which is capable of generating oxygen concentrations of up to 99.7 percent directly from air. The process is comprised of four absorbent beds, two containing a zeolite molecular sieve and two containing a carbon molecular sieve. This new process may find use in aircraft and medical breathing systems, and industrial air separation systems. The commercial potential of the process is currently being evaluated.

Miller, G. W.↗
NASA NTRS2002

GTX Reference Vehicle Structural Verification Methods and Weight Summary

The design of a single-stage-to-orbit air breathing propulsion system requires the simultaneous development of a reference launch vehicle in order to achieve the optimal mission performance. Accordingly, for the GTX study a 300-lb payload reference vehicle was preliminarily sized to a gross liftoff weight (GLOW) of 238,000 lb. A finite element model of the integrated vehicle/propulsion system was subjected to the trajectory environment and subsequently optimized for structural efficiency. This study involved the development of aerodynamic loads mapped to finite element models of the integrated system in order to assess vehicle margins of safety. Commercially available analysis codes were used in the process along with some internally developed spreadsheets and FORTRAN codes specific to the GTX geometry for mapping of thermal and pressure loads. A mass fraction of 0.20 for the integrated system dry weight has been the driver for a vehicle design consisting of state-of-the-art composite materials in order to meet the rigid weight requirements. This paper summarizes the methodology used for preliminary analyses and presents the current status of the weight optimization for the structural components of the integrated system.

Hunter, J. E.↗
NASA NTRS2002

GTX Reference Vehicle Structural Verification Methods and Weight Summary

The design of a single-stage-to-orbit air breathing propulsion system requires the simultaneous development of a reference launch vehicle in order to achieve the optimal mission performance. Accordingly, for the GTX study a 300-lb payload reference vehicle was preliminary sized to a gross liftoff weight (GLOW) of 238,000 lb. A finite element model of the integrated vehicle/propulsion system was subjected to the trajectory environment and subsequently optimized for structural efficiency. This study involved the development of aerodynamic loads mapped to finite element models of the integrated system in order to assess vehicle margins of safety. Commercially available analysis codes were used in the process along with some internally developed spread-sheets and FORTRAN codes specific to the GTX geometry for mapping of thermal and pressure loads. A mass fraction of 0.20 for the integrated system dry weight has been the driver for a vehicle design consisting of state-of-the-art composite materials in order to meet the rigid weight requirements. This paper summarizes the methodology used for preliminary analyses and presents the current status of the weight optimization for the structural components of the integrated system.

Hunter, J. E.↗
NASA NTRS2002

Influence of vestibular activation on respiration in humans

The purpose of this study was to determine the effects of the semicircular canals and otolith organs on respiration in humans. On the basis of animal studies, we hypothesized that vestibular activation would elicit a vestibulorespiratory reflex. To test this hypothesis, respiratory measures, arterial blood pressure, and heart rate were measured during engagement of semicircular canals and/or otolith organs. Dynamic upright pitch and roll (15 cycles/min), which activate the otolith organs and semicircular canals, increased respiratory rate (Delta2 +/- 1 and Delta3 +/- 1 breaths/min, respectively; P < 0.05). Dynamic yaw and lateral pitch (15 cycles/min), which activate the semicircular canals, increased respiration similarly (Delta3 +/- 1 and Delta2 +/- 1, respectively; P < 0.05). Dynamic chair rotation (15 cycles/min), which mimics dynamic yaw but eliminates neck muscle afferent, increased respiration (Delta3 +/- 1; P < 0.05) comparable to dynamic yaw (15 cycles/min). Increases in respiratory rate were graded as greater responses occurred during upright (Delta5 +/- 2 breaths/min) and lateral pitch (Delta4 +/- 1) and roll (Delta5 +/- 1) performed at 30 cycles/min. Increases in breathing frequency resulted in increases in minute ventilation during most interventions. Static head-down rotation, which activates otolith organs, did not alter respiratory rate (Delta1 +/- 1 breaths/min). Collectively, these data indicate that semicircular canals, but not otolith organs or neck muscle afferents, mediate increased ventilation in humans and support the concept that vestibular activation alters respiration in humans.

NASA Discipline Neuroscience↗
NASA NTRS2004

Upgrades at the Propulsion Systems Lab (PSL)

The Propulsion Systems Lab (PSL) does ground testing on full size air breathing engines. These engines range from those on commercial airplanes to fighter jets. At the PSL, engineers receive test requirements from customers and put together the necessary instrumentation, data systems, power requirements, electrical control valves, and engine controls. The engineers are also responsible for facility maintenance, repairs and upgrades. There are four major sections at the PSL; the Test floor, the Data room, the Control room and, the WDPF room. On the test floor are two test cells, cell #3 and cell #4. It is within these cells that the actual engine resides for ground testing. The cells, once sealed and taken up to altitude, are capable of reaching engine inlet temperatures of 1000 F to -90 F, and various atmospheric pressures. The engine, when operational, takes in air and gives out exhaust of up to 2000 F. The exhaust is led to another section of the cell where it is cooled to 150 F before finally redirected to the appropriate disposer. Temperature and pressure transducers detect the conditions within the cell and transmit them to the data room where the results are captured, processed, analyzed, and translated to a more comprehensive language. This is made possible with the aid of several programmable logic controllers (PLCs) and instrumentation and control systems. The translated data is then sent, via the LAN, to the control room where the results can be viewed on monitors by the engineers and customers. From the control room, the test cell conditions can be changed whenever desired. During tests, a lot takes place in the facility. The WDPF control system monitors and controls all facility parameters. This summer, I will assist the engineers; on an upgrade to the facility's distributed control and dynamic data system, in preparation for an engine test that will begin in September, the installation of control systems and various miscellaneous projects around the PSL.

Audu, Abdullahi M.↗
NASA NTRS1991

Integrated trajectory and control analysis for generic hypersonic vehicles

A tool which supports integrated assessment of air-breathing hypersonic vehicle trajectory management and control strategy has been developed by Hattis and Malchow (1991). Using a two-point boundary value problem solution technique, the tool can construct an entire near-minimum-fuel trajectory and desired control strategy from takeoff to orbit, while satisfying physically derived inequality constraints and while achieving efficient propulsive mode phasing. A strategy is also formulated to construct preliminary trajectory and control history representations with less computational burden than required for the overall flight profile assessment. A tabulated example hypersonic vehicle model has been used to demonstrate the integrated analysis methodology. To assure good numerical behavior when the algorithm uses tabulated model data, a multidimensional cubic spline curve smoothing routine has been developed. A dynamic-pressure-constrained, near-fuel-optimal trajectory was computed to provide examples of information that the methodology can provide. Previously unspecified propulsive discontinuities were located, flight regimes demanding rapid attitude changes were identified, available and closed-loop controller authority was ascertained, and inadequacies in vehicle and specific subsystem model representations were determined.

Hattis, Philip D.↗
NASA NTRS1997

Thrust Enhancement in Hypervelocity Nozzles by Chemical Catalysis

In the hypersonic flight regime, the air-breathing supersonic combustion ramjet (scramjet) has been shown to be a viable propulsion system. The current designs of scramjet engines provide performance benefits only up to a Mach number of 14. Performance losses increase rapidly as the Mach number increases. To extend the applicability of scram'jets beyond Mach 14, research is being conducted in the area of inlet and wave drag reduction, skin-friction and heat-transfer reduction, nozzle loss minimization, low-loss mixing, and combustion enhancement. For high Mach number applications, hydrogen is the obvious fuel choice because of its high energy content per unit mass in comparison with conventional fuels. These flight conditions require engines to operate at supersonic internal velocities, high combustor temperatures, and low static pressures. The high static temperature condition enhances the production of radicals such as H and OH, and the low-pressure condition slows the reaction rates, particularly the recombination reactions. High-temperature and low-pressure constraints, in combination with a small residence time, result in a radical-rich exhaust gas mixture exiting the combustor. At high Mach number conditions (due to low residence time), H and OH do not have enough time to recombine ; thus, a significant amount of energy is lost as these high-energy free radical are exhausted. The objective of the present study is to conduct a flowfield analysis for a typical nozzle geometry for NASP-type vehicle to assess for thrust enhancement in hypervelocity nozzles by substituting small amount of phosphine for hydrogen.

Singh, D. J.↗
NASA NTRS1991

Air-breathing hypersonic vehicle guidance and control studies; An integrated trajectory/control analysis methodology: Phase 1

A tool which generates optimal trajectory/control histories in an integrated manner is generically adapted to the treatment of single-stage-to-orbit air-breathing hypersonic vehicles. The methodology is implemented as a two point boundary value problem solution technique. Its use permits an assessment of an entire near-minimum-fuel trajectory and desired control strategy from takeoff to orbit while satisfying physically derived inequality constraints and while achieving efficient propulsive mode phasing. A simpler analysis strategy that partitions the trajectory into several boundary condition matched segments is also included to construct preliminary trajectory and control history representations with less computational burden than is required for the overall flight profile assessment. A demonstration was accomplished using a tabulated example (winged-cone accelerator) vehicle model that is combined with a newly developed multidimensional cubic spline data smoothing routine. A constrained near-fuel-optimal trajectory, imposing a dynamic pressure limit of 1000 psf, was developed from horizontal takeoff to 20,000 ft/sec relative air speed while aiming for a polar orbit. Previously unspecified propulsive discontinuities were located. Flight regimes demanding rapid attitude changes were identified, dictating control effector and closed-loop controller authority was ascertained after evaluating effector use for vehicle trim. Also, inadequacies in vehicle model representations and specific subsystem models with insufficient fidelity were determined based on unusual control characteristics and/or excessive sensitivity to uncertainty.

Hattis, Philip D.↗
NASA NTRSDate not supplied

Auxiliary Inlet Design Study for Mach 1.4

Auxiliary inlet configurations were developed for inlets designed for a NASA Mach 1.4 Supersonic Technology Concept Aeroplane (STCA). The inlets included axisymmetric pitot and axisymmetric spike inlets. The auxiliary inlets were full circumferential slots within the cowl. Computational fluid dynamics (CFD) simulations were performed to obtain inlet performance metrics of total pressure recovery and radial distortion at the engine face for take-off and approach conditions. Methods of design of experiments were used to explore the statistical significance of the design factors for the auxiliary inlets which included the axial location, length of the opening, and auxiliary inlet angle. The results demonstrate the effect of the design factors and show the potential for improved inlet performance with the use of auxiliary inlets.

supersonic inlets↗
NASA NTRS2005

Heat Transfer and Thermal Stability Research for Advanced Hydrocarbon Fuel Technologies

In recent years there has been increased interest in the development of a new generation of high performance boost rocket engines. These efforts, which will represent a substantial advancement in boost engine technology over that developed for the Space Shuttle Main Engines in the early 1970s, are being pursued both at NASA and the United States Air Force. NASA, under its Space Launch Initiative s Next Generation Launch Technology Program, is investigating the feasibility of developing a highly reliable, long-life, liquid oxygen/kerosene (RP-1) rocket engine for launch vehicles. One of the top technical risks to any engine program employing hydrocarbon fuels is the potential for fuel thermal stability and material compatibility problems to occur under the high-pressure, high-temperature conditions required for regenerative fuel cooling of the engine combustion chamber and nozzle. Decreased heat transfer due to carbon deposits forming on wetted fuel components, corrosion of materials common in engine construction (copper based alloys), and corrosion induced pressure drop increases have all been observed in laboratory tests simulating rocket engine cooling channels. To mitigate these risks, the knowledge of how these fuels behave in high temperature environments must be obtained. Currently, due to the complexity of the physical and chemical process occurring, the only way to accomplish this is empirically. Heated tube testing is a well-established method of experimentally determining the thermal stability and heat transfer characteristics of hydrocarbon fuels. The popularity of this method stems from the low cost incurred in testing when compared to hot fire engine tests, the ability to have greater control over experimental conditions, and the accessibility of the test section, facilitating easy instrumentation. These benefits make heated tube testing the best alternative to hot fire engine testing for thermal stability and heat transfer research. This investigation used the Heated Tube Facility at the NASA Glenn Research Center to perform a thermal stability and heat transfer characterization of RP-1 in an environment simulating that of a high chamber pressure, regenerative cooled rocket engine. The first step in the research was to investigate the carbon deposition process of previous heated tube experiments by performing scanning electron microscopic analysis in conjunction with energy dispersive spectroscopy on the tube sections. This analysis gave insight into the carbon deposition process and the effect that test conditions played in the formation of deleterious coke. Furthermore, several different formations were observed and noted. One other crucial finding of this investigation was that in sulfur containing hydrocarbon fuels, the interaction of the sulfur components with copper based wall materials presented a significant corrosion problem. This problem in many cases was more life limiting than those posed by the carbon deposition process. The results of this microscopic analysis was detailed and presented at the December 2003 JANNAF Air-Breathing Propulsion Meeting as a Materials Compatibility and Thermal Stability Analysis of common Hydrocarbon Fuels (reference 1).

DeWitt, Kenneth↗
NASA NTRS1996

Propulsion Selection for 85kft Remotely Piloted Atmospheric Science Aircraft

This paper describes how a 3 stage turbocharged gasoline engine was selected to power NASA's atmospheric science unmanned aircraft now under development. The airplane, whose purpose is to fly sampling instruments through targeted regions of the upper atmosphere at the exact location and time (season, time of day) where the most interesting chemistry is taking place, must have a round trip range exceeding 1000 km, carry a payload of about 500 lb to altitudes exceeding 80 kft over the site, and be able to remain above that altitude for at least 30 minutes before returning to base. This is a subsonic aircraft (the aerodynamic heating and shock associated with supersonic flight could easily destroy the chemical species that are being sampled) and it must be constructed so it will operate out of small airfields at primitive remote sites worldwide, under varying climate and weather conditions. Finally it must be low cost, since less than $50 M is available for its development. These requirements put severe constraints on the aircraft design (for example, wing loading in the vicinity of 10 psf) and have in turn limited the propulsion choices to already-existing hardware, or limited adaptations of existing hardware. The only candidate that could emerge under these circumstances was a propeller driven aircraft powered by spark ignited (SI) gasoline engines, whose intake pressurization is accomplished by multiple stages of turbo-charging and intercooling. Fortunately the turbocharged SI powerplant, owing to its rich automotive heritage and earlier intensive aero powerplant development during WWII, enjoys in addition to its potentially low development costs some subtle physical advantages (arising from its near-stochiometric combustion) that may make it smaller and lighter than either a turbine engine or a diesel for these altitudes. Just as fortunately, the NASA/industry team developing this aircraft includes the same people who built multi-stage turbocharged SI powerplants for unmanned military spyplanes in the early 1980's. Now adapting hardware developed for reconaissance at 65-70 kft to the interests of atmospheric science at 80-90 kft, their efforts should yield an aero powerplant that pushes the altitude limits of subsonic air breathing propulsion.

Bents, David J.↗
NASA NTRS2005

Parametric (On-Design) Cycle Analysis for a Separate-Exhaust Turbofan Engine With Interstage Turbine Burner

Today s modern aircraft is based on air-breathing jet propulsion systems, which use moving fluids as substances to transform energy carried by the fluids into power. Throughout aero-vehicle evolution, improvements have been made to the engine efficiency and pollutants reduction. The major advantages associated with the addition of ITB are an increase in thermal efficiency and reduction in NOx emission. Lower temperature peak in the main combustor results in lower thermal NOx emission and lower amount of cooling air required. This study focuses on a parametric (on-design) cycle analysis of a dual-spool, separate-flow turbofan engine with an Interstage Turbine Burner (ITB). The ITB considered in this paper is a relatively new concept in modern jet engine propulsion. The ITB serves as a secondary combustor and is located between the high- and the low-pressure turbine, i.e., the transition duct. The objective of this study is to use design parameters, such as flight Mach number, compressor pressure ratio, fan pressure ratio, fan bypass ratio, and high-pressure turbine inlet temperature to obtain engine performance parameters, such as specific thrust and thrust specific fuel consumption. Results of this study can provide guidance in identifying the performance characteristics of various engine components, which can then be used to develop, analyze, integrate, and optimize the system performance of turbofan engines with an ITB. Visual Basic program, Microsoft Excel macrocode, and Microsoft Excel neuron code are used to facilitate Microsoft Excel software to plot engine performance versus engine design parameters. This program computes and plots the data sequentially without forcing users to open other types of plotting programs. A user s manual on how to use the program is also included in this report. Furthermore, this stand-alone program is written in conjunction with an off-design program which is an extension of this study. The computed result of a selected design-point engine will be exported to an engine reference data file that is required in off-design calculation.

Liew, K. H.↗
NASA NTRS1985

The use of superoxide mixtures as air-revitalization chemicals in hyperbaric, self-contained, closed-circuit breathing apparatus

In portable breathing apparatus applications at 1 atm, potassium superoxide (KO2) has exhibited low-utilization efficiency of the available oxygen (O2) and diminished carbon dioxide-(CO2) scrubbing capacity caused by the formation of a fused, hydrated-hydroxide/carbonate product coating on the superoxide granules. In earlier work, it was discovered that granules fabricated from an intimate mixture of KO2 and calcium superoxide, Ca(O2)2, did not exhibit formation of a fused product coating and the utilization efficiency with respect to both O2 release and CO2 absorption was superior to KO2 granules when both types of granules were reacted with humidified CO2 under identified conditions. In the work described here, single pellets of KO2, KO2/Ca(O2), mixtures and commercially available KO2 tables and granules were reacted with a flow of humidified CO2 in helium at 1- and 10-atm total pressure and at an initial temperature of 40 C. In the 1-atm flow tests, the reaction rates and utilization efficiency of the KO2/Ca(O2)2 pellets were markedly superior to the KO2 pellets, tablets, and granules when the samples were reacted under identical conditions. However, at 10 atm, the rates of O2 release and CO2 absorption, as well as the utilization efficiencies of all the superoxide samples, were one-third to one-eighth of the values observed at 1 atm. The decrease in reaction performance at 10 atm compared to that at 1 atm has been attributed principally to the lower bulk diffusivity of the CO2 and H2O reactants in helium at the higher pressure and secondarily to the moderation of the reaction temperature caused by the higher heat capacity of the 10-atm helium.

Wood, P. C.↗