Search NASA⌕ Search

SEARCH · Search NASA

Results for “PROPELLANT OXIDIZER”

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.

Quote a phrase for an exact phrase match. Source license links do not imply unrestricted reuse.

At least 145 records · Page 8
All sourcesNASA NTRSDOE OSTI
NASA NTRS1993

Advanced Small Rocket Chambers. Basic Program and Option 2: Fundamental Processes and Material Evaluation

Propellants, chamber materials, and processes for fabrication of small high performance radiation cooled liquid rocket engines were evaluated to determine candidates for eventual demonstration in flight-type thrusters. Both storable and cryogenic propellant systems were considered. The storable propellant systems chosen for further study were nitrogen tetroxide oxidizer with either hydrazine or monomethylhydrazine as fuel. The cryogenic propellants chosen were oxygen with either hydrogen or methane as fuel. Chamber material candidates were chemical vapor deposition (CVD) rhenium protected from oxidation by CVD iridium for the chamber hot section, and film cooled wrought platinum-rhodium or regeneratively cooled stainless steel for the front end section exposed to partially reacted propellants. Laser diagnostics of the combustion products near the hot chamber surface and measurements at the surface layer were performed in a collaborative program at Sandia National Laboratories, Livermore, CA. The Material Sample Test Apparatus, a laboratory system to simulate the combustion environment in terms of gas and material temperature, composition, and pressure up to 6 Atm, was developed for these studies. Rocket engine simulator studies were conducted to evaluate the materials under simulated combustor flow conditions, in the diagnostic test chamber. These tests used the exhaust species measurement system, a device developed to monitor optically species composition and concentration in the chamber and exhaust by emission and absorption measurements.

Jassowski, Donald M.↗
NASA NTRS1996

Testing of electroformed deposited iridium/powder metallurgy rhenium rockets

High-temperature, oxidation-resistant chamber materials offer the thermal margin for high performance and extended lifetimes for radiation-cooled rockets. Rhenium (Re) coated with iridium (Ir) allow hours of operation at 2200 C on Earth-storable propellants. One process for manufacturing Ir/Re rocket chambers is the fabrication of Re substrates by powder metallurgy (PM) and the application of Ir coatings by using electroformed deposition (ED). ED Ir coatings, however, have been found to be porous and poorly adherent. The integrity of ED Ir coatings could be improved by densification after the electroforming process. This report summarizes the testing of two 22-N, ED Ir/PM Re rocket chambers that were subjected to post-deposition treatments in an effort to densify the Ir coating. One chamber was vacuum annealed, while the other chamber was subjected to hot isostatic pressure (HIP). The chambers were tested on gaseous oxygen/gaseous hydrogen propellants, at mixture ratios that simulated the oxidizing environments of Earth-storable propellants. ne annealed ED Ir/PM Re chamber was tested for a total of 24 firings and 4.58 hr at a mixture ratio of 4.2. After only 9 firings, the annealed ED Ir coating began to blister and spall upstream of the throat. The blistering and spalling were similar to what had been experienced with unannealed, as-deposited ED Ir coatings. The HIP ED Ir/PM Re chamber was tested for a total of 91 firings and 11.45 hr at mixture ratios of 3.2 and 4.2. The HIP ED Ir coating remained adherent to the Re substrate throughout testing; there were no visible signs of coating degradation. Metallography revealed, however, thinning of the HIP Ir coating and occasional pores in the Re layer upstream of the throat. Pinholes in the Ir coating may have provided a path for oxidation of the Re substrate at these locations. The HIP ED Ir coating proved to be more effective than vacuum annealed and as-deposited ED Ir. Further densification is still required to match the integrity of chemically vapor deposited Ir coatings. Despite this, the successful long duration testing of the HIP ED Ir chamber, in an oxidizing environment comparable to Earth-storable propellants, demonstrated the viability of this Ir/Re rocket fabrication process.

Reed, Brian D.↗
NASA NTRS2004

Removal of Perfluorinated Grease Components from NTO Oxidizer

Perfluorinated greases are typically used as a thread lubricant in the assembly of non-welded nitrogen tetroxide (NTO) oxidizer systems. These greases, typically a perfluoroalkylether, with suspended polytetrafluoroethylene (PTFE) micro-powder, have attractive lubricating properties toward threaded components and are relatively chemically inert toward NTO oxidizers. A major drawback, however, is that perfluoroalkylether greases are soluble or dispersible in NTO oxidizers and can contaminate the propellant. The result is propellant that fails the non-volatile residue (NVR) specification analyses and that may have negative effects on test hardware performance and lifetime. Consequently, removal of the grease contaminants from NTO may be highly desirable. Methods for the removal of perfluorinated grease components from NTO oxidizers including distillation, adsorption, filtration, and adjustment of temperature are investigated and reported in this work. Solubility or dispersibility data for the perfluoroalkylether oil (Krytox(tm)143 AC) component of a perfluorinated grease (Krytox 240 AC) and for Krytox 240 AC in NTO were determined and are reported.

McClure, Mark B.↗
NASA NTRS2015

Steam Methane Reformation Testing for Air-Independent Solid Oxide Fuel Cell Systems

Recently, NASA has been looking into utilizing landers that can be propelled by LOX-CH (sub 4), to be used for long duration missions. Using landers that utilize such propellants, also provides the opportunity to use solid oxide fuel cells as a power option, especially since they are able to process methane into a reactant through fuel reformation. One type of reformation, called steam methane reformation, is a process to reform methane into a hydrogen-rich product by reacting methane and steam (fuel cell exhaust) over a catalyst. A steam methane reformation system could potentially use the fuel cell's own exhaust to create a reactant stream that is hydrogen-rich, and requires less internal reforming of the incoming methane. Also, steam reformation may hold some advantages over other types of reforming, such as partial oxidation (PROX) reformation. Steam reformation does not require oxygen, while up to 25 percent can be lost in PROX reformation due to unusable CO (sub 2) reformation. NASA's Johnson Space Center has conducted various phases of steam methane reformation testing, as a viable solution for in-space reformation. This has included using two different types of catalysts, developing a custom reformer, and optimizing the test system to find the optimal performance parameters and operating conditions.

Mwara, Kamwana N.↗
NASA NTRS1990

RSRM-13 (360Q013) ballistics mass properties flight designation STS-41

The propulsion performance and reconstructed mass properties data from Thiokol's RSRM-13 motors which were assigned to the STS-41 launch are presented. The SRM propellant, TP-H1148, is a composite type solid propellant, formulated of polybutadiene acrylic acid acryonitrile terpolymer binder, epoxy curing agent, ammonium perchlorate oxidizer, and aluminum powder fuel. A small amount of burning rate catalyst (iron oxide) was added to achieve the desired propellant burn rate. The propellant evaluation and raw material information are also presented. The presented ballistic performance was based on the Operational Flight Instrumentation. The adjustments made to the raw data on this flight include biasing the data to correct ambient pressure before liftoff. The performance from each motor as well as matched pair performance values were well within the CEI Specification requirements.

Laubacher, Brian A.↗
NASA NTRS1991

RSRM-11 (36OW011) ballistics mass properties (STS-35)

The propulsion performance and reconstructed mass properties data from Thiolol's RSRM-11 motors which were assigned to the STS-35 launch are contained. The Thiokol manufacturing designations for the motors were 360W011A/360W011B, which are referred to as RSRM-11A and RSRM-B, respectively. The launch of STS-35 occurred on 2 December 1990 at the Eastern Test Range (ETR). The data contained herein was input to the STS-35 Flight Evaluation Report. The SRM propellant, TP-H1148, is a composite type solid propellants, formulated of polybutediene acrylic acid, acryonitrile terpolymer binder, epoxy curing agent, ammonium perchlorate oxidizer, and aluminum powder fuel. A small amount of burning rate catalyst (iron oxide) was added to achieve the desired propellant burn rate. The propellant evaluation and raw material information for the RSRM-11 are included. The ballistic performance presented was based on the Operational Flight Instrumentation (OFI) 12.5 sample per second pressure data for the steady state and tail off portion of the pressure trace. Recent studies have shown that the transducers are affected by the measuring system at KSC and temperature gradients created by the igniter heaters. Therefore, an adjustment to the data from each transducer is made to make the initial reading match the atmospheric pressure at the time of launch.

Hutchinson, B. J.↗
NASA NTRS1993

Design, analysis, and fabrication of oxide-coated iridium/rhenium combustion chambers

Iridium-coated rhenium (Ir/Re) combustion chambers provide high temperature, oxidation-resistant operation for radiation-cooled liquid-fueled rocket engines. A 22-N (5-lb(sub f)) chamber has been operated for 15 hours at 2200 C (4000 F) using nitrogen tetroxide/monomethyl hydrazine (NTO/MMH) propellant, with negligible internal erosion. The oxidation resistance of these chambers could be further increased by the addition of refractory oxide coatings, providing longer life and/or operation in more oxidizing and higher temperature environments. The oxide coatings would serve as a thermal and diffusion barrier for the iridium coating, lowering the temperature of the iridium layer while also preventing the ingress of oxygen and egress of iridium oxides. This would serve to slow the failure mechanisms of Ir/Re chambers, namely the diffusion of rhenium to the inner surface and the oxidation of iridium. Such protection could extend chamber lifetimes by tens or perhaps hundreds of hours, and allow chamber operation on stoichiometric or higher mixture ratio oxygen/hydrogen (O2/H2) propellant. Extensive thermomechanical, thermochemical, and mass transport modeling was performed as a key material/structure design tool. Based on the results of these analyses, several 22-N oxide-coated Ir/Re chambers were fabricated and delivered to NASA Lewis Research Center for hot-fire testing.

Jang, Q.↗
NASA NTRS1986

The effect of Mars surface and Phobos propellant production on Earth launch mass

Fuel and oxidizer produced on the surface of Mars and on the Martian Moon Phobos can reduce the cumulative mass of fuel and oxidizer which must be launched to low Earth orbit for Mars exploration missions. A scenario in which ten conjunction class trajectory missions over a twenty year period land a surface base and propellant production facilities on the Martian surface and on Phobos was examined. Production of oxygen on Phobos provides the greatest benefit. If all the propellant for Mars operations and Earth return is produced at Phobos and on Mars, a 30% reduction in cumulative low Earth orbit mass can be achieved at the end of the 20 year period.

Gus R Babb↗
NASA NTRS2018

Chapter 12: Materials for Liquid Propulsion Systems

Earth to orbit launch vehicles are propelled by rocket engines and motors, both liquid and solid. This chapter will discuss liquid engines. The heart of a launch vehicle is its engine. The remainder of the vehicle (with the notable exceptions of the payload and guidance system) is an aero structure to support the propellant tanks which provide the fuel and oxidizer to feed the engine or engines. The basic principle behind a rocket engine is straightforward. The engine is a means to convert potential thermochemical energy of one or more propellants into exhaust jet kinetic energy. Fuel and oxidizer are burned in a combustion chamber where they create hot gases under high pressure. These hot gases are allowed to expand through a nozzle. The molecules of hot gas are first constricted by the throat of the nozzle (de-Laval nozzle) which forces them to accelerate; then as the nozzle flares outwards, they expand and further accelerate. It is the mass of the combustion gases times their velocity, reacting against the walls of the combustion chamber and nozzle, which produce thrust according to Newton's third law: for every action there is an equal and opposite reaction. Solid rocket motors are cheaper to manufacture and offer good values for their cost. Liquid propellant engines offer higher performance, that is, they deliver greater thrust per unit weight of propellant burned. They also have a considerably higher thrust to weigh ratio. Since liquid rocket engines can be tested several times before flight, they have the capability to be more reliable, and their ability to shut down once started provides an extra margin of safety. Liquid propellant engines also can be designed with restart capability to provide orbital maneuvering capability. In some instances, liquid engines also can be designed to be reusable. On the solid side, hybrid solid motors also have been developed with the capability to stop and restart. Solid motors are covered in detail in chapter 11. Liquid rocket engine operational factors can be described in terms of extremes: temperatures ranging from that of liquid hydrogen (-423 F) to 6000 F hot gases; enormous thermal shock (7000 F/sec); large temperature differentials between contiguous components; reactive propellants; extreme acoustic environments; high rotational speeds for turbo machinery and extreme power densities. These factors place great demands on materials selection and each must be dealt with while maintaining an engine of the lightest possible weight. This chapter will describe the design considerations for the materials used in the various components of liquid rocket engines and provide examples of usage and experiences in each.

Halchak, John A.↗
NASA NTRS1990

RSRM-9 (360L009): Ballistics mass properties

The propulsion performance and reconstructed mass properties data from Thiokol's RSRM-9 motors which were assigned to the STS-36 launch are presented. The SRM propellant, TP-H1148, is a composite type solid propellant, formulated of polybutadiene acrylic acid acryonitrile terpolymer binder (PBAN), epoxy curing agent, ammonium perchlorate oxidizer and aluminum powder fuel. A small amount of burning rate catalyst (iron oxide) was added to achieve the desired propellant burn rate. The propellant evaluation and raw material information for the RSRM-9 are included. The propellant grain design consists of four segments. There is a forward segment with an eleven point star with a transition into a tapered circular perforated (CP) configuration. There are two center segments that result in a double tapered CP configuration and an aft segment with a triple taper CP configuration and a cutout for the partially submerged nozzle. The ballistic performance presented is based on the Operational Flight Instrumentation (OFI) 12.5 sample per second pressure data for the steady state and tail off portion of the pressure trace. No high sample rate pressure gauges, Development Flight Instrumentation (DFI), were used on this flight and therefore no ignition data is presented.

Drendel, Albert S.↗
NASA NTRSDate not supplied

Analysis of Oxygen Liquefaction with Transient Flow Rates for ISRU Systems

Human occupation of lunar and Martian surfaces requires in-situ resource utilization (ISRU) to create a sustainable environment with the limited resources available in space. Fuel and oxidizer generation is essential for developing a refueling capability for spacecraft, significantly reducing the propellant mass required for landing. Liquefaction systems are a key step in producing and storing cryogenic liquids such as oxygen, hydrogen, and methane. All those fluids are critical to propulsion, life support, and other spacecraft systems. NASA has previously confirmed the presence of water on the Moon and an electrolysis process can be used to separate the oxygen from the hydrogen molecules. Oxygen has also been found within lunar regolith and can be separated through ISRU processes. The extracted oxygen gas can then be liquefied inside a storage tank with tube-on-tank heat exchangers for future use as an oxidizer for propellants. While steady state liquefaction performance is easy to model, power limitations and cyclical environments caused by the change in Sun exposure between day and night periods potentially call for transient operations. The following analysis was performed using the Thermal Desktop software to investigate how various transient gaseous oxygen (GOX) flow rates impact the production and storage of liquid oxygen (LOX) in a 1-g liquefaction system. Variations of sinusoidal and exponential step functions were selected to model potential GOX flow rates that could be experienced by ISRU systems on lunar or Martian surfaces. This early computational analysis provides some insight on how transient operations impact oxygen liquefaction systems and helps illustrate operational questions to explore experimentally throughout the design process.

Cryogenic Fluid Management↗
NASA NTRS2004

Advanced Chemical Propulsion Study

A study was performed of advanced chemical propulsion technology application to space science (Code S) missions. The purpose was to begin the process of selecting chemical propulsion technology advancement activities that would provide greatest benefits to Code S missions. Several missions were selected from Code S planning data, and a range of advanced chemical propulsion options was analyzed to assess capabilities and benefits re these missions. Selected beneficial applications were found for higher-performing bipropellants, gelled propellants, and cryogenic propellants. Technology advancement recommendations included cryocoolers and small turbopump engines for cryogenic propellants; space storable propellants such as LOX-hydrazine; and advanced monopropellants. It was noted that fluorine-bearing oxidizers offer performance gains over more benign oxidizers. Potential benefits were observed for gelled propellants that could be allowed to freeze, then thawed for use.

Woodcock, Gordon↗
NASA NTRS2004

Advanced Chemical Propulsion Study

A study was performed of advanced chemical propulsion technology application to space science (Code S) missions. The purpose was to begin the process of selecting chemical propulsion technology advancement activities that would provide greatest benefits to Code S missions. Several missions were selected from Code S planning data, and a range of advanced chemical propulsion options was analyzed to assess capabilities and benefits re these missions. Selected beneficial applications were found for higher-performing bipropellants, gelled propellants, and cryogenic propellants. Technology advancement recommendations included cryocoolers and small turbopump engines for cryogenic propellants; space storable propellants such as LOX-hydrazine; and advanced monopropellants. It was noted that fluorine-bearing oxidizers offer performance gains over more benign oxidizers. Potential benefits were observed for gelled propellants that could be allowed to freeze, then thawed for use.

Woodcock, Gordon↗
NASA NTRS1990

Advanced launch vehicle upper stages using liquid propulsion and metallized propellants

Metallized propellants are liquid propellants with a metal additive suspended in a gelled fuel or oxidizer. Typically, aluminum particles are the metal additives. These propellants provide increase in the density and/or the specific impulse of the propulsion system. Using metallized propellants for volume- and mass-constrained upper stages can deliver modest increases in performance for Low Earth Orbit to Geosynchronous Earth Orbit and other Earth orbital transfer missions. Metallized propellants, however, can enable very fast planetary missions with a single-stage upper stage system. Trade studies comparing metallized propellant stage performance with non-metallized upper stages and the Inertial Upper Stage are presented. These upper stages are both one- and two-stage vehicles that provide the added energy to send payloads to altitudes and onto trajectories that are unattainable with only the launch vehicle. The stage designs are controlled by the volume and the mass constraints of the Space Transportation System and Space Transportation System-Cargo launch vehicles. The influences of the density and specific impulse increases enabled by metallized propellants are examined for a variety of different stage and propellant combinations.

Palaszewski, B. A.↗