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Results for “STORABLE PROPELLANT”

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 217 records · Page 12
All sourcesNASA NTRSDOE OSTI
NASA NTRS1993

Small rocket research and technology

Small chemical rockets are used on nearly all space missions. The small rocket program provides propulsion technology for civil and government space systems. Small rocket concepts are developed for systems which encompass reaction control for launch and orbit transfer systems, as well as on-board propulsion for large space systems and earth orbit and planetary spacecraft. Major roles for on-board propulsion include apogee kick, delta-V, de-orbit, drag makeup, final insertions, north-south stationkeeping, orbit change/trim, perigee kick, and reboost. The program encompasses efforts on earth-storable, space storable, and cryogenic propellants. The earth-storable propellants include nitrogen tetroxide (NTO) as an oxidizer with monomethylhydrazine (MMH) or anhydrous hydrazine (AH) as fuels. The space storable propellants include liquid oxygen (LOX) as an oxidizer with hydrazine or hydrocarbons such as liquid methane, ethane, and ethanol as fuels. Cryogenic propellants are LOX or gaseous oxygen (GOX) as oxidizers and liquid or gaseous hydrogen as fuels. Improved performance and lifetime for small chemical rockets are sought through the development of new predictive tools to understand the combustion and flow physics, the introduction of high temperature materials to eliminate fuel film cooling and its associated combustion inefficiency, and improved component designs to optimize performance. Improved predictive technology is sought through the comparison of both local and global predictions with experimental data. Results indicate that modeling of the injector and combustion process in small rockets needs improvement. High temperature materials require the development of fabrication processes, a durability data base in both laboratory and rocket environments, and basic engineering property data such as strength, creep, fatigue, and work hardening properties at both room and elevated temperature. Promising materials under development include iridium-coated rhenium and a ceramic composite of mixed hafnium carbide and tantalum carbide reinforced with graphite fibers.

Schneider, Steven↗
NASA NTRS2003

Microfabricated Liquid Rocket Motors

Under NASA Glenn Research Center sponsorship, MIT has developed the concept of micromachined, bipropellant, liquid rocket engines. This is potentially a breakthrough technology changing the cost-performance tradeoffs for small propulsion systems, enabling new applications, and redefining the meaning of the term low-cost-access-to-space. With this NASA support, a liquid-cooled, gaseous propellant version of the thrust chamber and nozzle was designed, built, and tested as a first step. DARPA is currently funding MIT to demonstrate turbopumps and controls. The work performed herein was the second year of a proposed three-year effort to develop the technology and demonstrate very high power density, regeneratively cooled, liquid bipropellant rocket engine thrust chamber and nozzles. When combined with the DARPA turbopumps and controls, this work would enable the design and demonstration of a complete rocket propulsion system. The original MIT-NASA concept used liquid oxygen-ethanol propellants. The military applications important to DARPA imply that storable liquid propellants are needed. Thus, MIT examined various storable propellant combinations including N2O4 and hydrazine, and H2O2 and various hydrocarbons. The latter are preferred since they do not have the toxicity of N2O4 and hydrazine. In reflection of the newfound interest in H2O2, it is once again in production and available commercially. A critical issue for the microrocket engine concept is cooling of the walls in a regenerative design. This is even more important at microscale than for large engines due to cube-square scaling considerations. Furthermore, the coolant behavior of rocket propellants has not been characterized at microscale. Therefore, MIT designed and constructed an apparatus expressly for this purpose. The report details measurements of two candidate microrocket fuels, JP-7 and JP-10.

Epstein, Alan H.↗
NASA NTRS2003

Study of a High-Energy Upper Stage for Future Shuttle Missions

Space Shuttle Orbiters are likely to remain in service to 2020 or beyond for servicing the International Space Station and for launching very high value spacecraft. There is a need for a new STS-deployable upper stage that can boost certain Orbiter payloads to higher energy orbits, up to and including Earth-escape trajectories. The inventory of solid rocket motor Inertial Upper Stages has been depleted, and it is unlikely that a LOX/LH2-fueled upper stage can fly on Shuttle due to safety concerns. This paper summarizes the results of a study that investigated a low cost, low risk approach to quickly developing a new large upper stage optimized to fly on the existing Shuttle fleet. Two design reference missions (DRMs) were specified: the James Webb Space Telescope (JWST) and the Space Interferometry Mission (SIM). Two categories of upper stage propellants were examined in detail: a storable liquid propellant and a storable gel propellant. Stage subsystems 'other than propulsion were based largely on heritage hardware to minimize cost, risk and development schedule span. The paper presents the ground rules and guidelines for conducting the study, the preliminary conceptual designs margins, assessments of technology readiness/risk, potential synergy with other programs, and preliminary estimates of development and production costs and schedule spans. Although the Orbiter Columbia was baselined for the study, discussion is provided to show how the results apply to the remaining STS Orbiter fleet.

Dressler, Gordon A.↗
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 NTRS1998

Evaluation of Some Candidate Propulsion Technologies for Mars Ascent Vehicle

A Mars ascent propulsion system trade study was conducted to determine 1) what propulsion technologies allow a Mars sample return mission to be launched on a Delta III class launch verhicle, and 2) whether more exotic technologies, such as in-situ propellant production, allow major cost savings by enabling the use of a smaller launch vehicle or a direct return from the Martian surface to Earth without the need of a rendezvous in the Martian orbit.

lightweight components↗
NASA NTRSDate not supplied

Updated Human Mars Ascent Vehicle Concept in Support of NASA’s Strategic Analysis Cycle 2021

The NASA Artemis program has brought significant change to the agency’s human exploration strategy over the last several years. The human exploration of Mars remains an ultimate objective of the overall human exploration strategy. However, how the agency intends to execute initial human exploration of Mars has shifted to better align with current policy. These changes find their way into the design of key architecture elements, such as the Mars Ascent Vehicle (MAV). Of the numerous changes in the conceptual human Mars architecture, two in particular have had significant impact on the design of the human MAV: the desire for minimal surface infrastructure for initial human Mars missions, and technology investment timelines in support of initial human missions to Mars. The first leads to a surface mission architecture targeting a 30-day surface duration with two crewmembers. The second drives out in-situ resource utilization (ISRU) from the initial sortie. As a result, the reference MAV design for initial human Mars missions in the Human Exploration and Operations Mission Directorate’s Strategic Analysis Cycle 2021 (SAC21) is based around supporting the return of 2 crew from the surface of Mars, without surface ISRU. This has led to several design changes since 2019, when details of a reference human MAV were last published. The MAV concept for SAC21 consists of a two-stage vehicle capable of supporting two crew for 84 hours. This duration is derived from a baseline ascent trajectory targeting an apoapsis rendezvous with the habitation element in a nominal 5-sol Mars aggregation orbit. With the goal of minimizing technology investments for the initial surface sortie removing ISRU from the trade space, the MAV must now be landed either fully loaded with propellant, or with some portion of the propellant off-loaded. Landing a partially-fueled MAV will require additional surface assets to perform robotic propellant transfer on the surface of Mars from a pre-emplaced propellant depot. Furthermore, cryogenic fluid management and storage for extended duration with no losses represents additional technology investment that would have to be made to support cryogenic-based propellant concepts that have been the baseline in past architectures. As a result, the SAC21 architecture utilizes a storable propellant-based MAV propulsion system to minimize these potential technology investments incurred by cryogenic-based propulsion systems. Recent efforts also focused on evaluating the sensitivity of key design parameters, such as landing site latitude, elevation, and local atmospheric conditions. All play a large role in determining the predicted propellant requirements of the MAV. Understanding their potential impacts is important because of the rippling effects of changes to the MAV design on other architecture elements, such as entry, descent, landing, and transportation systems. Results indicate up to +2% to -4% wet mass variation due to landing site latitude, with higher latitudes greater than 70 degrees North resulting in greater than 6% wet mass increases. The MAV saw reduced impact due to atmospheric design parameters, which only accounting for -0.6% to +0.4% wet mass variation.

Douglas J. Trent↗
NASA NTRS2014

Green Propellant Infusion Mission Program Development and Technology Maturation

The NASA Space Technology Mission Directorate's (STMD) Green Propellant Infusion Mission (GPIM) Technology Demonstration Mission (TDM) is comprised of a cross-cutting team of domestic spacecraft propulsion and storable green propellant technology experts. This TDM is led by Ball Aerospace & Technologies Corp. (BATC), who will use their BCP- 100 spacecraft to carry a propulsion system payload consisting of one 22 N thruster for primary divert (DeltaV) maneuvers and four 1 N thrusters for attitude control, in a flight demonstration of the AF-M315E technology. The GPIM project has technology infusion team members from all three major market sectors: Industry, NASA, and the Department of Defense (DoD). The GPIM project team includes BATC, includes Aerojet Rocketdyne (AR), Air Force Research Laboratory, Aerospace Systems Directorate, Edwards AFB (AFRL), NASA Glenn Research Center (GRC), NASA Kennedy Space Center (KSC), and NASA Goddard Space Flight Center (GSFC). STMD programmatic and technology oversight is provided by NASA Marshall Space Flight Center. The GPIM project shall fly an operational AF-M315E green propulsion subsystem on a Ball-built BCP-100 spacecraft.

AF-M315E↗
NASA NTRS1988

The ram accelerator - A chemically driven mass launcher

The ram accelerator, a chemically propelled mass driver, is presented as a viable new approach for directly launching acceleration-insensitive payloads into low earth orbit. The propulsion principle is similar to that of a conventional air-breathing ramjet. The cargo vehicle resembles the center-body of a ramjet and travels through a tube filled with a pre-mixed fuel and oxidizer mixture. The launch tube acts as the outer cowling of the ramjet and the combustion process travels with the vehicle. Two drive modes of the ram accelerator propulsion system are described, which when used in sequence are capable of accelerating the vehicle to as high as 10 km/sec. The requirements are examined for placing a 2000 kg vehicle into a 500 km orbit with a minimum of on-board rocket propellant for circularization maneuvers. It is shown that aerodynamic heating during atmospheric transit results in very little ablation of the nose. An indirect orbital insertion scenario is selected, utilizing a three step maneuver consisting of two burns and aerobraking. An on-board propulsion system using storable liquid propellants is chosen in order to minimize propellant mass requirements, and the use of a parking orbit below the desired final orbit is suggested as a means to increase the flexibility of the mass launch concept. A vehicle design using composite materials is proposed that will best meet the structural requirements, and a preliminary launch tube design is presented.

Kaloupis, P.↗
NASA NTRS1993

Inherent stability of central element coaxial liquid-liquid injectors

Most TRW liquid bi-propellant rocket engines built over the past thirty-plus years have employed a central element coaxial pintle injector and have operated with liquid/liquid propellant injection. This injector is a patented design exclusive to TRW and has unique features that make the rocket engine combustion characteristics different from those of other types of injector engine designs. Its many benefits include excellent combustion performance, efficient deep throttling, adaptability to low cost manufacturing, and high reliability. Approximately 200 pintle injector engines of various sizes and operating on a variety of propellants have been flown without a single inflight failure. An especially important feature of the pintle injector engine is its apparent inherent combustion stability. In over thirty years of development, testing, and production, TRW has never experienced combustion instability in any of its pintle injector engine designs. This has been true of engines operating over a range of thrust from 5 to 250,000 lbs. on earth-storable hypergolic propellants and a large number of smaller engines operating on a variety of propellants (21 combinations) in long duration-firing, pulsing (down to 2 msec), and deep throttling (as much as 19:1) modes. Operating chamber pressures have ranged from 10 to 3,500 psia. This record is particularly impressive given that typical TRW design practice does not consider combustion instability as an issue and no pintle engine has ever employed stability-enhancing features, such as baffles or acoustically resonant chambers. In spite of this, TRW engines have operated stably in regimes not possible with other types of injectors. Various physical explanations and combustion process models for this favorable stability characteristic have been postulated. However, a definitive study that unequivocally establishes the important stabilizing mechanisms still remains to be conducted.

Stoddard, Frank J.↗
NASA NTRS1973

An advanced regulator for the helium pressurization systems of the Space Shuttle OMS and RCS

The Space Shuttle Orbit Maneuvering System and Reaction Control System are pressure-fed rocket propulsion systems utilizing earth storable hypergolic propellants and featuring engines of 6000 lbs and 900 lbs thrust, respectively. The helium pressurization system requirements for these propulsion systems are defined and the current baseline pressurization systems are described. An advanced helium pressure regulator capable of meeting both OMS and RCS helium pressurization system requirements is presented and its operating characteristics and predicted performance characteristics are discussed.

Wichmann, H.↗
NASA NTRS1979

High performance N2O4/amine elements blowapart

Hypergolic earth-storable N2O4/Amine propellants are used for a wide range of liquid propulsion system applications. These propellants are highly reactive and can experience reactive stream separation (RSS, i.e., blowapart) which can inhibit intraelement mixing, hence reducing the overall spray mixture ratio and mass distribution, which can result in altered combustion efficiency, gas-side heat transfer coefficient, and stability. An experimental and analytical program was conducted to develop an understanding of the mechanisms controlling RSS. The program and product was development of design criteria for coping with RSS to allow the design of high performing, stable injectors. RSS mechanisms were identified using high speed color photography to observe reactive stream mixing of single element injectors tested with N2O4/MMH, N2O4/A-50, and N2O4/N2H4 propellants. The parametric characterization included modeling of the Space Shuttle Orbital Maneuvering System and Reaction Control System engine injectors.

Source record↗
NASA NTRS1986

Evaluation of a storable fluorine based pilot for scramjets

An ignitor/pilot device for scramjet engines has been tested with C2H4, C2H6, and CH4 gaseous hydrocarbon fuels at Mach 2. The pilot was designed to operate with storable nontoxic propellants rather than hazardous pyrophorics; the pilot fluid is a mixture of ethylene, oxygen, and sulfur hexafluoride. The pilot reduced the minimum temperature at which all three hydrocarbon fuels ignited; a comparison was then conducted between combustion efficiency calculations for these test results and those of pilots using nitrogen in place of sulfur hexafluoride, and using silane-hydrogen (yielding equivalent energy). The nitrogen-substituted pilot exhibited the highest efficiency.

Diskin, G. S.↗
NASA NTRS1999

Upper-Stage Flight Experiment

For propulsion applications that require that the propellants are storable for long periods, have a high density impulse, and are environmentally clean and non-toxic, the best choice is a combination of high-concentration hydrogen peroxide (High Test Peroxide, or HTP) and a liquid hydrocarbon (LHC) fuel. The HTP/LHC combination is suitable for low-cost launch vehicles, space taxi and space maneuvering vehicles, and kick stages. Orbital Sciences Corporation is under contract with the NASA Marshall Space Flight Center in cooperation with the Air Force Research Lab to design, develop and demonstrate a new low-cost liquid upper stage based on HTP and JP-8. The Upper Stage Flight Experiment (USFE) focuses on key technologies necessary to demonstrate the operation of an inherently simple propulsion system with an innovative, state-of-the-art structure. Two key low-cost vehicle elements will be demonstrated - a 10,000 lbf thrust engine and an integrated composite tank structure. The suborbital flight test of the USFE is scheduled for 2001. Preceding the flight tests are two major series of ground tests at NASA Stennis Space Center and a subscale tank development program to identify compatible composite materials and to verify their compatibility over long periods of time. The ground tests include a thrust chamber development test series and an integrated stage test. This paper summarizes the results from the first phase of the thrust chamber development tests and the results to date from the tank material compatibility tests. Engine and tank configurations that meet the goals of the program are described.

Anderson, W. E.↗
NASA NTRS2003

Hydrogen Peroxide - Material Compatibility Studied by Microcalorimetry

Environmental and toxicity concerns with current hypergolic propellants have led to a renewed interest in propellant grade hydrogen peroxide (HP) for propellant applications. Storability and stability has always been an issue with HP. Contamination or contact of HP with metallic surfaces may cause decomposition, which can result in the evolution of heat and gas leading to increased pressure or thermal hazards. The NASA Johnson Space Center White Sands Test Facility has developed a technique to monitor the decompositions of hydrogen peroxide at temperatures ranging from 25 to 60 C. Using isothermal microcalorimetry we have measured decomposition rates at the picomole/s/g level showing the catalytic effects of materials of construction. In this paper we will present the results of testing with Class 1 and 2 materials in 90 percent hydrogen peroxide.

Homung, Steven D.↗
NASA NTRS2018

Green Propulsion Advancement and Infusion

All space missions benefit from increased propulsion system performance, allowing lower spacecraft launch mass, larger scientific payloads, or extended on-orbit lifetimes. Likewise, long-term storable liquid propellant candidates that offer significant reduction in personnel hazards and shorter payload processing schedules present a more attractive propulsion subsystem solution to spacecraft builders. Aiming to reduce risk to potential infusion missions and fully comprehend the alternative propellant performance, the work presented herein represents many years of development and collaborative efforts to successfully align higher performance, low toxicity green propellants into NASA Goddard Space Flight Center (GSFC) missions. High Performance Green Propulsion (HPGP), and the associated propellant technology, has advanced significantly in maturity with increased familiarity with LMP-103S propellant handling, the proven reduction in loading hazards, successful launches conducted at multiple international Ranges, and HPGP on-orbit flight heritage. As science missions move forward to the potential infusion of HPGP technology, the National Aeronautics and Space Administration (NASA) and its partners are working to address gaps in system performance and operational considerations.

Spacecraft Propulsion and Power↗