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Explore propulsion reports from NASA NTRS and DOE OSTI. Search rocket engines, electric propulsion and thrusters; verify designs and performance in the source documents.

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At least 1,207 records · Page 67
All sourcesNASA NTRSDOE OSTI
NASA NTRS1997

Characterization of the breakup of the Pegasus rocket body 1994-029B

The breakup of a Pegasus hydrazine auxiliary propulsion system in June 1996, officially recognized as the worst satellite breakup in terms of cataloged debris, is considered. The fragmentation event is analyzed and it is discussed how these debris contribute to the current and future near earth space environment. The low altitude of the breakup and the large range of ejection velocities present concerns for other earth orbiting space vehicles, especially the Space Shuttle and the Hubble Space Telescope. In addition to orbit data collected by the U.S. Space Surveillance Network, observations were conducted with ground-based radar observatories. These observations show that the overabundance of debris is not limited to the trackable population, but also extends down to debris with sizes of less than 1 cm. Attempts to detect the debris with optical sensors were less successful.

Matney, Mark↗
NASA NTRS1999

Simulation of Non-Acoustic Combustion Instability in a Hybrid Rocket Motor

A transient model of a hybrid motor was formulated to study the cause and elimination of non-acoustic combustion instability. The transient model was used to simulate four key tests out of a series of seventeen hybrid motor tests conducted by Thiokol, Rocketdyne and Martin Marietta at NASA/Marshall Space Flight Center (NASA/MSFC). These tests were performed under the Hybrid Propulsion Technology for Launch Vehicle Boosters (HPTLVB) program. The first test resulted in stable combustion. The second test resulted in large-amplitude, 6.5 Hz chamber pressure oscillations that gradually damped away by the end of the test. The third test resulted in large-amplitude, 7.5 Hz chamber pressure oscillations that were sustained throughout the test. The seventh test resulted in the elimination of combustion instability with the installation of an orifice immediately upstream of the injector. The formulation and implementation of the model are the scope of this presentation. The current model is an independent continuation of modeling presented previously by joint Thiokol-Rocketdyne collaborators Boardman, Hawkins, Wassom, and Claflin. The previous model simulated an unstable IR&D hybrid motor test performed by Thiokol. There was very good agreement between the model and the test data. Like the previous model, the current model was developed using Matrix-x simulation software. However, the tests performed at NASA/MSFC under the HPTLVB program were actually simulated. In the current model, the hybrid motor consisting of the liquid oxygen (LOX) injector, the multi-port solid fuel grain and the nozzle was simulated. Also, simulated in the model was the LOX feed system consisting of the tank, venturi, valve and feed lines. All components of the hybrid motor and LOX feed system are treated by a lumped-parameter approach. Agreement between the results of the transient model and the actual test data was very good. This agreement between simulated and actual test data indicated that the combustion instability in the hybrid motor was due to two causes. The first cause was a LOX feed system of insufficient stiffness. The second cause was a LOX injector with an impedance or pressure drop that was too low to provide damping against the feed system oscillations. Also, it was discovered that testing with a new grain of solid fuel sustained the combustion instability. However, testing with a used grain of solid fuel caused the combustion instability to gradually decay.

Rocker, Marvin↗
NASA NTRS1999

Simulation of Non-Acoustic Combustion Instability in a Hybrid Rocket Motor

A transient model of a hybrid motor was formulated to study the cause and elimination of non-acoustic combustion instability. The transient model was used to simulate four key tests out of a series of seventeen hybrid motor tests conducted by Thiokol, Rocketdyne and Martin Marietta at NASA/Marshall Space Flight Center (NASAIMSFC). These tests were performed under the Hybrid Propulsion Technology for Launch Vehicle Boosters (HPTLVB) program. The first test resulted in stable combustion. The second test resulted in large-amplitude, 6.5 Hz chamber pressure oscillations that gradually damped away by the end of the test. The third test resulted in large-amplitude, 7.5 Hz chamber pressure oscillations that were sustained throughout the test. The seventh test resulted in the elimination of combustion instability with the installation of an orifice immediately upstream of the injector. The formulation and implementation of the model are the scope of this presentation. The current model is an independent continuation of modeling presented previously by joint Thiokol-Rocketdyne collaborators Boardman, Hawkins, Wassom, and Claflin. The previous model simulated an unstable IR&D hybrid motor test performed by Thiokol. There was very good agreement between the model and the test data. Like the previous model, the current model was developed using Matrix-x simulation software. However, the tests performed at NASA/MSFC under the HPTLVB program were actually simulated. In the current model, the hybrid motor consisting of the liquid oxygen (LOX) injector, the multi-port solid fuel grain and the nozzle was simulated. Also, simulated in the model was the LOX feed system consisting of the tank, venturi, valve and feed lines. All components of the hybrid motor and LOX feed system are treated by a lumped-parameter approach. Agreement between the results of the transient model and the actual test data was very good. This agreement between simulated and actual test data indicated that the combustion instability in the hybrid motor was due to two causes. The first cause was a LOX feed system of insufficient stiffness. The second cause was a LOX injector with an impedance or pressure drop that was too low to provide damping against the feed system oscillations. Also, it was discovered that testing with a new grain of solid fuel sustained the combustion instability. However, testing with a used grain of solid fuel caused the combustion instability to gradually decay.

Rocker, Marvin↗
NASA NTRS2000

Preliminary Analysis and Design of a Rocket Based Combined Cycle for Efficient Access to Space

Access to space is presently limited by cost. The cost of delivering a payload to low earth orbit (LEO) is on average $10,000 per pound of payload, in the United States. Much of this cost is incurred from the operation of vehicles developed with 30-40 year old technology. The old technology and design practices have resulted in expensive hardware and intensive maintenance requirements for current launch vehicles. In order to alleviate the cost factor, the technological advances throughout the next millennium must bring affordable development and a new invigorating desire to space exploration. National Aeronautical and Space Administration (NASA), Department of Defense (DOD), and private industry are addressing this issue by focusing on incremental improvements in the Earth- to Orbit (ETO) costs. These improvements have investigated two different approaches: 1) make space vehicles as inexpensive as possible (i.e. Evolved Expendable Launch Vehicle (EELV) and Delta IV) 2) make space vehicles as reusable as airplanes so that the initial cost of investment can be recaptured (i.e. Reusable Launch Vehicle (RLV), X-33, X-34, and X-37) These programs have made notable progress in new material, propulsion, structures, and avionics technologies, during the last 3-5 years. So far, these programs are targeted to reduce present costs by, as much as, five times the current cost. The year 2025 goal for continued space advancement is to have ETO costs reduced by a factor of ten (i.e., tenfold), as low as $100 - $200/lb payload. For the RLV, this goal translates into very low maintenance costs and higher expected reliability per flight must be obtained. Therefore, making higher launch rates possible. For the expendable vehicle, the cost of maintenance cost is minimal, but a greater reliability must exist to insure the payload; since, there would be no way to recover the payload if the mission was to an abort or failure. Overall the cost of vehicle, payload and operations of an expendable may be too high compared to those of the RLV.

Harris, Lawanna↗
NASA NTRS2001

Robust Nonlinear Feedback Control of Aircraft Propulsion Systems

This is the final report on the research performed under NASA Glen grant NASA/NAG-3-1975 concerning feedback control of the Pratt & Whitney (PW) STF 952, a twin spool, mixed flow, after burning turbofan engine. The research focussed on the design of linear and gain-scheduled, multivariable inner-loop controllers for the PW turbofan engine using H-infinity and linear, parameter-varying (LPV) control techniques. The nonlinear turbofan engine simulation was provided by PW within the NASA Rocket Engine Transient Simulator (ROCETS) simulation software environment. ROCETS was used to generate linearized models of the turbofan engine for control design and analysis as well as the simulation environment to evaluate the performance and robustness of the controllers. Comparison between the H-infinity, and LPV controllers are made with the baseline multivariable controller and developed by Pratt & Whitney engineers included in the ROCETS simulation. Simulation results indicate that H-infinity and LPV techniques effectively achieve desired response characteristics with minimal cross coupling between commanded values and are very robust to unmodeled dynamics and sensor noise.

Garrard, William L.↗
NASA NTRS2001

Blood Pump Development Using Rocket Engine Flow Simulation Technology

This paper reports the progress made towards developing complete blood flow simulation capability in humans, especially in the presence of artificial devices such as valves and ventricular assist devices. Devices modeling poses unique challenges different from computing the blood flow in natural hearts and arteries. There are many elements needed to quantify the flow in these devices such as flow solvers, geometry modeling including flexible walls, moving boundary procedures and physiological characterization of blood. As a first step, computational technology developed for aerospace applications was extended to the analysis and development of a ventricular assist device (VAD), i.e., a blood pump. The blood flow in a VAD is practically incompressible and Newtonian, and thus an incompressible Navier-Stokes solution procedure can be applied. A primitive variable formulation is used in conjunction with the overset grid approach to handle complex moving geometry. The primary purpose of developing the incompressible flow analysis capability was to quantify the flow in advanced turbopump for space propulsion system. The same procedure has been extended to the development of NASA-DeBakey VAD that is based on an axial blood pump. Due to massive computing requirements, high-end computing is necessary for simulating three-dimensional flow in these pumps. Computational, experimental, and clinical results are presented.

Kwak, Dochan↗
NASA NTRS2002

Minimally Intrusive and Nonintrusive Supersonic Injectors for LANTR and RBCC/Scramjet Propulsion Systems

A family of supersonic injectors for use on spaceplanes, rockets and missiles and the like is disclosed and claimed. Each injector maintains a specific constant (uniform) Mach number along its length when used while being minimally intrusive at significantly higher injectant pressure than combuster freestream total pressure. Each injector is substantially non-intrusive when it is not being used. The injectors may be used individually or in a group. Different orientations of the injectors in a group promotes greater penetration and mixing of fuel or oxidizer into a supersonic combustor. The injectors can be made from single piece of Aluminum, investment cast metal, or ceramic or they can be made from starboard and port blocks strapped together to accurately control the throat area. Each injector includes an elongated body having an opening which in cross section is an hour glass (venturi shaped) and the opening diverges in width and depth from the bow section to the stem section of the opening.

Alvin E Buggele↗
NASA NTRS2003

Asteroids as Propulsion Systems of Space Ships

Currently, rockets are used to change the trajectory of space ships and probes. This method is very expensive and requires a lot of fuel, which limits the feasibility of space stations, interplanetary space ships, and probes. Sometimes space probes use the gravity field of a planet However, there am only nine planets in the Solar System, all separated by great distances. There are tons of millions of asteroids in outer space. This paper offers a revolutionary method for changing the trajectory of space probes. The method uses the kinetic or rotary energy of asteroids, comet nuclei, meteorites or other space bodies (small planets, natural planetary satellites, space debris, etc.) to increase (to decrease) ship (probe) speed up to 1000 m/sec (or more) and to achieve any new direction in outer space. The flight possibilities of space ships and probes are increased by a factor of millions.

Bolonkin, Alexander↗
NASA NTRS2003

High Conductivity Materials for High Heat Flux Applications in Space Propulsion Systems

GRCop-84 is a new copper base alloy with an excellent combination of strength and conductivity and has been developed to the point that it is a strong candidate for near term rocket engine applications. This work sought to establish the feasibility of new alloys with capabilities beyond GRCop-84. The use of coatings as environmental and thermal barriers adds further capability into the Copper base structure. Finally, Nickel-Aluminum based systems have also been explored.

Nathal, M. V.↗
NASA NTRS1957

Steady Nuclear Combustion in Rockets

The astrophysical theory of stationary nuclear reactions in stars is applied to the conditions that would be met in the practical engineering cases that would differ from the former, particularly with respect to the much lower combustion pressures, dimensions of the reacting volume, and burnup times. This application yields maximum rates of hear production per unit volume of reacting gas occurring at about 10(exp 8) K in the cases of reactions between the hydrogen isotopes, but yields higher rates for heavier atoms. For the former, with chamber pressures of the order of 100 atmospheres, the energy production for nuclear combustion reaches values of about 10(exp 4) kilocalories per cubic meter per second, which approaches the magnitude for the familiar chemical fuels. The values are substantially lower for heavier atoms, and increase with the square of the combustion pressure. The half-life of the burnup in the fastest reactions may drop to values as low as those for chemical fuels so that, despite the high temperature, the radiated energy can remain smaller than the energy produced, particularly if an inefficiently radiating (i.e., easily completely ionized reacting material like hydrogen), is used. On the other hand, the fraction of completely ionized particles in the gases undergoing nuclear combustion must not exceed a certain upper limit because the densities (approximately 10(exp -10) grams per cubic centimeter)) lie in the range of high vacua and only for the previously mentioned fraction of nonionized particles can mean free paths be retained small enough so that the chamber diameters of several dozen meters will suffice. Under these conditions it appears that continuously maintained stable nuclear reactions at practical pressures and dimensions are fundamentally possible and their application can be visualized as energy sources for power plants and propulsion units.

Saenger, E.↗
NASA NTRS2004

RTF Analysis-Propulsion/Flowliner

The objective of the current task is to provide a computational framework for design and analysis of the entire fuel supply system of a liquid rocket engine, including high fidelity unsteady turbopump flow analysis. Time-accurate results obtained from shuttle fuel flowliner analysis using 66 Million grid points with 262 overlapped zones will be shown. We will present analysis results along with performance data of the simulation runs on Supercomputers such as Columbia.

Kiris, Centin↗
NASA NTRS2005

Development of Ionic Liquid Monopropellants for In-Space Propulsion

A family of new, low toxicity, high energy monopropellants is currently being evaluated at NASA Marshall Space Flight Center for in-space rocket engine applications such as reaction control engines. These ionic liquid monopropellants, developed in recent years by the Air Force Research Laboratory, could offer system simplification, less in-flight thermal management, and reduced handling precautions, while increasing propellant energy density as compared to traditional storable in-space propellants such as hydrazine and nitrogen tetroxide. However, challenges exist in identifying ignition schemes for these ionic liquid monopropellants, which are known to burn at much hotter combustion temperatures compared to traditional monopropellants such as hydrazine. The high temperature combustion of these new monopropellants make the use of typical ignition catalyst beds prohibitive since the catalyst cannot withstand the elevated temperatures. Current research efforts are focused on monopropellant ignition and burn rate characterization, parameters that are important in the fundamental understanding of the monopropellant behavior and the eventual design of a thruster. Laboratory studies will be conducted using alternative ignition techniques such as laser-induced spark ignition and hot wire ignition. Ignition delay, defined as the time between the introduction of the ignition source and the first sign of light emission from a developing flame kernel, will be measured using Schlieren visualization. An optically-accessible liquid monopropellant burner, shown schematically in Figure 1 and similar in design to apparatuses used by other researchers to study solid and liquid monopropellants, will be used to determine propellant burn rate as a function of pressure and initial propellant temperature. The burn rate will be measured via high speed imaging through the chamber s windows.

Blevins, John A.↗
NASA NTRS2005

Reliability Improvements in Liquid Rocket Engine Instrumentation

Instrumentation hardware is often the weak link in advanced liquid fueled propulsion systems. The development of the Space Shuttle Main Engine (SSME) was no exception. By sheer necessity, a reusable, high energy, low weight engine system often relegates the instrumentation hardware to the backseat in the critical hardware development process. This produces less than optimum hardware constraints; including size, location, mounting, redundancy, and signal conditioning. This can negatively affect the development effort and ultimately the system reliability. The challenge was clear, however, the outcome was less certain. Unfortunately, the SSME hardware development culminated in series of measurement failures, most significant of which was the premature engine shutdown during the launch of STS-51F on July 29, 1985. The Return to Flight activities following the Challenger disaster redoubled our efforts to eliminate, once and for all, sensor malfunctions as the determining factor in overall engine reliability. This paper describes each phase of this effort in detail and includes discussion of the tasks related to improving measurement reliability.

Hill, A.↗
NASA NTRS2005

REIMR: A Process for Utilizing Propulsion-Oriented 'Lessons-Learned' to Mitigate Development Risk

This paper is a summary overview of a study conducted a t the NASA Marshall Space Flight Center (MSFC) during the initial phases of the Space Launch Initiative (SLI) program to evaluate a large number of technical problems associated with the design, development, test, evaluation and operation of several major liquid propellant rocket engine systems (i.e., SSME, Fastrac, J-2, F-1). The results of this study was the identification of the "Fundamental Root Causes" that enabled the technical problems to manifest, and practices that can be implemented to prevent them from recurring in future engine development efforts. This paper will discus the Fundamental Root Causes, cite some examples of how the technical problems arose from them, and provide a discussion of how they can be mitigated or avoided.

Ballard, Richard O.↗
NASA NTRS2005

Development of Ionic Liquid Monopropellants for In-Space Propulsion

A family of new, low toxicity, high energy monopropellants is currently being evaluated at NASA Marshall Space Flight Center for in-space rocket engine applications such as reaction control engines. These ionic liquid monopropellants, developed in recent years by the Air Force Research Laboratory, could offer system simplification, less in-flight thermal management, and reduced handling precautions, while increasing propellant energy density as compared to traditional storable in-space propellants such as hydrazine and nitrogen tetroxide. However, challenges exist in identifying ignition schemes for these ionic liquid monopropellants, which are known to burn at much hotter combustion temperatures compared to traditional monopropellants such as hydrazine. The high temperature combustion of these new monopropellants make the use of typical ignition catalyst beds prohibitive since the catalyst cannot withstand the elevated temperatures. Current research efforts are focused on monopropellant ignition and burn rate characterization, parameters that are important in the fundamental understanding of the monopropellant behavior and the eventual design of a thruster. Laboratory studies will be conducted using alternative ignition techniques such as laser-induced spark ignition and hot wire ignition. Ignition delay, defined as the time between the introduction of the ignition source and the first sign of light emission from a developing flame kernel, will be measured using Schlieren visualization. An optically-accessible liquid monopropellant burner will be used to determine propellant burn rate as a function of pressure and initial propellant temperature. The burn rate will be measured via high speed imaging through the chamber s windows.

Blevins, John A.↗
NASA NTRS2005

Using CFD as a Rocket Injector Design Tool: Recent Progress at Marshall Space Flight Center

New programs are forcing American propulsion system designers into unfamiliar territory. For instance, industry s answer to the cost and reliability goals set out by the Next Generation Launch Technology Program are engine concepts based on the Oxygen- Rich Staged Combustion Cycle. Historical injector design tools are not well suited for this new task. The empirical correlations do not apply directly to the injector concepts associated with the ORSC cycle. These legacy tools focus primarily on performance with environment evaluation a secondary objective. Additionally, the environmental capability of these tools is usually one-dimensional while the actual environments are at least two- and often three-dimensional. CFD has the potential to calculate performance and multi-dimensional environments but its use in the injector design process has been retarded by long solution turnaround times and insufficient demonstrated accuracy. This paper has documented the parallel paths of program support and technology development currently employed at Marshall Space Flight Center in an effort to move CFD to the forefront of injector design. MSFC has established a long-term goal for use of CFD for combustion devices design. The work on injector design is the heart of that vision and the Combustion Devices CFD Simulation Capability Roadmap that focuses the vision. The SRL concept, combining solution fidelity, robustness and accuracy, has been established as a quantitative gauge of current and desired capability. Three examples of current injector analysis for program support have been presented and discussed. These examples are used to establish the current capability at MSFC for these problems. Shortcomings identified from this experience are being used as inputs to the Roadmap process. The SRL evaluation identified lack of demonstrated solution accuracy as a major issue. Accordingly, the MSFC view of code validation and current MSFC-funded validation efforts were discussed in some detail. The objectives of each effort were noted. Issues relative to code validation for injector design were discussed in some detail. The requirement for CFD support during the design of the experiment was noted and discussed in terms of instrumentation placement and experimental rig uncertainty. In conclusion, MSFC has made significant progress in the last two years in advancing CFD toward the goal of application to injector design. A parallel effort focused on program support and technology development via the SCIT Task have enabled the progress.

Tucker, Kevin↗