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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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NASA NTRS2003

Pulse Detonation Rocket Engine Research at NASA Marshall

This viewgraph representation provides an overview of research being conducted on Pulse Detonation Rocket Engines (PDRE) by the Propulsion Research Center (PRC) at the Marshall Space Flight Center. PDREs have a theoretical thermodynamic advantage over Steady-State Rocket Engines (SSREs) although unsteady blowdown processes complicate effective use of this advantage in practice; PRE is engaged in a fundamental study of PDRE gas dynamics to improve understanding of performance issues. Topics covered include: simplified PDRE cycle, comparison of PDRE and SSRE performance, numerical modeling of quasi 1-D rocket flows, time-accurate thrust calculations, finite-rate chemistry effects in nozzles, effect of F-R chemistry on specific impulse, effect of F-R chemistry on exit species mole fractions and PDRE performance optimization studies.

Morris, Christopher I.↗

NASA NTRS2003

Hybrid Propulsion Demonstration Program 250K Hybrid Motor

The Hybrid Propulsion Demonstration Program (HPDP) program was formed to mature hybrid propulsion technology to a readiness level sufficient to enable commercialization for various space launch applications. The goal of the HPDP was to develop and test a 250,000 pound vacuum thrust hybrid booster in order to demonstrate hybrid propulsion technology and enable manufacturing of large hybrid boosters for current and future space launch vehicles. The HPDP has successfully conducted four tests of the 250,000 pound thrust hybrid rocket motor at NASA's Stennis Space Center. This paper documents the test series.

Story, George↗

NASA NTRS2003

Use of Soft Computing Technologies For Rocket Engine Control

The problem to be addressed in this paper is to explore how the use of Soft Computing Technologies (SCT) could be employed to further improve overall engine system reliability and performance. Specifically, this will be presented by enhancing rocket engine control and engine health management (EHM) using SCT coupled with conventional control technologies, and sound software engineering practices used in Marshall s Flight Software Group. The principle goals are to improve software management, software development time and maintenance, processor execution, fault tolerance and mitigation, and nonlinear control in power level transitions. The intent is not to discuss any shortcomings of existing engine control and EHM methodologies, but to provide alternative design choices for control, EHM, implementation, performance, and sustaining engineering. The approaches outlined in this paper will require knowledge in the fields of rocket engine propulsion, software engineering for embedded systems, and soft computing technologies (i.e., neural networks, fuzzy logic, and Bayesian belief networks), much of which is presented in this paper. The first targeted demonstration rocket engine platform is the MC-1 (formerly FASTRAC Engine) which is simulated with hardware and software in the Marshall Avionics & Software Testbed laboratory that

Trevino, Luis C.↗

NASA NTRS1961

Rocket-Model Investigation of Lateral Stability Characteristics and Power Effects of a Jet-Engine Airplane Configuration with Tail Boom at Mach Numbers from 1.15 to 1.37

Results from a limited research program initiated to study the effects of a hot propulsive jet on the lateral stability characteristics of a fighter-type airplane configuration are presented. The data were obtained on a rocket-boosted free-flight model and a Mach number range from 1.15 to 1.37 was covered. The configuration tested had sweptback-wing and tail surfaces and a tail boom of rectangular cross section. A solid-propellant rocket motor was used to simulate a turbojet engine with afterburner operating. Pulse rockets provided yaw disturbances during both power-on and power-off flight.

Curry, T. Bradley, Jr.↗

NASA NTRS2004

In-Space Chemical Propulsion System Model

Multiple, new technologies for chemical systems are becoming available and include high temperature rockets, very light propellant tanks and structures, new bipropellant and monopropellant options, lower mass propellant control components, and zero boil off subsystems. Such technologies offer promise of increasing the performance of in-space chemical propulsion for energetic space missions. A mass model for pressure-fed, Earth and space-storable, advanced chemical propulsion systems (ACPS) was developed in support of the NASA MSFC In-Space Propulsion Program. Data from flight systems and studies defined baseline system architectures and subsystems and analyses were formulated for parametric scaling relationships for all ACPS subsystems. The paper will first provide summary descriptions of the approaches used for the systems and the subsystems and then present selected analyses to illustrate use of the model for missions with characteristics of current interest.

Byers, David C.↗

NASA NTRS2004

In-Space Chemical Propulsion System Model

Multiple, new technologies for chemical systems are becoming available and include high temperature rockets, very light propellant tanks and structures, new bipropellant and monopropellant options, lower mass propellant control components, and zero boil off subsystems. Such technologies offer promise of increasing the performance of in-space chemical propulsion for energetic space missions. A mass model for pressure-fed, Earth and space-storable, advanced chemical propulsion systems (ACPS) was developed in support of the NASA MSFC In-Space Propulsion Program. Data from flight systems and studies defined baseline system architectures and subsystems and analyses were formulated for parametric scaling relationships for all ACPS subsystem. The paper will first provide summary descriptions of the approaches used for the systems and the subsystems and then present selected analyses to illustrate use of the model for missions with characteristics of current interest.

Byers, David C.↗

NASA NTRS2005

Liquid Rocket Engine Testing Overview

Contents include the following: Objectives and motivation for testing. Technology, Research and Development Test and Evaluation (RDT&E), evolutionary. Representative Liquid Rocket Engine (LRE) test compaigns. Apollo, shuttle, Expandable Launch Vehicles (ELV) propulsion. Overview of test facilities for liquid rocket engines. Boost, upper stage (sea-level and altitude). Statistics (historical) of Liquid Rocket Engine Testing. LOX/LH, LOX/RP, other development. Test project enablers: engineering tools, operations, processes, infrastructure.

Rahman, Shamim↗

NASA NTRS2005

SOLID SOLUTION CARBIDES ARE THE KEY FUELS FOR FUTURE NUCLEAR THERMAL PROPULSION

Nuclear thermal propulsion uses nuclear energy to directly heat a propellant (such as liquid hydrogen) to generate thrust for space transportation. In the 1960 s, the early Rover/Nuclear Engine for Rocket Propulsion Application (NERVA) program showed very encouraging test results for space nuclear propulsion but, in recent years, fuel research has been dismal. With NASA s renewed interest in long-term space exploration, fuel researchers are now revisiting the RoverMERVA findings, which indicated several problems with such fuels (such as erosion, chemical reaction of the fuel with propellant, fuel cracking, and cladding issues) that must be addressed. It is also well known that the higher the temperature reached by a propellant, the larger the thrust generated from the same weight of propellant. Better use of fuel and propellant requires development of fuels capable of reaching very high temperatures. Carbides have the highest melting points of any known material. Efforts are underway to develop carbide mixtures and solid solutions that contain uranium carbide, in order to achieve very high fuel temperatures. Binary solid solution carbides (U, Zr)C have proven to be very effective in this regard. Ternary carbides such as (U, Zr, X) carbides (where X represents Nb, Ta, W, and Hf) also hold great promise as fuel material, since the carbide mixtures in solid solution generate a very hard and tough compact material. This paper highlights past experience with early fuel materials and bi-carbides, technical problems associated with consolidation of the ingredients, and current techniques being developed to consolidate ternary carbides as fuel materials.

Panda, Binayak↗

NASA NTRS2004

The Use of RF Waves in Space Propulsion Systems

This paper will review the ways in which RF and microwave radiation may be used in the design of electric propulsion systems for spacecraft. RF power has been used or proposed in electric propulsion systems to ionize, to heat, and to accelerate the propellant, or to produce plasma used to inflate a magnetic field for solar sail purposes. Direct RF propulsion using radiation pressure or ponderomotive forces is impractical owing to efficiency considerations. Examples of various systems that have been developed or proposed will be reviewed. The Variable Specific Impulse Magnetoplasma Rocket (VASIMR) uses RF for producing, heating and accelerating plasma. Inductive RF and microwave ion thruster schemes use e-m waves to ionize the plasma, which is then accelerated by use of dc grids. The details of the VASIMR, an inductive RF thruster, and a microwave ion thruster are discussed and contrasted with related RF systems.

Bering, Edgar A., III↗

NASA NTRS2006

Design Considerations for the Nuclear Thermal Rocket Element Environmental Simulator (NTREES)

Nuclear Thermal Rockets or NTR's have been suggested as a propulsion system option for vehicles traveling to the moon or Mars. These engines are capable of providing high thrust at specific impulses at least twice that of today s best chemical engines. The performance constraints on these engines are mainly the result of temperature limitations on the fuel coupled with a limited ability to withstand chemical attack by the hot hydrogen propellant. To operate at maximum efficiency, fuel forms are desired which can withstand the extremely hot, hostile environment characteristic of NTR operation for at least several hours. The simulation of such an environment would require an experimental device which could simultaneously approximate the power, flow, and temperature conditions which a nuclear fuel element (or partial element) would encounter during NTR operation. Such a simulation would allow detailed studies of the fuel behavior and hydrogen flow characteristics under reactor like conditions to be performed. The goal of these simulations would be directed toward expanding the performance envelope of NTR engines over that which was demonstrated during the Rover and NERVA nuclear rocket programs of the 1970's. Currently, such a simulator is nearing completion at the Marshall Space Flight Center, and will shortly be used in the future to evaluate a wide variety of he1 element designs and the materials of which they are constructed. This present work addresses the initial experimental objectives of the Nuclear Thermal Rocket Element Environmental Simulator or NTREES and some of the design considerations which were considered prior to and during its construction.

Emrich, Bill↗

NASA NTRS2007

Advanced Chemical Propulsion System Study

A detailed; mission-level systems study has been performed to show the benefit resulting from engine performance gains that will result from NASA's In-Space Propulsion ROSS Cycle 3A NRA, Advanced Chemical Technology sub-topic. The technology development roadmap to accomplish the NRA goals are also detailed in this paper. NASA-Marshall and NASA-JPL have conducted mission-level studies to define engine requirements, operating conditions, and interfaces. Five reference missions have been chosen for this analysis based on scientific interest, current launch vehicle capability and trends in space craft size: a) GTO to GEO, 4800 kg, delta-V for GEO insertion only approx.1830 m/s; b) Titan Orbiter with aerocapture, 6620 kg, total delta V approx.210 m/s, mostly for periapsis raise after aerocapture; c) Enceladus Orbiter (Titan aerocapture) 6620 kg, delta V approx.2400 m/s; d) Europa Orbiter, 2170 kg, total delta V approx.2600 m/s; and e) Mars Orbiter, 2250 kg, total delta V approx.1860 m/s. The figures of merit used to define the benefit of increased propulsion efficiency at the spacecraft level include propulsion subsystem wet mass, volume and overall cost. The objective of the NRA is to increase the specific impulse of pressure-fed earth storable bipropellant rocket engines to greater than 330 seconds with nitrogen tetroxide and monomothylhydrazine propellants and greater than 335 , seconds with nitrogen tetroxide and hydrazine. Achievement of the NRA goals will significantly benefit NASA interplanetary missions and other government and commercial opportunities by enabling reduced launch weight and/or increased payload. The study also constitutes a crucial stepping stone to future development, such as pump-fed storable engines.

Portz, Ron↗

NASA NTRS2008

Status Update Report for the Peregrine 100km Sounding Rocket Project

The Peregrine Sounding Rocket Program is a joint basic research program of NASA Ames Research Center, NASA Wallops, Stanford University and the Space Propulsion Group, Inc. (SPG). The goal is to determine the applicability of liquifying hybrid technology to a small launch system. The approach is to design, build, test and y a stable, efficient liquefying fuel hybrid rocket vehicle to an altitude of 100 km. The program was kicked o in October of 2006 and has seen considerable progress in the subsequent 18 months. Two virtually identical vehicles will be constructed and own out of the NASA Sounding Rocket Facility at Wallops Island. This paper presents the current status of the project as of June 2008. For background on the project, the reader is referred to last year's paper.

Dyer, Jonny↗

NASA NTRS2010

2010 JANNAF Abstract Propulsion System Status: Ares First Stage

In November 2005, NASA created the Constellation Program to develop an entirely new fleet of spacecraft to include the Ares I Crew Launch Vehicle and the Ares V Cargo Launch vehicles. The Ares I First Stage Team has made significant progress on the design of a propulsion system capable of lofting that new fleet. To minimize technical risks and development costs, the Solid Rocket Boosters (SRBs) of Shuttle served as a starting point in the design of a new motor that would meet the requirements of those new vehicles. By bringing in a fifth segment, this new propulsive element will provide greater total impulse to loft a safer, more powerful fleet of space flight vehicles. Additional modifications include the use of composite structures, new metallic forward structures, new parachutes, a new propellant grain design, and a new nozzle. Although still ongoing and in process, significant design, development and concept testing has already been conducted. This includes a full-scale motor firing, deceleration subsystem testing, and testing of several complete avionics line replaceable units. The suborbital development flight, Ares I-X, provided the opportunity to gather critical data about the flight dynamics of the integrated launch vehicle stack. The Ares I-X test flight has also informed our Ares First Stage design. It brought a more comprehensive understanding in a number of areas, which are now being analyzed. This paper will provide an overview of the design, development, challenges, and progress on the production of the new Ares First Stage propulsion system. Challenges and mitigation strategies will be addressed. An overview will also be included of testing to date.

Tiller, Bruce K.↗

NASA NTRS2010

Rocket Engine Altitude Simulation Technologies

John C. Stennis Space Center is embarking on a very ambitious era in its rocket engine propulsion test history. The first new large rocket engine test stand to be built at Stennis Space Center in over 40 years is under construction. The new A3 Test Stand is designed to test very large (294,000 Ibf thrust) cryogenic propellant rocket engines at a simulated altitude of 100,000 feet. A3 Test Stand will have an engine testing chamber where the engine will be fired after the air in the chamber has been evacuated to a pressure at the simulated altitude of less than 0.16 PSIA. This will result in a very unique environment with extremely low pressures inside a very large chamber and ambient pressures outside this chamber. The test chamber is evacuated of air using a 2-stage diffuser / ejector system powered by 5000 lb/sec of steam produced by 27 chemical steam generators. This large amount of power and flow during an engine test will result in a significant acoustic and vibrational environment in and around A3 Test Stand.

Woods, Jody L.↗

NASA NTRS2010

Performance and Stability Analyses of Rocket Combustion Devices Using Liquid Oxygen/Liquid Methane Propellants

Liquid rocket engines using oxygen and methane propellants are being considered by the National Aeronautics and Space Administration (NASA) for in-space vehicles. This propellant combination has not been previously used in flight-qualified engine systems, so limited test data and analysis results are available at this stage of early development. NASA has funded several hardware-oriented programs with oxygen and methane propellants over the past several years with the Propulsion and Cryogenic Advanced Development (PCAD) project, under the Exploration Technology Development Program. As part of this effort, NASA Marshall Space Flight Center has conducted combustion, performance, and combustion stability analyses of several of the configurations on these programs. This paper summarizes these analyses. Test and analysis results of impinging and coaxial element injectors using liquid oxygen and liquid methane propellants are included. Several cases with gaseous methane are included for reference. Several different thrust chamber configurations have been modeled, including thrust chambers with multi-element like-on-like and swirl coax element injectors tested at NASA MSFC, and a unielement chamber with shear and swirl coax injectors tested at The Pennsylvania State University. Configurations were modeled with two one-dimensional liquid rocket combustion analysis codes, the Rocket Combustor Interaction Design and Analysis (ROCCID), and the Coaxial Injector Combustion Model (CICM). Significant effort was applied to show how these codes can be used to model combustion and performance with oxygen/methane propellants a priori, and what anchoring or calibrating features need to be applied or developed in the future. This paper describes the test hardware configurations, presents the results of all the analyses, and compares the results from the two analytical methods.

Hulka, James R.↗

NASA NTRS2006

Numerical Propulsion System Simulation

The NASA Glenn Research Center, in partnership with the aerospace industry, other government agencies, and academia, is leading the effort to develop an advanced multidisciplinary analysis environment for aerospace propulsion systems called the Numerical Propulsion System Simulation (NPSS). NPSS is a framework for performing analysis of complex systems. The initial development of NPSS focused on the analysis and design of airbreathing aircraft engines, but the resulting NPSS framework may be applied to any system, for example: aerospace, rockets, hypersonics, power and propulsion, fuel cells, ground based power, and even human system modeling. NPSS provides increased flexibility for the user, which reduces the total development time and cost. It is currently being extended to support the NASA Aeronautics Research Mission Directorate Fundamental Aeronautics Program and the Advanced Virtual Engine Test Cell (AVETeC). NPSS focuses on the integration of multiple disciplines such as aerodynamics, structure, and heat transfer with numerical zooming on component codes. Zooming is the coupling of analyses at various levels of detail. NPSS development includes capabilities to facilitate collaborative engineering. The NPSS will provide improved tools to develop custom components and to use capability for zooming to higher fidelity codes, coupling to multidiscipline codes, transmitting secure data, and distributing simulations across different platforms. These powerful capabilities extend NPSS from a zero-dimensional simulation tool to a multi-fidelity, multidiscipline system-level simulation tool for the full development life cycle.

Naiman, Cynthia↗

NASA NTRS2010

Comparisons Between Stability Prediction and Measurements for the Reusable Solid Rocket Motor

The Space Transportation System has used the solid rocket boosters for lift-off and ascent propulsion over the history of the program. Part of the structural loads assessment of the assembled vehicle is the contribution due to solid rocket booster thrust oscillations. These thrust oscillations are a consequence of internal motor pressure oscillations active during operation. Understanding of these pressure oscillations is key to predicting the subsequent thrust oscillations and vehicle loading. The pressure oscillation characteristics of the Reusable Solid Rocket Motor (RSRM) design are reviewed in this work. Dynamic pressure data from the static test and flight history are shown, with emphasis on amplitude, frequency, and timing of the oscillations. Physical mechanisms that cause these oscillations are described by comparing data observations to predictions made by the Solid Stability Prediction (SSP) code.

Fischbach, Sean R.↗

NASA NTRS2011

Pulsed Electric Propulsion Thrust Stand Calibration Method

The evaluation of the performance of any propulsion device requires the accurate measurement of thrust. While chemical rocket thrust is typically measured using a load cell, the low thrust levels associated with electric propulsion (EP) systems necessitate the use of much more sensitive measurement techniques. The design and development of electric propulsion thrust stands that employ a conventional hanging pendulum arm connected to a balance mechanism consisting of a secondary arm and variable linkage have been reported in recent publications by Polzin et al. These works focused on performing steady-state thrust measurements and employed a static analysis of the thrust stand response. In the present work, we present a calibration method and data that will permit pulsed thrust measurements using the Variable Amplitude Hanging Pendulum with Extended Range (VAHPER) thrust stand. Pulsed thrust measurements are challenging in general because the pulsed thrust (impulse bit) occurs over a short timescale (typically 1 micros to 1 millisecond) and cannot be resolved directly. Consequently, the imparted impulse bit must be inferred through observation of the change in thrust stand motion effected by the pulse. Pulsed thrust measurements have typically only consisted of single-shot operation. In the present work, we discuss repetition-rate pulsed thruster operation and describe a method to perform these measurements. The thrust stand response can be modeled as a spring-mass-damper system with a repetitive delta forcing function to represent the impulsive action of the thruster.

Wong, Andrea R.↗