Microfabricated Liquid Rocket Motors

NASA Technical Reports Server (NTRS)

Epstein, Alan H.; Joppin, C.; Kerrebrock, J. L.; Schneider, Steven J. (Technical Monitor)

2003-01-01

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.

KSC-2012-6224

NASA Image and Video Library

2012-11-09

CAPE CANAVERAL, Fla. -- At the Neo Liquid Propellant Testbed inside a facility near Kennedy Space Center’s Shuttle Landing Facility in Florida, engineers are working on the buildup of the Neo test fixture and an Injector 71 engine that uses super-cooled propellants. NASA engineers are working on the design and assembly of the Neo Liquid Propellant Testbed as part of the Engineering Directorate’s Rocket University training program. Photo credit: NASA/Frankie Martin

KSC-2012-6223

NASA Image and Video Library

2012-11-09

CAPE CANAVERAL, Fla. -- At the Neo Liquid Propellant Testbed inside a facility near Kennedy Space Center’s Shuttle Landing Facility in Florida, engineers are working on the buildup of the Neo test fixture and an Injector 71 engine that uses super-cooled propellants. NASA engineers are working on the design and assembly of the Neo Liquid Propellant Testbed as part of the Engineering Directorate’s Rocket University training program. Photo credit: NASA/Frankie Martin

Lessons Learned with Metallized Gelled Propellants

NASA Technical Reports Server (NTRS)

1996-01-01

During testing of metallized gelled propellants in a rocket engine, many changes had to be made to the normal test program for traditional liquid propellants. The lessons learned during the testing and the solutions for many of the new operational conditions posed with gelled fuels will help future programs run more smoothly. The major factors that influenced the success of the testing were propellant settling, piston-cylinder tank operation, control of self pressurization, capture of metal oxide particles, and a gelled-fuel protective layer. In these ongoing rocket combustion experiments at the NASA Lewis Research Center, metallized, gelled liquid propellants are used in a small modular engine that produces 30 to 40 lb of thrust. Traditional liquid RP-1 and gelled RP-1 with 0-, 5-, and 55-wt% loadings of aluminum are used with gaseous oxygen as the oxidizer. The figure compares the thrust chamber efficiencies of different engines.

Liquid rocket engine fluid-cooled combustion chambers

NASA Technical Reports Server (NTRS)

1972-01-01

A monograph on the design and development of fluid cooled combustion chambers for liquid propellant rocket engines is presented. The subjects discussed are (1) regenerative cooling, (2) transpiration cooling, (3) film cooling, (4) structural analysis, (5) chamber reinforcement, and (6) operational problems.

Liquid propellant rocket combustion instability

NASA Technical Reports Server (NTRS)

Harrje, D. T.

1972-01-01

The solution of problems of combustion instability for more effective communication between the various workers in this field is considered. The extent of combustion instability problems in liquid propellant rocket engines and recommendations for their solution are discussed. The most significant developments, both theoretical and experimental, are presented, with emphasis on fundamental principles and relationships between alternative approaches.

Prediction of high frequency combustion instability in liquid propellant rocket engines

NASA Technical Reports Server (NTRS)

Kim, Y. M.; Chen, C. P.; Ziebarth, J. P.; Chen, Y. S.

1992-01-01

The present use of a numerical model developed for the prediction of high-frequency combustion stabilities in liquid propellant rocket engines focuses on (1) the overall behavior of nonlinear combustion instabilities (2) the effects of acoustic oscillations on the fuel-droplet vaporization and combustion process in stable and unstable engine operating conditions, oscillating flowfields, and liquid-fuel trajectories during combustion instability, and (3) the effects of such design parameters as inlet boundary conditions, initial spray conditions, and baffle length. The numerical model has yielded predictions of the tangential-mode combustion instability; baffle length and droplet size variations are noted to have significant effects on engine stability.

Cryogenic Impinging Jets Subjected to High Frequency Transverse Acoustic Forcing in a High Pressure Environment

DTIC Science & Technology

2016-07-27

for liquid propellant atomization in rocket engines1- 2. Liquid rocket engines like the F-1 have successfully used like-on-like impinging jet...impingement of the two cylindrical jets. Another drawback, perhaps the most critical, is that rocket engine using impinging jets sacrifice performance in...The experimental results also suggested that impact waves seem to dominate the atomization process over most of the conditions relevant to rocket

Concept of a self-pressurized feed system for liquid rocket engines and its fundamental experiment results

NASA Astrophysics Data System (ADS)

Matsumoto, Jun; Okaya, Shunichi; Igoh, Hiroshi; Kawaguchi, Junichiro

2017-04-01

A new propellant feed system referred to as a self-pressurized feed system is proposed for liquid rocket engines. The self-pressurized feed system is a type of gas-pressure feed system; however, the pressurization source is retained in the liquid state to reduce tank volume. The liquid pressurization source is heated and gasified using heat exchange from the hot propellant using a regenerative cooling strategy. The liquid pressurization source is raised to critical pressure by a pressure booster referred to as a charger in order to avoid boiling and improve the heat exchange efficiency. The charger is driven by a part of the generated pressurization gas using a closed-loop self-pressurized feed system. The purpose of this study is to propose a propellant feed system that is lighter and simpler than traditional gas pressure feed systems. The proposed system can be applied to all liquid rocket engines that use the regenerative cooling strategy. The concept and mathematical models of the self-pressurized feed system are presented first. Experiment results for verification are then shown and compared with the mathematical models.

Hybrid propulsion technology program: Phase 1, volume 4

NASA Technical Reports Server (NTRS)

Claflin, S. E.; Beckman, A. W.

1989-01-01

The use of a liquid oxidizer-solid fuel hybrid propellant combination in booster rocket motors appears extremely attractive due to the integration of the best features of liquid and solid propulsion systems. The hybrid rocket combines the high performance, clean exhaust, and safety of liquid propellant engines with the low cost and simplicity of solid propellant motors. Additionally, the hybrid rocket has unique advantages such as an inert fuel grain and a relative insensitivity to fuel grain and oxidizer injection anomalies. The advantages mark the hybrid rocket as a potential replacement or alternative for current and future solid propellant booster systems. The issues are addressed and recommendations are made concerning oxidizer feed systems, injectors, and ignition systems as related to hybrid rocket propulsion. Early in the program a baseline hybrid configuration was established in which liquid oxygen would be injected through ports in a solid fuel whose composition is based on hydroxyl terminated polybutadiene (HTPB). Liquid oxygen remained the recommended oxidizer and thus all of the injector concepts which were evaluated assumed only liquid would be used as the oxidizer.

38th JANNAF Combustion Subcommittee Meeting. Volume 1

NASA Technical Reports Server (NTRS)

Fry, Ronald S. (Editor); Eggleston, Debra S. (Editor); Gannaway, Mary T. (Editor)

2002-01-01

This volume, the first of two volumes, is a collection of 55 unclassified/unlimited-distribution papers which were presented at the Joint Army-Navy-NASA-Air Force (JANNAF) 38th Combustion Subcommittee (CS), 26 th Airbreathing Propulsion Subcommittee (APS), 20th Propulsion Systems Hazards Subcommittee (PSHS), and 21 Modeling and Simulation Subcommittee. The meeting was held 8-12 April 2002 at the Bayside Inn at The Sandestin Golf & Beach Resort and Eglin Air Force Base, Destin, Florida. Topics cover five major technology areas including: 1) Combustion - Propellant Combustion, Ingredient Kinetics, Metal Combustion, Decomposition Processes and Material Characterization, Rocket Motor Combustion, and Liquid & Hybrid Combustion; 2) Liquid Rocket Engines - Low Cost Hydrocarbon Liquid Rocket Engines, Liquid Propulsion Turbines, Liquid Propulsion Pumps, and Staged Combustion Injector Technology; 3) Modeling & Simulation - Development of Multi- Disciplinary RBCC Modeling, Gun Modeling, and Computational Modeling for Liquid Propellant Combustion; 4) Guns Gun Propelling Charge Design, and ETC Gun Propulsion; and 5) Airbreathing - Scramjet an Ramjet- S&T Program Overviews.

In-situ propellant rocket engines for Mars missions ascent vehicle

NASA Technical Reports Server (NTRS)

Roncace, Elizabeth A.

1991-01-01

When contemplating the human exploration of Mars, many scenarios using various propulsion systems have been considered. One propulsion option among them is a vehicle stage with multiple, pump fed rocket engines capable of operating on propellants available on Mars. This reduces the earth launch mass requirements, resulting in economic and payload benefits. No plentiful sources of hydrogen on Mars have been identified on the surface of Mars, so most commonly used high performance liquid fuels, such as hydrogen and hydrocarbons, can be eliminated as possible in situ propellants. But 95 pct of the Martian atmosphere consists of carbon dioxide, which can be converted into carbon monoxide and oxygen. The carbon monoxide oxygen propellant combination is a candidate for a Martian in situ propellant rocket engine. The feasibility is analyzed of a pump fed engine cycle using the propellant combination of carbon monoxide and oxygen.

In-situ propellant rocket engines for Mars mission ascent vehicle

NASA Technical Reports Server (NTRS)

Roncace, Elizabeth A.

1991-01-01

When comtemplating the human exploration of Mars, many scenarios using various propulsion systems have been considered. One propulsion option among them is a vehicle stage with multiple, pump fed rocket engines capable of operating on propellants available on Mars. This reduces the Earth launch mass requirements, resulting in economic and payload benefits. No plentiful sources of hydrogen on Mars have been identified on the surface of Mars, so most commonly used high performance liquid fuels, such as hydrogen and hydrocarbons, can be eliminated as possible in-situ propellants. But 95 pct. of the Martian atmosphere consists of carbon dioxide, which can be converted into carbon monoxide and oxygen. The carbon monoxide oxygen propellant conbination is a candidate for a Martian in-situ propellant rocket engine. The feasibility is analyzed of a pump fed engine cycle using the propellant combination of carbon monoxide and oxygen.

Liquid rocket valve assemblies

NASA Technical Reports Server (NTRS)

1973-01-01

The design and operating characteristics of valve assemblies used in liquid propellant rocket engines are discussed. The subjects considered are as follows: (1) valve selection parameters, (2) major design aspects, (3) design integration of valve subassemblies, and (4) assembly of components and functional tests. Information is provided on engine, stage, and spacecraft checkout procedures.

Liquid fuel injection elements for rocket engines

NASA Technical Reports Server (NTRS)

Cox, George B., Jr. (Inventor)

1993-01-01

Thrust chambers for liquid propellant rocket engines include three principal components. One of these components is an injector which contains a plurality of injection elements to meter the flow of propellants at a predetermined rate, and fuel to oxidizer mixture ratio, to introduce the mixture into the combustion chamber, and to cause them to be atomized within the combustion chamber so that even combustion takes place. Evolving from these injectors are tube injectors. These tube injectors have injection elements for injecting the oxidizer into the combustion chamber. The oxidizer and fuel must be metered at predetermined rates and mixture ratios in order to mix them within the combustion chamber so that combustion takes place smoothly and completely. Hence tube injectors are subject to improvement. An injection element for a liquid propellant rocket engine of the bipropellant type is provided which includes tangential fuel metering orifices, and a plurality of oxidizer tube injection elements whose injection tubes are also provided with tangential oxidizer entry slots and internal reed valves.

Liquid Rocket Lines, Bellows, Flexible Hoses, and Filters

NASA Technical Reports Server (NTRS)

1977-01-01

Fluid-flow components in a liquid propellant rocket engine and the rocket vehicle which it propels are interconnected by lines, bellows, and flexible hoses. Elements involved in the successful design of these components are identified and current technologies pertaining to these elements are reviewed, assessed, and summarized to provide a technology base for a checklist of rules to be followed by project managers in guiding a design or assessing its adequacy. Recommended procedures for satisfying each of the design criteria are included.

Analytical and experimental studies of impinging liquid jets

NASA Technical Reports Server (NTRS)

Ryan, H. M.; Anderson, W. E.; Pal, S.; Santoro, R. J.

1994-01-01

Impinging injectors are a common type of injector used in liquid propellant rocket engines and are typically used in engines where both propellants are injected as a liquid, e.g., engines using LOX/hydrocarbon and storable propellant combinations. The present research program is focused on providing the requisite fundamental understanding associated with impinging jet injectors for the development of an advanced a priori combustion stability design analysis capability. To date, a systematic study of the atomization characteristics of impinging liquid jets under cold-flow conditions have been completed. Effects of orifice diameter, impingement angle, pre-impingement length, orifice length-to-diameter ratio, fabrication procedure, jet flow condition and jet velocity under steady and oscillating, and atmospheric- and high-pressure environments have been investigated. Results of these experimental studies have been compared to current models of sheet breakup and drop formation. In addition, the research findings have been scrutinized to provide a fundamental explanation for a proven empirical correlation used in the design of stable impinging injector-based rocket engines.

High-speed schlieren imaging of rocket exhaust plumes

NASA Astrophysics Data System (ADS)

Coultas-McKenney, Caralyn; Winter, Kyle; Hargather, Michael

2016-11-01

Experiments are conducted to examine the exhaust of a variety of rocket engines. The rocket engines are mounted in a schlieren system to allow high-speed imaging of the engine exhaust during startup, steady state, and shutdown. A variety of rocket engines are explored including a research-scale liquid rocket engine, consumer/amateur solid rocket motors, and water bottle rockets. Comparisons of the exhaust characteristics, thrust and cost for this range of rockets is presented. The variety of nozzle designs, target functions, and propellant type provides unique variations in the schlieren imaging.

Liquid Methane/Liquid Oxygen Propellant Conditioning Feed System (PCFS) Test Rigs

NASA Technical Reports Server (NTRS)

Skaff, A.; Grasl, S.; Nguyen, C.; Hockenberry S.; Schubert, J.; Arrington, L.; Vasek, T.

2008-01-01

As part of their Propulsion and Cryogenic Advanced Development (PCAD) program, NASA has embarked upon an effort to develop chemical rocket engines which utilize non-toxic, cryogenic propellants such as liquid oxygen (LO2) and liquid methane (LCH4). This effort includes the development and testing of a 100 lbf Reaction Control Engine (RCE) that will be used to evaluate the performance of a LO2/LCH4 rocket engine over a broad range of propellant temperatures and pressures. This testing will take place at NASA-Glenn Research Center's (GRC) Research Combustion Laboratory (RCL) test facility in Cleveland, OH, and is currently scheduled to begin in late 2008. While the initial tests will be performed at sea level, follow-on testing will be performed at NASA-GRC's Altitude Combustion Stand (ACS) for altitude testing. In support of these tests, Sierra Lobo, Inc. (SLI) has designed, developed, and fabricated two separate portable propellant feed systems under the Propellant Conditioning and Feed System (PCFS) task: one system for LCH4, and one for LO2. These systems will be capable of supplying propellants over a large range of conditions from highly densified to several hundred pounds per square inch (psi) saturated. This paper presents the details of the PCFS design and explores the full capability of these propellant feed systems.

Swirl-Stabilized Injector Flow and Combustion Dynamics for Liquid Propellants at Supercritical Conditions

DTIC Science & Technology

2007-02-08

was employed to study the vapor cavitation during liquid carbon dioxide expansion through a sharp-orifice nozzle. Numerical experiments demonstrated...Combustion Dynamics for 6b. GRANT NUMBER Liquid Propellants at Supercritical Conditions FA9550-04-1-0014 5c. PROGRAM ELEMENT NUMBER 6. AUTHOR(S) 5d. PROJECT...fundamental knowledge of supercritical combustion of liquid propellants under conditions representative of contemporary rocket engines. Both shear and

PC programs for the prediction of the linear stability behavior of liquid propellant propulsion systems and application to current MSFC rocket engine test programs, volume 1

NASA Technical Reports Server (NTRS)

Doane, George B., III; Armstrong, W. C.

1990-01-01

Research on propulsion stability (chugging and acoustic modes), and propellant valve control was investigated. As part of the activation of the new liquid propulsion test facilities, it is necessary to analyze total propulsion system stability. To accomplish this, several codes were built to run on desktop 386 machines. These codes enable one to analyze the stability question associated with the propellant feed systems. In addition, further work was adapted to this computing environment and furnished along with other codes. This latter inclusion furnishes those interested in high frequency oscillatory combustion behavior (that does not couple to the feed system) a set of codes for study of proposed liquid rocket engines.

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

NASA Technical Reports Server (NTRS)

Hulka, James R.; Jones, G. W.

2010-01-01

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.

Theoretical Studies of Ionic Liquids and Nanoclusters as Hybrid Fuels

DTIC Science & Technology

2016-08-17

Acknowledgements Distribution A: Approved for Public Release; Distribution Unlimited. PA# 16409 Aerospace Systems Directorate RQ-West (EAFB, CA)  Rocket ...Engines & Motors  Satellite Propulsion  Combustion Devices  Fuels and Propellants  System Analysis  R&D Rocket Testing RQ-East (WPAFB, OH)  Air...Distribution A: Approved for Public Release; Distribution Unlimited. PA# 16409 5 Identify and develop advanced chemical propellants for rocket

Cold Flow Propulsion Test Complex Pulse Testing

NASA Technical Reports Server (NTRS)

McDougal, Kris

2016-01-01

When the propellants in a liquid rocket engine burn, the rocket not only launches and moves in space, it causes forces that interact with the vehicle itself. When these interactions occur under specific conditions, the vehicle's structures and components can become unstable. One instability of primary concern is termed pogo (named after the movement of a pogo stick), in which the oscillations (cycling movements) cause large loads, or pressure, against the vehicle, tanks, feedlines, and engine. Marshall Space Flight Center (MSFC) has developed a unique test technology to understand and quantify the complex fluid movements and forces in a liquid rocket engine that contribute strongly to both engine and integrated vehicle performance and stability. This new test technology was established in the MSFC Cold Flow Propulsion Test Complex to allow injection and measurement of scaled propellant flows and measurement of the resulting forces at multiple locations throughout the engine.

The hard start phenomena in hypergolic engines. Volume 1: Bibliography

NASA Technical Reports Server (NTRS)

Miron, Y.; Perlee, H. E.

1974-01-01

A bibliography of reports pertaining to the hard start phenomenon in attitude control rocket engines on Apollo spacecraft is presented. Some of the subjects discussed are; (1) combustion of hydrazine, (2) one dimensional theory of liquid fuel rocket combustion, (3) preignition phenomena in small pulsed rocket engines, (4) experimental and theoretical investigation of the fluid dynamics of rocket combustion, and (5) nonequilibrium combustion and nozzle flow in propellant performance.

Design of a 2000 lbf LOX/LCH4 Throttleable Rocket Engine for a Vertical Lander

NASA Astrophysics Data System (ADS)

Lopez, Israel

Liquid oxygen (LOX) and liquid methane (LCH4) has been recognized as an attractive rocket propellant combination because of its in-situ resource utilization (ISRU) capabilities, namely in Mars. ISRU would allow launch vehicles to carry greater payloads and promote missions to Mars. This has led to an increasing interest to develop spacecraft technologies that employ this propellant combination. The UTEP Center for Space Exploration and Technology Research (cSETR) has focused part of its research efforts to developing LOX/LCH4 systems. One of those projects includes the development of a vertical takeoff and landing vehicle called JANUS. This vehicle will employ a LOX/LCH 4 propulsion system. The main propulsion engine is called CROME-X and is currently being developed as part of this project. This rocket engine will employ LOX/LCH4 propellants and is intended to operate from 2000-500 lbf thrust range. This thesis describes the design and development of CROME-X. Specifically, it describes the design process for the main engine components, the design criteria for each, and plans for future engine development.

Space shuttle with common fuel tank for liquid rocket booster and main engines (supertanker space shuttle)

NASA Technical Reports Server (NTRS)

Thorpe, Douglas G.

1991-01-01

An operation and schedule enhancement is shown that replaces the four-body cluster (Space Shuttle Orbiter (SSO), external tank, and two solid rocket boosters) with a simpler two-body cluster (SSO and liquid rocket booster/external tank). At staging velocity, the booster unit (liquid-fueled booster engines and vehicle support structure) is jettisoned while the remaining SSO and supertank continues on to orbit. The simpler two-bodied cluster reduces the processing and stack time until SSO mate from 57 days (for the solid rocket booster) to 20 days (for the liquid rocket booster). The areas in which liquid booster systems are superior to solid rocket boosters are discussed. Alternative and future generation vehicles are reviewed to reveal greater performance and operations enhancements with more modifications to the current methods of propulsion design philosophy, e.g., combined cycle engines, and concentric propellant tanks.

Development of Life Prediction Capabilities for Liquid Propellant Rocket Engines. Task 4. Post-Fire Diagnostic System for the SSME System Architecture Study.

DTIC Science & Technology

1991-07-31

90 START MCC LN CAV PR 3 UNDERSHOOT ABOVE THRESHOLD YES MI A2-492 2/13/90 MAINSTAGE HPOT DS TMP CHANNEL A/B DIVERGENCE NO MI A2-492 2/13/90 MAINSTAGE ...System for the SSME System Architecture Study Y, , Contract NAS 3 -25883 JUL 31󈧣 CR-187112 Prepared for: National Aeronautics and Space...Liquid Propellant Rocket Engines Contract No. NAS 3 -25883 Eli Ki ,,, July 31, 1991 BY Dist Prepared By.: Mr. Mark Gage Aerojet Propulsion Division Box

Control Room at the NACA’s Rocket Engine Test Facility

NASA Image and Video Library

1957-05-21

Test engineers monitor an engine firing from the control room of the Rocket Engine Test Facility at the National Advisory Committee for Aeronautics (NACA) Lewis Flight Propulsion Laboratory. The Rocket Engine Test Facility, built in the early 1950s, had a rocket stand designed to evaluate high-energy propellants and rocket engine designs. The facility was used to study numerous different types of rocket engines including the Pratt and Whitney RL-10 engine for the Centaur rocket and Rocketdyne’s F-1 and J-2 engines for the Saturn rockets. The Rocket Engine Test Facility was built in a ravine at the far end of the laboratory because of its use of the dangerous propellants such as liquid hydrogen and liquid fluorine. The control room was located in a building 1,600 feet north of the test stand to protect the engineers running the tests. The main control and instrument consoles were centrally located in the control room and surrounded by boards controlling and monitoring the major valves, pumps, motors, and actuators. A camera system at the test stand allowed the operators to view the tests, but the researchers were reliant on data recording equipment, sensors, and other devices to provide test data. The facility’s control room was upgraded several times over the years. Programmable logic controllers replaced the electro-mechanical control devices. The new controllers were programed to operate the valves and actuators controlling the fuel, oxidant, and ignition sequence according to a predetermined time schedule.

Liquid rocket performance computer model with distributed energy release

NASA Technical Reports Server (NTRS)

Combs, L. P.

1972-01-01

Development of a computer program for analyzing the effects of bipropellant spray combustion processes on liquid rocket performance is described and discussed. The distributed energy release (DER) computer program was designed to become part of the JANNAF liquid rocket performance evaluation methodology and to account for performance losses associated with the propellant combustion processes, e.g., incomplete spray gasification, imperfect mixing between sprays and their reacting vapors, residual mixture ratio striations in the flow, and two-phase flow effects. The DER computer program begins by initializing the combustion field at the injection end of a conventional liquid rocket engine, based on injector and chamber design detail, and on propellant and combustion gas properties. It analyzes bipropellant combustion, proceeding stepwise down the chamber from those initial conditions through the nozzle throat.

Space Shuttle with rail system and aft thrust structure securing solid rocket boosters to external tank

NASA Technical Reports Server (NTRS)

Vonpragenau, G. L. (Inventor)

1984-01-01

The configuration and relationship of the external propellant tank and solid rocket boosters of space transportation systems such as the space shuttle are described. The space shuttle system with the improved propellant tank is shown. The external tank has a forward pressure vessel for liquid hydrogen and an aft pressure vessel for liquid oxygen. The solid rocket boosters are joined together by a thrust frame which extends across and behind the external tank. The thrust of the orbiter's main rocket engines are transmitted to the aft portion of the external tank and the thrust of the solid rocket boosters are transmitted to the aft end of the external tank.

Computational analysis of liquid hypergolic propellant rocket engines

NASA Technical Reports Server (NTRS)

Krishnan, A.; Przekwas, A. J.; Gross, K. W.

1992-01-01

The combustion process in liquid rocket engines depends on a number of complex phenomena such as atomization, vaporization, spray dynamics, mixing, and reaction mechanisms. A computational tool to study their mutual interactions is developed to help analyze these processes with a view of improving existing designs and optimizing future designs of the thrust chamber. The focus of the article is on the analysis of the Variable Thrust Engine for the Orbit Maneuvering Vehicle. This engine uses a hypergolic liquid bipropellant combination of monomethyl hydrazine as fuel and nitrogen tetroxide as oxidizer.

Scaling of Performance in Liquid Propellant Rocket Engine Combustors

NASA Technical Reports Server (NTRS)

Hulka, James R.

2007-01-01

This paper discusses scaling of combustion and combustion performance in liquid propellant rocket engine combustion devices. In development of new combustors, comparisons are often made between predicted performance in a new combustor and measured performance in another combustor with different geometric and thermodynamic characteristics. Without careful interpretation of some key features, the comparison can be misinterpreted and erroneous information used in the design of the new device. This paper provides a review of this performance comparison, including a brief review of the initial liquid rocket scaling research conducted during the 1950s and 1960s, a review of the typical performance losses encountered and how they scale, a description of the typical scaling procedures used in development programs today, and finally a review of several historical development programs to see what insight they can bring to the questions at hand.

Scaling of Performance in Liquid Propellant Rocket Engine Combustion Devices

NASA Technical Reports Server (NTRS)

Hulka, James R.

2008-01-01

This paper discusses scaling of combustion and combustion performance in liquid propellant rocket engine combustion devices. In development of new combustors, comparisons are often made between predicted performance in a new combustor and measured performance in another combustor with different geometric and thermodynamic characteristics. Without careful interpretation of some key features, the comparison can be misinterpreted and erroneous information used in the design of the new device. This paper provides a review of this performance comparison, including a brief review of the initial liquid rocket scaling research conducted during the 1950s and 1960s, a review of the typical performance losses encountered and how they scale, a description of the typical scaling procedures used in development programs today, and finally a review of several historical development programs to see what insight they can bring to the questions at hand.

Ricardo Dyrgalla (1910-1970), pioneer of rocket development in Argentina

NASA Astrophysics Data System (ADS)

de León, Pablo

2009-12-01

One of the most important developers of liquid propellant rocket engines in Argentina was Polish-born Ricardo Dyrgalla. Dyrgalla immigrated to Argentina from the United Kingdom in 1946, where he had been studying German weapons development at the end of the Second World War. A trained pilot and aeronautical engineer, he understood the intricacies of rocket propulsion and was eager to find practical applications to his recently gained knowledge. Dyrgalla arrived in Argentina during Juan Perón's first presidency, a time when technicians from all over Europe were being recruited to work in various projects for the recently created Argentine Air Force. Shortly after immigrating, Dyrgalla proposed to develop an advanced air-launched weapon, the Tábano, based on a rocket engine of his design, the AN-1. After a successful development program, the Tábano was tested between 1949 and 1951; however, the project was canceled by the government shortly after. Today, the AN-1 rocket engine is recognized as the first liquid propellant rocket to be developed in South America. Besides the AN-1, Dyrgalla also developed several other rockets systems in Argentina, including the PROSON, a solid-propellant rocket launcher developed by the Argentine Institute of Science and Technology for the Armed Forces (CITEFA). In the late 1960s, Dyrgalla and his family relocated to Brazil due mostly to the lack of continuation of rocket development in Argentina. There, he worked for the Institute of Aerospace Technology (ITA) until his untimely death in 1970. Ricardo Dyrgalla deserves to be recognized among the world's rocket pioneers and his contribution to the science and engineering of rocketry deserves a special place in the history of South America's rocketry and space flight advocacy programs.

Independent Review of the Failure Modes of F-1 Engine and Propellants System

NASA Technical Reports Server (NTRS)

Ray, Paul

2003-01-01

The F-1 is the powerful engine, that hurdled the Saturn V launch vehicle from the Earth to the moon on July 16,1969. The force that lifted the rocket overcoming the gravitational force during the first stage of the flight was provided by a cluster of five F-1 rocket engines, each of them developing over 1.5 million pounds of thrust (MSFC-MAN-507). The F-1 Rocket engine used RP-1 (Rocket Propellant-1, commercially known as Kerosene), as fuel with lox (liquid Oxygen) as oxidizer. NASA terminated Saturn V activity and has focused on Space Shuttle since 1972. The interest in rocket system has been revived to meet the National Launch System (NLS) program and a directive from the President to return to the Moon and exploration of the space including Mars. The new program Space Launch Initiative (SLI) is directed to drastically reduce the cost of flight for payloads, and adopt a reusable launch vehicle (RLV). To achieve this goal it is essential to have the ability of lifting huge payloads into low earth orbit. Probably requiring powerful boosters as strap-ons to a core vehicle, as was done for the Saturn launch vehicle. The logic in favor of adopting Saturn system, a proven technology, to meet the SLI challenge is very strong. The F-1 engine was the largest and most powerful liquid rocket engine ever built, and had exceptional performance. This study reviews the failure modes of the F-1 engine and propellant system.

Testing of a Liquid Oxygen/Liquid Methane Reaction Control Thruster in a New Altitude Rocket Engine Test Facility

NASA Technical Reports Server (NTRS)

Meyer, Michael L.; Arrington, Lynn A.; Kleinhenz, Julie E.; Marshall, William M.

2012-01-01

A relocated rocket engine test facility, the Altitude Combustion Stand (ACS), was activated in 2009 at the NASA Glenn Research Center. This facility has the capability to test with a variety of propellants and up to a thrust level of 2000 lbf (8.9 kN) with precise measurement of propellant conditions, propellant flow rates, thrust and altitude conditions. These measurements enable accurate determination of a thruster and/or nozzle s altitude performance for both technology development and flight qualification purposes. In addition the facility was designed to enable efficient test operations to control costs for technology and advanced development projects. A liquid oxygen-liquid methane technology development test program was conducted in the ACS from the fall of 2009 to the fall of 2010. Three test phases were conducted investigating different operational modes and in addition, the project required the complexity of controlling propellant inlet temperatures over an extremely wide range. Despite the challenges of a unique propellant (liquid methane) and wide operating conditions, the facility performed well and delivered up to 24 hot fire tests in a single test day. The resulting data validated the feasibility of utilizing this propellant combination for future deep space applications.

Liquid rocket valve components

NASA Technical Reports Server (NTRS)

1973-01-01

A monograph on valves for use with liquid rocket propellant engines is presented. The configurations of the various types of valves are described and illustrated. Design criteria and recommended practices for the various valves are explained. Tables of data are included to show the chief features of valve components in use on operational vehicles.

Future space transport

NASA Technical Reports Server (NTRS)

Grishin, S. D.; Chekalin, S. V.

1984-01-01

Prospects for the mastery of space and the basic problems which must be solved in developing systems for both manned and cargo spacecraft are examined. The achievements and flaws of rocket boosters are discussed as well as the use of reusable spacecraft. The need for orbiting satellite solar power plants and related astrionics for active control of large space structures for space stations and colonies in an age of space industrialization is demonstrated. Various forms of spacecraft propulsion are described including liquid propellant rocket engines, nuclear reactors, thermonuclear rocket engines, electrorocket engines, electromagnetic engines, magnetic gas dynamic generators, electromagnetic mass accelerators (rail guns), laser rocket engines, pulse nuclear rocket engines, ramjet thermonuclear rocket engines, and photon rockets. The possibilities of interstellar flight are assessed.

Development of Mechanics in Support of Rocket Technology in Ukraine

NASA Astrophysics Data System (ADS)

Prisnyakov, Vladimir

2003-06-01

The paper analyzes the advances of mechanics made in Ukraine in resolving various problems of space and rocket technology such as dynamics and strength of rockets and rocket engines, rockets of different purpose, electric rocket engines, and nonstationary processes in various systems of rockets accompanied by phase transitions of working media. Achievements in research on the effect of vibrations and gravitational fields on the behavior of space-rocket systems are also addressed. Results obtained in investigating the reliability and structural strength durability conditions for nuclear installations, solid- and liquid-propellant engines, and heat pipes are presented

Combustion and Performance Analyses of Coaxial Element Injectors with Liquid Oxygen/Liquid Methane Propellants

NASA Technical Reports Server (NTRS)

Hulka, J. R.; Jones, G. W.

2010-01-01

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 a flight-qualified engine system, so limited test data and analysis results are available at this stage of early development. NASA has funded several hardware-oriented activities 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, the NASA Marshall Space Flight Center has conducted combustion, performance, and combustion stability analyses of several of the configurations. This paper summarizes the analyses of combustion and performance as a follow-up to a paper published in the 2008 JANNAF/LPS meeting. Combustion stability analyses are presented in a separate paper. The current paper includes test and analysis results of coaxial element injectors using liquid oxygen and liquid methane or gaseous methane propellants. Several thrust chamber configurations have been modeled, including thrust chambers with multi-element swirl coax element injectors tested at the NASA MSFC, and a uni-element 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

Thermo-mechanical concepts applied to modeling liquid propellant rocket engine stability

NASA Astrophysics Data System (ADS)

Kassoy, David R.; Norris, Adam

2016-11-01

The response of a gas to transient, spatially distributed energy addition can be quantified mathematically using thermo-mechanical concepts available in the literature. The modeling demonstrates that the ratio of the energy addition time scale to the acoustic time scale of the affected volume, and the quantity of energy added to that volume during the former determine the whether the responses to heating can be described as occurring at nearly constant volume, fully compressible or nearly constant pressure. Each of these categories is characterized by significantly different mechanical responses. Application to idealized configurations of liquid propellant rocket engines provides an opportunity to identify physical conditions compatible with gasdynamic disturbances that are sources of engine instability. Air Force Office of Scientific Research.

ARC-1980-AC80-0107-4

NASA Image and Video Library

1980-02-06

Outfitting the Space Shuttle Orbiter Columbia with the three main rocket engines that will boost the 75 ton spacecraft into orbit on its first flight is completed with the installation of Engine #2007 (top). At liftoff, each engine will be producing about 375,000 pounds of thrust, or about 12 million horsepower each, and gulping down its liquid oxygen and liquid hydrogen propellants at a rate of about 1,100 pounts per second. The Shuttle's main engines, the most efficient rocket engines ever built, are reusable and designed t operate over a life span of 55 missions.

Characterization and Analyses of Valves, Feed Lines and Tanks used in Propellant Delivery Systems at NASA SSC

NASA Technical Reports Server (NTRS)

Ryan, Harry M.; Coote, David J.; Ahuja, Vineet; Hosangadi, Ashvin

2006-01-01

Accurate modeling of liquid rocket engine test processes involves assessing critical fluid mechanic and heat and mass transfer mechanisms within a cryogenic environment, and accurately modeling fluid properties such as vapor pressure and liquid and gas densities as a function of pressure and temperature. The Engineering and Science Directorate at the NASA John C. Stennis Space Center has developed and implemented such analytic models and analysis processes that have been used over a broad range of thermodynamic systems and resulted in substantial improvements in rocket propulsion testing services. In this paper, we offer an overview of the analyses techniques used to simulate pressurization and propellant fluid systems associated with the test stands at the NASA John C. Stennis Space Center. More specifically, examples of the global performance (one-dimensional) of a propellant system are provided as predicted using the Rocket Propulsion Test Analysis (RPTA) model. Computational fluid dynamic (CFD) analyses utilizing multi-element, unstructured, moving grid capability of complex cryogenic feed ducts, transient valve operation, and pressurization and mixing in propellant tanks are provided as well.

Propellant Management and Conditioning within the X-34 Main Propulsion System

NASA Technical Reports Server (NTRS)

Brown, T. M.; McDonald, J. P.; Hedayat, A.; Knight, K. C.; Champion, R. H., Jr.

1998-01-01

The X-34 hypersonic flight vehicle is currently under development by Orbital Sciences Corporation (Orbital). The Main Propulsion ystem as been designed around the liquid propellant Fastrac rocket engine currently under development at NASA Marshall Space Flight Center. This paper presents analyses of the MPS subsystems used to manage the liquid propellants. These subsystems include the propellant tanks, the tank vent/relief subsystem, and the dump/fill/drain subsystem. Analyses include LOX tank chill and fill time estimates, LOX boil-off estimates, propellant conditioning simulations, and transient propellant dump simulations.

Advanced active health monitoring system of liquid rocket engines

NASA Astrophysics Data System (ADS)

Qing, Xinlin P.; Wu, Zhanjun; Beard, Shawn; Chang, Fu-Kuo

2008-11-01

An advanced SMART TAPE system has been developed for real-time in-situ monitoring and long term tracking of structural integrity of pressure vessels in liquid rocket engines. The practical implementation of the structural health monitoring (SHM) system including distributed sensor network, portable diagnostic hardware and dedicated data analysis software is addressed based on the harsh operating environment. Extensive tests were conducted on a simulated large booster LOX-H2 engine propellant duct to evaluate the survivability and functionality of the system under the operating conditions of typical liquid rocket engines such as cryogenic temperature, vibration loads. The test results demonstrated that the developed SHM system could survive the combined cryogenic temperature and vibration environments and effectively detect cracks as small as 2 mm.

Extended temperature range ACPS thruster investigation

NASA Technical Reports Server (NTRS)

Blubaugh, A. L.; Schoenman, L.

1974-01-01

The successful hot fire demonstration of a pulsing liquid hydrogen/liquid oxygen and gaseous hydrogen/liquid oxygen attitude control propulsion system thruster is described. The test was the result of research to develop a simple, lightweight, and high performance reaction control system without the traditional requirements for extensive periods of engine thermal conditioning, or the use of complex equipment to convert both liquid propellants to gas prior to delivery to the engine. Significant departures from conventional injector design practice were employed to achieve an operable design. The work discussed includes thermal and injector manifold priming analyses, subscale injector chilldown tests, and 168 full scale and 550 N (1250 lbF) rocket engine tests. Ignition experiments, at propellant temperatures ranging from cryogenic to ambient, led to the generation of a universal spark ignition system which can reliably ignite an engine when supplied with liquid, two phase, or gaseous propellants. Electrical power requirements for spark igniter are very low.

Coolant Design System for Liquid Propellant Aerospike Engines

NASA Astrophysics Data System (ADS)

McConnell, Miranda; Branam, Richard

2015-11-01

Liquid propellant rocket engines burn at incredibly high temperatures making it difficult to design an effective coolant system. These particular engines prove to be extremely useful by powering the rocket with a variable thrust that is ideal for space travel. When combined with aerospike engine nozzles, which provide maximum thrust efficiency, this class of rockets offers a promising future for rocketry. In order to troubleshoot the problems that high combustion chamber temperatures pose, this research took a computational approach to heat analysis. Chambers milled into the combustion chamber walls, lined by a copper cover, were tested for their efficiency in cooling the hot copper wall. Various aspect ratios and coolants were explored for the maximum wall temperature by developing our own MATLAB code. The code uses a nodal temperature analysis with conduction and convection equations and assumes no internal heat generation. This heat transfer research will show oxygen is a better coolant than water, and higher aspect ratios are less efficient at cooling. This project funded by NSF REU Grant 1358991.

Propulsion Estimates for High Energy Lunar Missions Using Future Propellants

NASA Technical Reports Server (NTRS)

Palaszewski, Bryan A.; Bennett, Gary L.

2016-01-01

High energy propellants for human lunar missions are analyzed, focusing on very advanced ozone and atomic hydrogen. One of the most advanced launch vehicle propulsion systems, such as the Space Shuttle Main Engine (SSME), used hydrogen and oxygen and had a delivered specific impulse of 453 seconds. In the early days of the space program, other propellants (or so called metapropellants) were suggested, including atomic hydrogen and liquid ozone. Theoretical and experimental studies of atomic hydrogen and ozone were conducted beginning in the late 1940s. This propellant research may have provided screenwriters with the idea of an atomic hydrogen-ozone rocket engine in the 1950 movie, Rocketship X-M. This paper presents analyses showing that an atomic hydrogen-ozone rocket engine could produce a specific impulse over a wide range of specific impulse values reaching as high as 1,600 s. A series of single stage and multistage rocket vehicle analyses were conducted to find the minimum specific impulse needed to conduct high energy round trip lunar missions.

Researcher Poses with a Nuclear Rocket Model

NASA Image and Video Library

1961-11-21

A researcher at the NASA Lewis Research Center with slide ruler poses with models of the earth and a nuclear-propelled rocket. The Nuclear Engine for Rocket Vehicle Applications (NERVA) was a joint NASA and Atomic Energy Commission (AEC) endeavor to develop a nuclear-powered rocket for both long-range missions to Mars and as a possible upper-stage for the Apollo Program. The early portion of the program consisted of basic reactor and fuel system research. This was followed by a series of Kiwi reactors built to test nuclear rocket principles in a non-flying nuclear engine. The next phase, NERVA, would create an entire flyable engine. The AEC was responsible for designing the nuclear reactor and overall engine. NASA Lewis was responsible for developing the liquid-hydrogen fuel system. The nuclear rocket model in this photograph includes a reactor at the far right with a hydrogen propellant tank and large radiator below. The payload or crew would be at the far left, distanced from the reactor.

High energy-density liquid rocket fuel performance

NASA Technical Reports Server (NTRS)

Rapp, Douglas C.

1990-01-01

A fuel performance database of liquid hydrocarbons and aluminum-hydrocarbon fuels was compiled using engine parametrics from the Space Transportation Engine Program as a baseline. Propellant performance parameters are introduced. General hydrocarbon fuel performance trends are discussed with respect to hydrogen-to-carbon ratio and heat of formation. Aluminum-hydrocarbon fuel performance is discussed with respect to aluminum metal loading. Hydrocarbon and aluminum-hydrocarbon fuel performance is presented with respect to fuel density, specific impulse, and propellant density specific impulse.

Method for Determining Optimum Injector Inlet Geometry

NASA Technical Reports Server (NTRS)

Myers, W. Neill (Inventor); Trinh, Huu P. (Inventor)

2015-01-01

A method for determining the optimum inlet geometry of a liquid rocket engine swirl injector includes obtaining a throttleable level phase value, volume flow rate, chamber pressure, liquid propellant density, inlet injector pressure, desired target spray angle and desired target optimum delta pressure value between an inlet and a chamber for a plurality of engine stages. The method calculates the tangential inlet area for each throttleable stage. The method also uses correlation between the tangential inlet areas and delta pressure values to calculate the spring displacement and variable inlet geometry of a liquid rocket engine swirl injector.

Thrust augmentation nozzle (TAN) concept for rocket engine booster applications

NASA Astrophysics Data System (ADS)

Forde, Scott; Bulman, Mel; Neill, Todd

2006-07-01

Aerojet used the patented thrust augmented nozzle (TAN) concept to validate a unique means of increasing sea-level thrust in a liquid rocket booster engine. We have used knowledge gained from hypersonic Scramjet research to inject propellants into the supersonic region of the rocket engine nozzle to significantly increase sea-level thrust without significantly impacting specific impulse. The TAN concept overcomes conventional engine limitations by injecting propellants and combusting in an annular region in the divergent section of the nozzle. This injection of propellants at moderate pressures allows for obtaining high thrust at takeoff without overexpansion thrust losses. The main chamber is operated at a constant pressure while maintaining a constant head rise and flow rate of the main propellant pumps. Recent hot-fire tests have validated the design approach and thrust augmentation ratios. Calculations of nozzle performance and wall pressures were made using computational fluid dynamics analyses with and without thrust augmentation flow, resulting in good agreement between calculated and measured quantities including augmentation thrust. This paper describes the TAN concept, the test setup, test results, and calculation results.

Video of SLS Liquid Hydrogen Tank Qualification Structural Test Article Being Moved to Cell E at NASA’s Michoud Assembly Facility

NASA Image and Video Library

2017-06-29

This video shows the Space Launch System liquid hydrogen tank structural qualification test article being moved to Building 110, Cell at NASA's Michoud Assembly Facility in New Orleans. The rocket's liquid hydrogen tank, which is the propellant tank that joins to the engine section of the 212-foot tall core stage, will carry cryogenic liquid hydrogen that propels the rocket. This test article build at Michoud is being prepared for testing at NASA's Marshall Space Flight Center in Huntsville, Alabama. There, it will be subjected to millions of pounds of force during testing to ensure the hardware can withstand the incredible stresses of launch.

Reduced Basis and Stochastic Modeling of Liquid Propellant Rocket Engine as a Complex System

DTIC Science & Technology

2015-07-02

additions, the approach will be extended to a real- gas system so that it can be used to investigate model multi-element liquid rocket combustors in a...Sirignano (2010). In the following discussion, we examine the various conservation principles for the gas and liquid phases. The hyperbolic nature of the...conservation equations for the gas and liquid phases. Mass conservation of individual chemical species or of individual classes of liquid droplets will

Radiation shielding estimates for manned Mars space flight.

PubMed

Dudkin, V E; Kovalev, E E; Kolomensky, A V; Sakovich, V A; Semenov, V F; Demin, V P; Benton, E V

1992-01-01

In the analysis of the required radiation shielding protection of spacecraft during a Mars flight, specific effects of solar activity (SA) on the intensity of galactic and solar cosmic rays were taken into consideration. Three spaceflight periods were considered: (1) maximum SA; (2) minimum SA; and (3) intermediate SA, when intensities of both galactic and solar cosmic rays are moderately high. Scenarios of spaceflights utilizing liquid-propellant rocket engines, low- and intermediate-thrust nuclear electrojet engines, and nuclear rocket engines, all of which have been designed in the Soviet Union, are reviewed. Calculations were performed on the basis of a set of standards for radiation protection approved by the U.S.S.R. State Committee for Standards. It was found that the lowest estimated mass of a Mars spacecraft, including the radiation shielding mass, obtained using a combination of a liquid propellant engine with low and intermediate thrust nuclear electrojet engines, would be 500-550 metric tons.

History of Sulphur Content Effects on the Thermal Stability of RP-1 under Heated Conditions

NASA Technical Reports Server (NTRS)

Irvine, Solveig A.; Schoettmer, Amanda K.; Bates, Ronald W.; Meyer, Michael L.

2004-01-01

As technologies advance in the aerospace industry, a strong desire has emerged to design more efficient, longer life, reusable liquid hydrocarbon fueled rocket engines. To achieve this goal, a more complete understanding of the thermal stability and chemical makeup of the hydrocarbon propellant is needed. Since the main fuel used in modern liquid hydrocarbon systems is RP-1, there is concern that Standard Grade RP-1 may not be a suitable propellant for future-generation rocket engines due to concern over the outdated Mil-Specification for the fuel. This current specification allows high valued limits on contaminants such as sulfur compounds, and also lacks specification of required thermal stability qualifications for the fuel. Previous studies have highlighted the detrimental effect of high levels of mercaptan sulfur content (^50 ppm) on copper rocket engine materials, but the fuel itself has not been studied. While the role of sulfur in other fuels (e.g., aviation, diesel, and automotive fuels) has been extensively studied, little has been reported on the effects of sulfur levels in rocket fuels. Lower RP-1 sulfur concentrations need to be evaluated and an acceptable sulfur limit established before RP-1 can be recommended for use as the propellant for future launch vehicles. (5 tables, 8 figures, 9 refs.)

Materials Problems in Chemical Liquid-Propellant Rocket Systems

NASA Technical Reports Server (NTRS)

Gilbert, L. L.

1959-01-01

With the advent of the space age, new adjustments in technical thinking and engineering experience are necessary. There is an increasing and extensive interest in the utilization of materials for components to be used at temperatures ranging from -423 to over 3500 deg F. This paper presents a description of the materials problems associated with the various components of chemical liquid rocket systems. These components include cooled and uncooled thrust chambers, injectors, turbine drive systems, propellant tanks, and cryogenic propellant containers. In addition to materials limitations associated with these components, suggested research approaches for improving materials properties are made. Materials such as high-temperature alloys, cermets, carbides, nonferrous alloys, plastics, refractory metals, and porous materials are considered.

High-Energy Propellant Rocket Firing at the Rocket Lab

NASA Image and Video Library

1955-01-21

A rocket using high-energy propellant is fired from the Rocket Laboratory at the National Advisory Committee for Aeronautics (NACA) Lewis Flight Propulsion Laboratory. The Rocket Lab was a collection of ten one-story cinderblock test cells located behind earthen barriers at the western edge of the campus. The rocket engines tested there were comparatively small, but the Lewis researchers were able to study different configurations, combustion performance, and injectors and nozzle design. The rockets were generally mounted horizontally and fired, as seen in this photograph of Test Cell No. 22. A group of fuels researchers at Lewis refocused their efforts after World War II in order to explore high energy propellants, combustion, and cooling. Research in these three areas began in 1945 and continued through the 1960s. The group of rocket researches was not elevated to a division branch until 1952. The early NACA Lewis work led to the development of liquid hydrogen as a viable propellant in the late 1950s. Following the 1949 reorganization of the research divisions, the rocket group began working with high-energy propellants such as diborane, pentaborane, and hydrogen. The lightweight fuels offered high levels of energy but were difficult to handle and required large tanks. In late 1954, Lewis researchers studied the combustion characteristics of gaseous hydrogen in a turbojet combustor. Despite poor mixing of the fuel and air, it was found that the hydrogen yielded more than a 90-percent efficiency. Liquid hydrogen became the focus of Lewis researchers for the next 15 years.

Review on pressure swirl injector in liquid rocket engine

NASA Astrophysics Data System (ADS)

Kang, Zhongtao; Wang, Zhen-guo; Li, Qinglian; Cheng, Peng

2018-04-01

The pressure swirl injector with tangential inlet ports is widely used in liquid rocket engine. Commonly, this type of pressure swirl injector consists of tangential inlet ports, a swirl chamber, a converging spin chamber, and a discharge orifice. The atomization of the liquid propellants includes the formation of liquid film, primary breakup and secondary atomization. And the back pressure and temperature in the combustion chamber could have great influence on the atomization of the injector. What's more, when the combustion instability occurs, the pressure oscillation could further affects the atomization process. This paper reviewed the primary atomization and the performance of the pressure swirl injector, which include the formation of the conical liquid film, the breakup and atomization characteristics of the conical liquid film, the effects of the rocket engine environment, and the response of the injector and atomization on the pressure oscillation.

Nonlinear Modeling and Control of a Propellant Mixer

NASA Technical Reports Server (NTRS)

Barbieri, Enrique; Richter, Hanz; Figueroa, Fernando

2003-01-01

A mixing chamber used in rocket engine combustion testing at NASA Stennis Space Center is modeled by a second order nonlinear MIMO system. The mixer is used to condition the thermodynamic properties of cryogenic liquid propellant by controlled injection of the same substance in the gaseous phase. The three inputs of the mixer are the positions of the valves regulating the liquid and gas flows at the inlets, and the position of the exit valve regulating the flow of conditioned propellant. The outputs to be tracked and/or regulated are mixer internal pressure, exit mass flow, and exit temperature. The outputs must conform to test specifications dictated by the type of rocket engine or component being tested downstream of the mixer. Feedback linearization is used to achieve tracking and regulation of the outputs. It is shown that the system is minimum-phase provided certain conditions on the parameters are satisfied. The conditions are shown to have physical interpretation.

Scaling of Performance in Liquid Propellant Rocket Engine Combustors

NASA Technical Reports Server (NTRS)

Hulka, James

2008-01-01

The objectives are: a) Re-introduce to you the concept of scaling; b) Describe the scaling research conducted in the 1950s and early 1960s, and present some of their conclusions; c) Narrow the focus to scaling for performance of combustion devices for liquid propellant rocket engines; and d) Present some results of subscale to full-scale performance from historical programs. Scaling is "The ability to develop new combustion devices with predictable performance on the basis of test experience with old devices." Scaling can be used to develop combustion devices of any thrust size from any thrust size. Scaling is applied mostly to increase thrust. Objective is to use scaling as a development tool. - Move injector design from an "art" to a "science"

NASA Engineer Examines the Design of a Regeneratively-Cooled Rocket Engine

NASA Image and Video Library

1958-12-21

An engineer at the National Aeronautics and Space Administration (NASA) Lewis Research Center examines a drawing showing the assembly and details of a 20,000-pound thrust regeneratively cooled rocket engine. The engine was being designed for testing in Lewis’ new Rocket Engine Test Facility, which began operating in the fall of 1957. The facility was the largest high-energy test facility in the country that was capable of handling liquid hydrogen and other liquid chemical fuels. The facility’s use of subscale engines up to 20,000 pounds of thrust permitted a cost-effective method of testing engines under various conditions. The Rocket Engine Test Facility was critical to the development of the technology that led to the use of hydrogen as a rocket fuel and the development of lightweight, regeneratively-cooled, hydrogen-fueled rocket engines. Regeneratively-cooled engines use the cryogenic liquid hydrogen as both the propellant and the coolant to prevent the engine from burning up. The fuel was fed through rows of narrow tubes that surrounded the combustion chamber and nozzle before being ignited inside the combustion chamber. The tubes are visible in the liner sitting on the desk. At the time, Pratt and Whitney was designing a 20,000-pound thrust liquid-hydrogen rocket engine, the RL-10. Two RL-10s would be used to power the Centaur second-stage rocket in the 1960s. The successful development of the Centaur rocket and the upper stages of the Saturn V were largely credited to the work carried out Lewis.

Evaluation of Impinging Stream Vortex Chamber Concepts for Liquid Rocket Engine Applications

NASA Technical Reports Server (NTRS)

Trinh, Huu; Kopicz, Charles; Bullard, Brad; Michaels, Scott

2003-01-01

NASA Marshall Space Flight Center (MSFC) and the U. S. Army are jointly investigating vortex chamber concepts for cryogenic oxygen/hydrocarbon fuel rocket engine applications. One concept, the Impinging Stream Vortex Chamber Concept (ISVC), has been tested with gel propellants at AMCOM at Redstone Arsenal, Alabama. A version of this concept for the liquid oxygen (LOX)/hydrocarbon fuel (RP-1) propellant system is derived from the one for the gel propellant. An unlike impinging injector is employed to deliver the propellants to the chamber. MSFC has also designed two alternative injection schemes, called the chasing injectors, associated with this vortex chamber concept. In these injection techniques, both propellant jets and their impingement point are in the same chamber cross-sectional plane. One injector has a similar orifice size with the original unlike impinging injector. The second chasing injector has small injection orifices. The team has achieved their objectives of demonstrating the self-cooled chamber wall benefits of ISVC and of providing the test data for validating computational fluids dynamics (CFD) models. These models, in turn, will be used to design the optimum vortex chambers in the future.

Low-thrust chemical rocket engine study

NASA Technical Reports Server (NTRS)

Mellish, J. A.

1981-01-01

Engine data and information are presented to perform system studies on cargo orbit-transfer vehicles which would deliver large space structures to geosynchronous equatorial orbit. Low-thrust engine performance, weight, and envelope parametric data were established, preliminary design information was generated, and technologies for liquid rocket engines were identified. Two major engine design drivers were considered in the study: cooling and engine cycle options. Both film-cooled and regeneratively cooled engines were evaluated. The propellant combinations studied were hydrogen/oxygen, methane/oxygen, and kerosene/oxygen.

Potential low cost, safe, high efficiency propellant for future space program

NASA Astrophysics Data System (ADS)

Zhou, D.

2005-03-01

Mixtures of nanometer or micrometer sized carbon powder suspended in hydrogen and methane/hydrogen mixtures are proposed as candidates for low cost, high efficiency propellants for future space programs. While liquid hydrogen has low weight and high heat of combustion per unit mass, because of the low mass density the heat of combustion per unit volume is low, and the liquid hydrogen storage container must be large. The proposed propellants can produce higher gross heat combustion with small volume with trade off of some weight increase. Liquid hydrogen can serve as the fluid component of the propellant in the mixtures and thus used by current rocket engine designs. For example, for the same volume a mixture of 5% methane and 95% hydrogen, can lead to an increase in the gross heat of combustion by about 10% and an increase in the Isp (specific impulse) by 21% compared to a pure liquid hydrogen propellant. At liquid hydrogen temperatures of 20.3 K, methane will be in solid state, and must be formed as fine granules (or slush) to satisfy the requirement of liquid propellant engines.

Liquid-Propellant Rocket Engine Throttling: A Comprehensive Review

NASA Technical Reports Server (NTRS)

Casiano, Matthew; Hulka, James; Yang, Virog

2009-01-01

Liquid-Propellant Rocket Engines (LREs) are capable of on-command variable thrust or thrust modulation, an operability advantage that has been studied intermittently since the late 1930s. Throttleable LREs can be used for planetary entry and descent, space rendezvous, orbital maneuvering including orientation and stabilization in space, and hovering and hazard avoidance during planetary landing. Other applications have included control of aircraft rocket engines, limiting of vehicle acceleration or velocity using retrograde rockets, and ballistic missile defense trajectory control. Throttleable LREs can also continuously follow the most economical thrust curve in a given situation, compared to discrete throttling changes over a few select operating points. The effects of variable thrust on the mechanics and dynamics of an LRE as well as difficulties and issues surrounding the throttling process are important aspects of throttling behavior. This review provides a detailed survey of LRE throttling centered around engines from the United States. Several LRE throttling methods are discussed, including high-pressure-drop systems, dual-injector manifolds, gas injection, multiple chambers, pulse modulation, throat throttling, movable injector components, and hydrodynamically dissipative injectors. Concerns and issues surrounding each method are examined, and the advantages and shortcomings compared.

Propellant Vaporization as a Criterion for Rocket-Engine Design; Experimental Performance, Vaporization and Heat-Transfer Rates with Various Propellant Combinations

NASA Technical Reports Server (NTRS)

Clark, Bruce J.; Hersch, Martin; Priem, Richard J.

1959-01-01

Experimental combustion efficiencies of eleven propellant combinations were determined as a function of chamber length. Efficiencies were measured in terms of characteristic exhaust velocities at three chamber lengths and in terms of gas velocities. The data were obtained in a nominal 200-pound-thrust rocket engine. Injector and engine configurations were kept essentially the same to allow comparison of the performance. The data, except for those on hydrazine and ammonia-fluorine, agreed with predicted results based on the assumption that vaporization of the propellants determines the rate of combustion. Decomposition in the liquid phase may be.responsible for the anomalous behavior of hydrazine. Over-all heat-transfer rates were also measured for each combination. These rates were close to the values predicted by standard heat-transfer calculations except for the combinations using ammonia.

LOX droplet vaporization in a supercritical forced convective environment

NASA Technical Reports Server (NTRS)

Hsiao, Chia-Chun; Yang, Vigor

1993-01-01

Modern liquid rocket engines often use liquid oxygen (LOX) and liquid hydrogen (LH2) as propellants to achieve high performance, with the engine operational conditions in the supercritical regimes of the propellants. Once the propellant exceeds its critical state, it essentially becomes a puff of dense fluid. The entire field becomes a continuous medium, and no distinct interfacial boundary between the liquid and gas exists. Although several studies have been undertaken to investigate the supercritical droplet behavior at quiescent conditions, very little effort has been made to address the fundamental mechanisms associated with LOX droplet vaporization in a supercritical, forced convective environment. The purpose is to establish a theoretical framework within which supercritical droplet dynamics and vaporization can be studied systematically by means of an efficient and robust numerical algorithm.

Experimental Altitude Performance of JP-4 Fuel and Liquid-Oxygen Rocket Engine with an Area Ratio of 48

NASA Technical Reports Server (NTRS)

Fortini, Anthony; Hendrix, Charles D.; Huff, Vearl N.

1959-01-01

The performance for four altitudes (sea-level, 51,000, 65,000, and 70,000 ft) of a rocket engine having a nozzle area ratio of 48.39 and using JP-4 fuel and liquid oxygen as a propellant was evaluated experimentally by use of a 1000-pound-thrust engine operating at a chamber pressure of 600 pounds per square inch absolute. The altitude environment was obtained by a rocket-ejector system which utilized the rocket exhaust gases as the pumping fluid of the ejector. Also, an engine having a nozzle area ratio of 5.49 designed for sea level was tested at sea-level conditions. The following table lists values from faired experimental curves at an oxidant-fuel ratio of 2.3 for various approximate altitudes.

Primary atomization of liquid jets issuing from rocket engine coaxial injectors

NASA Astrophysics Data System (ADS)

Woodward, Roger D.

1993-01-01

The investigation of liquid jet breakup and spray development is critical to the understanding of combustion phenomena in liquid-propellant rocket engines. Much work has been done to characterize low-speed liquid jet breakup and dilute sprays, but atomizing jets and dense sprays have yielded few quantitative measurements due to their optical opacity. This work focuses on a characteristic of the primary breakup process of round liquid jets, namely the length of the intact liquid core. The specific application considered is that of shear-coaxial type rocket engine injectors. Real-time x-ray radiography, capable of imaging through the dense two-phase region surrounding the liquid core, has been used to make the measurements. Nitrogen and helium were employed as the fuel simulants while an x-ray absorbing potassium iodide aqueous solution was used as the liquid oxygen (LOX) simulant. The intact-liquid-core length data have been obtained and interpreted to illustrate the effects of chamber pressure (gas density), injected-gas and liquid velocities, and cavitation. The results clearly show that the effect of cavitation must be considered at low chamber pressures since it can be the dominant breakup mechanism. A correlation of intact core length in terms of gas-to-liquid density ratio, liquid jet Reynolds number, and Weber number is suggested. The gas-to-liquid density ratio appears to be the key parameter for aerodynamic shear breakup in this study. A small number of hot-fire, LOX/hydrogen tests were also conducted to attempt intact-LOX-core measurements under realistic conditions in a single-coaxial-element rocket engine. The tests were not successful in terms of measuring the intact core, but instantaneous imaging of LOX jets suggests that LOX jet breakup is qualitatively similar to that of cold-flow, propellant-simulant jets. The liquid oxygen jets survived in the hot-fire environment much longer than expected, and LOX was even visualized exiting the chamber nozzle under some conditions. This may be an effect of the single element configuration.

Atmospheric Mining in the Outer Solar System: Outer Planet Orbital Transfer and Lander Analyses

NASA Technical Reports Server (NTRS)

Palaszewski, Bryan

2016-01-01

High energy propellants for human lunar missions are analyzed, focusing on very advanced ozone and atomic hydrogen. One of the most advanced launch vehicle propulsion systems, such as the Space Shuttle Main Engine (SSME), used hydrogen and oxygen and had a delivered specific impulse of 453 seconds. In the early days of the space program, other propellants (or so called metapropellants) were suggested, including atomic hydrogen and liquid ozone. Theoretical and experimental studies of atomic hydrogen and ozone were conducted beginning in the late 1940s. This propellant research may have provided screenwriters with the idea of an atomic hydrogen-ozone rocket engine in the 1950 movie, Rocketship X-M. This paper presents analyses showing that an atomic hydrogen-ozone rocket engine could produce a specific impulse over a wide range of specific impulse values reaching as high as 1,600 seconds. A series of single stage and multistage rocket vehicle analyses were conducted to find the minimum specific impulse needed to conduct high energy round trip lunar missions.

Raman Gas Species Measurements in Hydrocarbon-Fueled Rocket Engine Injector Flows

NASA Technical Reports Server (NTRS)

Wehrmeyer, Joseph; Hartfield, Roy J., Jr.; Trinh, Huu P.; Dobson, Chris C.; Eskridge, Richard H.

2000-01-01

Rocket engine propellent injector development at NASA-Marshall includes experimental analysis using optical techniques, such as Raman, fluorescence, or Mie scattering. For the application of spontaneous Raman scattering to hydrocarbon-fueled flows a technique needs to be developed to remove the interfering polycyclic aromatic hydrocarbon fluorescence from the relatively weak Raman signals. A current application of such a technique is to the analysis of the mixing and combustion performance of multijet, impinging-jet candidate fuel injectors for the baseline Mars ascent engine, which will burn methane and liquid oxygen produced in-situ on Mars to reduce the propellent mass transported to Mars for future manned Mars missions. The Raman technique takes advantage of the strongly polarized nature of Raman scattering. It is shown to be discernable from unpolarized fluorescence interference by subtracting one polarized image from another. Both of these polarized images are obtained from a single laser pulse by using a polarization-separating calcite rhomb mounted in the imaging spectrograph. A demonstration in a propane-air flame is presented, as well as a high pressure demonstration in the NASA-Marshall Modular Combustion Test Artice, using the liquid methane-liquid oxygen propellant system

Hot Fire Ignition Test with Densified Liquid Hydrogen using a RL10B-2 Cryogenic H2/O2 Rocket Engine

NASA Technical Reports Server (NTRS)

McNelis, Nancy B.; Haberbusch, Mark S.

1997-01-01

Enhancements to propellants provide an opportunity to either increase performance of an existing vehicle, or reduce the size of a new vehicle. In the late 1980's the National AeroSpace Plane (NASP) reopened the technology chapter on densified propellants, in particular hydrogen. Since that point in time the NASA Lewis Research Center (LERC) in Cleveland, Ohio has been leading the way to provide critical research on the production and transfer of densified propellants. On October 4, 1996 NASA LeRC provided another key demonstration towards the advancement of densified propellants as a viable fuel. Successful ignition of an RL10B-2 engine was achieved with near triple point liquid hydrogen.

Analysis of rocket engine injection combustion processes

NASA Technical Reports Server (NTRS)

Salmon, J. W.

1976-01-01

A critique is given of the JANNAF sub-critical propellant injection/combustion process analysis computer models and application of the models to correlation of well documented hot fire engine data bases. These programs are the distributed energy release (DER) model for conventional liquid propellants injectors and the coaxial injection combustion model (CICM) for gaseous annulus/liquid core coaxial injectors. The critique identifies model inconsistencies while the computer analyses provide quantitative data on predictive accuracy. The program is comprised of three tasks: (1) computer program review and operations; (2) analysis and data correlations; and (3) documentation.

Rocket Propellant Talk at the 1957 NACA Lewis Inspection

NASA Image and Video Library

1957-10-21

A researcher works a demonstration board in the Rocket Engine Test Facility during the 1957 Inspection of the National Advisory Committee for Aeronautics (NACA) Lewis Flight Propulsion Laboratory in Cleveland, Ohio. Representatives from the military, aeronautical industry, universities, and the press were invited to the laboratory to be briefed on the NACA’s latest research efforts and tour the test facilities. Over 1700 people visited the Lewis during the October 7-10, 1957 Inspection. The Soviet Union launched their first Sputnik satellite just days before on October 4. NACA Lewis had been involved in small rockets and propellants research since 1945, but the NACA leadership was wary of involving itself too deeply with the work since ballistics traditionally fell under the military’s purview. The Lewis research was performed by the High Temperature Combustion section in the Fuels and Lubricants Division in a series of small cinderblock test cells. The rocket group was expanded in 1952 and made several test runs in late 1954 using liquid hydrogen as a propellant. A larger test facility, the Rocket Engine Test Facility, was approved and became operational just in time for the Inspection.

Amplification of Reynolds number dependent processes by wave distortion. [acoustic instability of liquid propellant rocket engines

NASA Technical Reports Server (NTRS)

Ventrice, M. B.; Fang, J. C.; Purdy, K. R.

1975-01-01

A system using a hot-wire transducer as an analog of a liquid droplet of propellant was employed to investigate the ingredients of the acoustic instability of liquid-propellant rocket engines. It was assumed that the combustion process was vaporization-limited and that the combustion chamber was acoustically similar to a closed-closed right-circular cylinder. Before studying the hot-wire closed-loop system (the analog system), a microphone closed-loop system, which used the response of a microphone as the source of a linear feedback exciting signal, was investigated to establish the characteristics of self-sustenance of acoustic fields. Self-sustained acoustic fields were found to occur only at resonant frequencies of the chamber. In the hot-wire closed-loop system, the response of hot-wire anemometer was used as the source of the feedback exciting signal. The self-sustained acoustic fields which developed in the system were always found to be harmonically distorted and to have as their fundamental frquency a resonant frequency for which there also existed a second resonant frequency which was approximately twice the fundamental frequency.

Combustion dynamics in liquid rocket engines

NASA Technical Reports Server (NTRS)

Mclain, W. H.

1971-01-01

A chemical analysis of the emission and absorption spectra in the combustion chamber of a nitrogen tetroxide/aerozine-50 rocket engine was conducted. Measurements were made under conditions of preignition, ignition, and post combustion operating periods. The cause of severe ignition overpressures sporadically observed during the vacuum startup of the Apollo reaction control system engine was investigated. The extent to which residual propellants or condensed intermediate reaction products remain after the engine has been operated in a pulse mode duty cycle was determined.

Guided Rocket Weapon,

DTIC Science & Technology

1982-06-11

nyn;tei% (fuel/propellant ir, extruded trent the tanks by *C Cc:’ njie d g a. Work liquid-propellant engines on the same principle, as on the ~c~A t...82052705 PAGE 44-- Fig. 26. Starting/launcing of the guided winged missile * Snack ". Page 49.1 Ballistic short-range missiles. The most widely used short

Schlieren image velocimetry measurements in a rocket engine exhaust plume

NASA Astrophysics Data System (ADS)

Morales, Rudy; Peguero, Julio; Hargather, Michael

2017-11-01

Schlieren image velocimetry (SIV) measures velocity fields by tracking the motion of naturally-occurring turbulent flow features in a compressible flow. Here the technique is applied to measuring the exhaust velocity profile of a liquid rocket engine. The SIV measurements presented include discussion of visibility of structures, image pre-processing for structure visibility, and ability to process resulting images using commercial particle image velocimetry (PIV) codes. The small-scale liquid bipropellant rocket engine operates on nitrous oxide and ethanol as propellants. Predictions of the exhaust velocity are obtained through NASA CEA calculations and simple compressible flow relationships, which are compared against the measured SIV profiles. Analysis of shear layer turbulence along the exhaust plume edge is also presented.

Liquid rocket propulsion: Retrospective and prospects

NASA Astrophysics Data System (ADS)

Rosenberg, Sanders D.

1993-02-01

Rocket propulsion has made a fundamental contribution to change in the human condition during the second half of the 20th Century. This paper presents a survey of the basic elements of and future prospects for liquid rocket propulsion systems, with emphasis placed on their bipropellant engines, which have contributed profoundly to the successes of this 'aerospace century.' Many technologies had to reach maturity simultaneously to enable our current progress: materials, electronics, guidance and control, systems engineering, and propulsion, made major contributions. However, chemical propellants and the engine systems required to extract and control their propulsive power successfully are at the heart of all that humankind has accomplished through space flight and the use of space for the betterment of all. And it is a fascinating story to tell.

Liquid Propulsion: Propellant Feed System Design. Chapter 2.3.11

NASA Technical Reports Server (NTRS)

Cannon, James L.

2010-01-01

The propellant feed system of a liquid rocket engine determines how the propellants are delivered from the tanks to the thrust chamber. They are generally classified as either pressure fed or pump fed. The pressure-fed system is simple and relies on tank pressures to feed the propellants into the thrust chamber. This type of system is typically used for space propulsion applications and auxiliary propulsion applications requiring low system pressures and small quantities of propellants. In contrast, the pump-fed system is used for high pressure, high performance applications. The selection of one propellant feed system over another is determined based on design trade studies at both the engine and vehicle levels. This chapter first provides a brief overview of the basic configurations of pressure-fed systems. Pump-fed systems are then discussed with greater detail given to the turbomachinery design. Selected design requirements and configurations are provided.

Liquid Rocket Booster Study. Volume 2, Book 1

NASA Technical Reports Server (NTRS)

1989-01-01

The recommended Liquid Rocket Booster (LRB) concept is shown which uses a common main engine with the Advanced Launch System (ALS) which burns LO2 and LH2. The central rationale is based on the belief that the U.S. can only afford one big new rocket engine development in the 1990's. A LO2/LH2 engine in the half million pound thrust class could satisfy STS LRB, ALS, and Shuttle C (instead of SSMEs). Development costs and higher production rates can be shared by NASA and USAF. If the ALS program does not occur, the LO2/RP-1 propellants would produce slight lower costs for and STS LRB. When the planned Booster Engine portion of the Civil Space Transportation Initiatives has provided data on large pressure fed LO2/RP-1 engines, then the choice should be reevaluated.

Development of LM10-MIRA LOX/LNG expander cycle demonstrator engine

NASA Astrophysics Data System (ADS)

Rudnykh, Mikhail; Carapellese, Stefano; Liuzzi, Daniele; Arione, Luigi; Caggiano, Giuseppe; Bellomi, Paolo; D'Aversa, Emanuela; Pellegrini, Rocco; Lobov, S. D.; Gurtovoy, A. A.; Rachuk, V. S.

2016-09-01

This article contains results of joint works by Konstruktorskoe Buro Khimavtomatiki (KBKhA, Russia) and AVIO Company (Italy) on creation of the LM10-MIRA liquid-propellant rocket demonstrator engine for the third stage of the upgraded "Vega" launcher.Scientific and research activities conducted by KBKhA and AVIO in 2007-2014 in the frame of the LYRA Program, funded by the Italian Space Agency, with ELV as Prime contractor, and under dedicated ASI-Roscosmos inter-agencies agreement, were aimed at development and testing of a 7.5 t thrust expander cycle demonstrator engine propelled by oxygen and liquid natural gas (further referred to as LNG).

RD860 and RD860L Engines with Deep Thrust Throttling and a High Technology Readiness Level (TRL)

NASA Astrophysics Data System (ADS)

Prokopchuk, O. O.; Shul'ga, V. A.; Dibrivnyi, O. V.; Kukhta, A. S.

2018-04-01

To solve the problems of delivering payloads to Mars surface and returning them to the orbit, liquid rocket engines, operating on storable propellants with deep throttling possibility, are needed, besides having high energy-mass characteristics.

Reusable Rocket Engine Advanced Health Management System. Architecture and Technology Evaluation: Summary

NASA Technical Reports Server (NTRS)

Pettit, C. D.; Barkhoudarian, S.; Daumann, A. G., Jr.; Provan, G. M.; ElFattah, Y. M.; Glover, D. E.

1999-01-01

In this study, we proposed an Advanced Health Management System (AHMS) functional architecture and conducted a technology assessment for liquid propellant rocket engine lifecycle health management. The purpose of the AHMS is to improve reusable rocket engine safety and to reduce between-flight maintenance. During the study, past and current reusable rocket engine health management-related projects were reviewed, data structures and health management processes of current rocket engine programs were assessed, and in-depth interviews with rocket engine lifecycle and system experts were conducted. A generic AHMS functional architecture, with primary focus on real-time health monitoring, was developed. Fourteen categories of technology tasks and development needs for implementation of the AHMS were identified, based on the functional architecture and our assessment of current rocket engine programs. Five key technology areas were recommended for immediate development, which (1) would provide immediate benefits to current engine programs, and (2) could be implemented with minimal impact on the current Space Shuttle Main Engine (SSME) and Reusable Launch Vehicle (RLV) engine controllers.

A-3 Test Stand construction

NASA Image and Video Library

2010-10-01

An 80,000-gallon liquid hydrogen tank is placed at the A-3 Test Stand construction site on Sept. 24, 2010. The tank will provide propellant for tests of next-generation rocket engines at the stand. It will be placed upright on top of the stand, helping to increase the overall height to 300 feet. Once completed, the A-3 Test Stand will enable operators to test rocket engines at simulated altitudes of up to 100,000 feet. The A-3 stand is the first large rocket engine test structure to be built at Stennis Space Center since the 1960s.

A-3 Test Stand construction

NASA Image and Video Library

2010-09-24

A 35,000-gallon liquid oxygen tank is placed at the A-3 Test Stand construction site on Sept. 24, 2010. The tank will provide propellant for tests of next-generation rocket engines at the stand. It will be placed upright on top of the stand, helping to increase the overall height to 300 feet. Once completed, the A-3 Test Stand will enable operators to test rocket engines at simulated altitudes of up to 100,000 feet. The A-3 stand is the first large rocket engine test structure to be built at Stennis Space Center since the 1960s.

Boiler and Pressure Balls Monopropellant Thermal Rocket Engine

NASA Technical Reports Server (NTRS)

Greene, William D. (Inventor)

2009-01-01

The proposed technology is a rocket engine cycle utilizing as the propulsive fluid a low molecular weight, cryogenic fluid, typically liquid hydrogen, pressure driven, heated, and expelled through a nozzle to generate high velocity and high specific impulse discharge gas. The proposed technology feeds the propellant through the engine cycle without the use of a separate pressurization fluid and without the use of turbomachinery. Advantages of the proposed technology are found in those elements of state-of-the-art systems that it avoids. It does not require a separate pressurization fluid or a thick-walled primary propellant tank as is typically required for a classical pressure-fed system. Further, it does not require the acceptance of intrinsic reliability risks associated with the use of turbomachinery

Russian Rocket Engine Test

NASA Technical Reports Server (NTRS)

1998-01-01

NASA engineers successfully tested a Russian-built rocket engine on November 4, 1998 at the Marshall Space Flight Center (MSFC) Advanced Engine Test Facility, which had been used for testing the Saturn V F-1 engines and Space Shuttle Main engines. The MSFC was under a Space Act Agreement with Lockheed Martin Astronautics of Denver to provide a series of test firings of the Atlas III propulsion system configured with the Russian-designed RD-180 engine. The tests were designed to measure the performance of the Atlas III propulsion system, which included avionics and propellant tanks and lines, and how these components interacted with the RD-180 engine. The RD-180 is powered by kerosene and liquid oxygen, the same fuel mix used in Saturn rockets. The RD-180, the most powerful rocket engine tested at the MSFC since Saturn rocket tests in the 1960s, generated 860,000 pounds of thrust.

Small rocket research and technology

NASA Technical Reports Server (NTRS)

Schneider, Steven; Biaglow, James

1993-01-01

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.

Fiberoptic sensors for rocket engine applications

NASA Technical Reports Server (NTRS)

Ballard, R. O.

1992-01-01

A research effort was completed to summarize and evaluate the current level of technology in fiberoptic sensors for possible applications in integrated control and health monitoring (ICHM) systems in liquid propellant engines. The environment within a rocket engine is particuarly severe with very high temperatures and pressures present combined with extremely rapid fluid and gas flows, and high-velocity and high-intensity acoustc waves. Application of fiberoptic technology to rocket engine health monitoring is a logical evolutionary step in ICHM development and presents a significant challenge. In this extremely harsh environment, the additional flexibility of fiberoptic techniques to augment conventional sensor technologies offer abundant future potential.

A-1 Test Stand modifications

NASA Image and Video Library

2011-09-14

Team members check the progress of a liquid nitrogen cold shock test on the A-1 Test Stand at Stennis Space Center on Sept. 15. The cold shock test is used to confirm the test stand's support system can withstand test conditions, when super-cold rocket engine propellant is piped. The A-1 Test Stand is preparing to conduct tests on the powerpack component of the J-2X rocket engine, beginning in early 2012.

LH tank installation

NASA Image and Video Library

2011-07-25

Stennis Space Center employees marked another construction milestone July 25 with installation of the 85,000-gallon liquid hydrogen tank atop the A-3 Test Stand. The 300-foot-tall stand is being built to test next-generation rocket engines that could carry humans into deep space once more. The liquid hydrogen tank and a 35,000-gallon liquid oxygen tank installed atop the steel structure earlier in June will provide fuel propellants for testing the engines.

Injector Design Tool Improvements: User's manual for FDNS V.4.5

NASA Technical Reports Server (NTRS)

Chen, Yen-Sen; Shang, Huan-Min; Wei, Hong; Liu, Jiwen

1998-01-01

The major emphasis of the current effort is in the development and validation of an efficient parallel machine computational model, based on the FDNS code, to analyze the fluid dynamics of a wide variety of liquid jet configurations for general liquid rocket engine injection system applications. This model includes physical models for droplet atomization, breakup/coalescence, evaporation, turbulence mixing and gas-phase combustion. Benchmark validation cases for liquid rocket engine chamber combustion conditions will be performed for model validation purpose. Test cases may include shear coaxial, swirl coaxial and impinging injection systems with combinations LOXIH2 or LOXISP-1 propellant injector elements used in rocket engine designs. As a final goal of this project, a well tested parallel CFD performance methodology together with a user's operation description in a final technical report will be reported at the end of the proposed research effort.

Radiation effect on rocket engine performance

NASA Technical Reports Server (NTRS)

Chiu, Huei-Huang

1988-01-01

The effects of radiation on the performance of modern rocket propulsion systems operating at high pressure and temperature were recognized as a key issue in the design and operation of various liquid rocket engines of the current and future generations. Critical problem areas of radiation coupled with combustion of bipropellants are assessed and accounted for in the formulation of a universal scaling law incorporated with a radiation-enhanced vaporization combustion model. Numerical algorithms are developed and the pertaining data of the Variable Thrust Engine (VTE) and Space Shuttle Main Engine (SSME) are used to conduct parametric sensitivity studies to predict the principal intercoupling effects of radiation. The analysis reveals that low enthalpy engines, such as the VTE, are vulnerable to a substantial performance set back by the radiative loss, whereas the performance of high enthalpy engines such as the SSME, are hardly affected over a broad range of engine operation. Additionally, combustion enhancement by the radiative heating of the propellant has a significant impact in those propellants with high absorptivity. Finally, the areas of research related with radiation phenomena in bipropellant engines are identified.

Plasma Igniter for Reliable Ignition of Combustion in Rocket Engines

NASA Technical Reports Server (NTRS)

Martin, Adam; Eskridge, Richard

2011-01-01

A plasma igniter has been developed for initiating combustion in liquid-propellant rocket engines. The device propels a hot, dense plasma jet, consisting of elemental fluorine and fluorine compounds, into the combustion chamber to ignite the cold propellant mixture. The igniter consists of two coaxial, cylindrical electrodes with a cylindrical bar of solid Teflon plastic in the region between them. The outer electrode is a metal (stainless steel) tube; the inner electrode is a metal pin (mild steel, stainless steel, tungsten, or thoriated-tungsten). The Teflon bar fits snugly between the two electrodes and provides electrical insulation between them. The Teflon bar may have either a flat surface, or a concave, conical surface at the open, down-stream end of the igniter (the igniter face). The igniter would be mounted on the combustion chamber of the rocket engine, either on the injector-plate at the upstream side of the engine, or on the sidewalls of the chamber. It also might sit behind a valve that would be opened just prior to ignition, and closed just after, in order to prevent the Teflon from melting due to heating from the combustion chamber.

Evaluation of Vortex Chamber Concepts for Liquid Rocket Engine Applications

NASA Technical Reports Server (NTRS)

Trinh, Huu Phuoc; Knuth, Williams; Michaels, Scott; Turner, James E. (Technical Monitor)

2000-01-01

Rocket-based combined-cycle engines (RBBC) being considered at NASA for future generation launch vehicles feature clusters of small rocket thrusters as part of the engine components. Depending on specific RBBC concepts, these thrusters may be operated at various operating conditions including power level and/or propellant mixture ratio variations. To pursue technology developments for future launch vehicles, NASA/Marshall Space Flight Center (MSFC) is examining vortex chamber concepts for the subject cycle engine application. Past studies indicated that the vortex chamber schemes potentially have a number of advantages over conventional chamber methods. Due to the nature of the vortex flow, relatively cooler propellant streams tend to flow along the chamber wall. Hence, the thruster chamber can be operated without the need of any cooling techniques. This vortex flow also creates strong turbulence, which promotes the propellant mixing process. Consequently, the subject chamber concepts not only offer the system simplicity but they also would enhance the combustion performance. The test results showed that the chamber performance was markedly high even at a low chamber length-to- diameter ratio (L/D). This incentive can be translated to a convenience in the thrust chamber packaging.

Evaluation of Proposed Rocket Engines for Earth-to-Orbit Vehicles

NASA Technical Reports Server (NTRS)

Martin, James A.; Kramer, Richard D.

1990-01-01

The objective is to evaluate recently analyzed rocket engines for advanced Earth-to-orbit vehicles. The engines evaluated are full-flow staged combustion engines and split expander engines, both at mixture ratios at 6 and above with oxygen and hydrogen propellants. The vehicles considered are single-stage and two-stage fully reusable vehicles and the Space Shuttle with liquid rocket boosters. The results indicate that the split expander engine at a mixture ratio of about 7 is competitive with the full-flow staged combustion engine for all three vehicle concepts. A key factor in this result is the capability to increase the chamber pressure for the split expander as the mixture ratio is increased from 6 to 7.

Nonlinear Combustion Instability Prediction

NASA Technical Reports Server (NTRS)

Flandro, Gary

2010-01-01

The liquid rocket engine stability prediction software (LCI) predicts combustion stability of systems using LOX-LH2 propellants. Both longitudinal and transverse mode stability characteristics are calculated. This software has the unique feature of being able to predict system limit amplitude.

A simplified computer program for the prediction of the linear stability behavior of liquid propellant combustors

NASA Technical Reports Server (NTRS)

Mitchell, C. E.; Eckert, K.

1979-01-01

A program for predicting the linear stability of liquid propellant rocket engines is presented. The underlying model assumptions and analytical steps necessary for understanding the program and its input and output are also given. The rocket engine is modeled as a right circular cylinder with an injector with a concentrated combustion zone, a nozzle, finite mean flow, and an acoustic admittance, or the sensitive time lag theory. The resulting partial differential equations are combined into two governing integral equations by the use of the Green's function method. These equations are solved using a successive approximation technique for the small amplitude (linear) case. The computational method used as well as the various user options available are discussed. Finally, a flow diagram, sample input and output for a typical application and a complete program listing for program MODULE are presented.

The Technique for CFD-Simulation of Fuel Valve from Pneumatic-Hydraulic System of Liquid-Propellant Rocket Engine

NASA Astrophysics Data System (ADS)

Shabliy, L. S.; Malov, D. V.; Bratchinin, D. S.

2018-01-01

In the article the description of technique for simulation of valves for pneumatic-hydraulic system of liquid-propellant rocket engine (LPRE) is given. Technique is based on approach of computational hydrodynamics (Computational Fluid Dynamics - CFD). The simulation of a differential valve used in closed circuit LPRE supply pipes of fuel components is performed to show technique abilities. A schematic and operation algorithm of this valve type is described in detail. Also assumptions made in the construction of the geometric model of the hydraulic path of the valve are described in detail. The calculation procedure for determining valve hydraulic characteristics is given. Based on these calculations certain hydraulic characteristics of the valve are given. Some ways of usage of the described simulation technique for research the static and dynamic characteristics of the elements of the pneumatic-hydraulic system of LPRE are proposed.

Characterization of typical platelet injector flow configurations. [liquid propellant rocket engines

NASA Technical Reports Server (NTRS)

Hickox, C. E.

1975-01-01

A study to investigate the hydraulic atomization characteristics of several novel injector designs for use in liquid propellant rocket engines is presented. The injectors were manufactured from a series of thin stainless steel platelets through which orifices were very accurately formed by a photoetching process. These individual platelets were stacked together and the orifices aligned so as to produce flow passages of prescribed geometry. After alignment, the platelets were bonded into a single, 'platelet injector', unit by a diffusion bonding process. Because of the complex nature of the flow associated with platelet injectors, it was necessary to use experimental techniques, exclusively, throughout the study. Large scale models of the injectors were constructed from aluminum plates and the appropriate fluids were modeled using a glycerol-water solution. Stop-action photographs of test configurations, using spark-shadowgraph or stroboscopic back-lighting, are shown.

A Historical Systems Study of Liquid Rocket Engine Throttling Capabilities

NASA Technical Reports Server (NTRS)

Betts, Erin M.; Frederick, Robert A., Jr.

2010-01-01

This is a comprehensive systems study to examine and evaluate throttling capabilities of liquid rocket engines. The focus of this study is on engine components, and how the interactions of these components are considered for throttling applications. First, an assessment of space mission requirements is performed to determine what applications require engine throttling. A background on liquid rocket engine throttling is provided, along with the basic equations that are used to predict performance. Three engines are discussed that have successfully demonstrated throttling. Next, the engine system is broken down into components to discuss special considerations that need to be made for engine throttling. This study focuses on liquid rocket engines that have demonstrated operational capability on American space launch vehicles, starting with the Apollo vehicle engines and ending with current technology demonstrations. Both deep throttling and shallow throttling engines are discussed. Boost and sustainer engines have demonstrated throttling from 17% to 100% thrust, while upper stage and lunar lander engines have demonstrated throttling in excess of 10% to 100% thrust. The key difficulty in throttling liquid rocket engines is maintaining an adequate pressure drop across the injector, which is necessary to provide propellant atomization and mixing. For the combustion chamber, cooling can be an issue at low thrust levels. For turbomachinery, the primary considerations are to avoid cavitation, stall, surge, and to consider bearing leakage flows, rotordynamics, and structural dynamics. For valves, it is necessary to design valves and actuators that can achieve accurate flow control at all thrust levels. It is also important to assess the amount of nozzle flow separation that can be tolerated at low thrust levels for ground testing.

Metallized Gelled Propellants: Oxygen/RP-1/Aluminum Rocket Engine Calorimeter Heat Transfer Measurements and Analysis

NASA Technical Reports Server (NTRS)

Palaszewski, Bryan

1997-01-01

A set of analyses was conducted to determine the heat transfer characteristics of metallized gelled liquid propellants in a rocket engine. The analyses used the data from experiments conducted with a small 30- to 40-lbf thrust engine composed of a modular injector, igniter, chamber and nozzle. The fuels used were traditional liquid RP-1 and gelled RP-1 with 0-wt %, 5-wt%, and 55-wt% loadings of aluminum with silicon dioxide gellant, and gaseous oxygen as the oxidizer. Heat transfer was computed based on measurements using calorimeter rocket chamber and nozzle hardware with a total of 31 cooling channels. A gelled fuel coating formed in the 0-, 5- and 55-wt% engines, and the coating was composed of unburned gelled fuel and partially combusted RP-1. The coating caused a large decrease in calorimeter engine heat flux in the last half of the chamber for the 0- and 5-wt% RP-1/Al. This heat flux reduction effect was analyzed by comparing engine runs and the changes in the heat flux during a run as well as from run to run. Heat transfer and time-dependent heat flux analyses and interpretations are provided. The 5- and 55-wt% RP-1/Al fueled engines had the highest chamber heat fluxes, with the 5-wt% fuel having the highest throat flux. This result is counter to the predicted result, where the 55 wt% fuel has the highest combustion and throat temperature, and therefore implies that it would deliver the highest throat heat flux. The 5-wt% RP-1/Al produced the most influence on the engine heat transfer and the heat flux reduction was caused by the formation of a gelled propellant layer in the chamber and nozzle.

Hydrazine engine plume contamination mapping. [measuring instruments for rocket exhaust from liquid propellant rocket engines

NASA Technical Reports Server (NTRS)

Chirivella, J. E.

1975-01-01

Instrumentation for the measurement of plume exhaust specie deposition rates were developed and demonstrated. The instruments, two sets of quartz crystal microbalances, were designed for low temperature operation in the back flow and variable temperature operation in the core flow regions of an exhaust plume. These quartz crystal microbalances performed nominally, and measurements of exhaust specie deposition rates for 8400 number of pulses for a 0.1-lb monopropellant thruster are reported.

Liquid and gelled sprays for mixing hypergolic propellants using an impinging jet injection system

NASA Astrophysics Data System (ADS)

James, Mark D.

The characteristics of sprays produced by liquid rocket injectors are important in understanding rocket engine ignition and performance. The includes, but is not limited to, drop size distribution, spray density, drop velocity, oscillations in the spray, uniformity of mixing between propellants, and the spatial distribution of drops. Hypergolic ignition and the associated ignition delay times are also important features in rocket engines, providing high reliability and simplicity of the ignition event. The ignition delay time is closely related to the level and speed of mixing between a hypergolic fuel and oxidizer, which makes the injection method and conditions crucial in determining the ignition performance. Although mixing and ignition of liquid hypergolic propellants has been studied for many years, the processes for injection, mixing, and ignition of gelled hypergolic propellants are less understood. Gelled propellants are currently under investigation for use in rocket injectors to combine the advantages of solid and liquid propellants, although not without their own difficulties. A review of hypergolic ignition has been conducted for selected propellants, and methods for achieving ignition have been established. This research is focused on ignition using the liquid drop-on-drop method, as well as the doublet impinging jet injector. The events leading up to ignition, known as pre-ignition stage are discussed. An understanding of desirable ignition and combustion performance requires a study of the effects of injection, temperature, and ambient pressure conditions. A review of unlike-doublet impinging jet injection mixing has also been conducted. This includes mixing factors in reactive and non-reactive sprays. Important mixing factors include jet momentum, jet diameter and length, impingement angle, mass distribution, and injector configuration. An impinging jet injection system is presented using an electro-mechanically driven piston for injecting liquid and gelled hypergolic propellants. A calibration of the system is done with water in preparation for hypergolic injection, and characteristics of individual water and gelled JP-8 jets are studied at velocities in the range of 3 ft/s to 61 ft/s. The piston response is also analyzed to characterize the startup and steady state liquid jet velocities using orifices of 0.02" in diameter. Using this injection system, water and gelled JP-8 sprays are formed and compared across injection velocities of 30 ft/s to 121 ft/s. The comparison includes sheet shape and disintegration, total number of drops, drop size distributions, drop eccentricity, most populated drop bin size, and mean drop sizes. A test matrix for investigating the effects of mixing on ignition of MMH and IRFNA through different injection conditions are presented. First, water and IRFNA are injected to create a spray in the combustion chamber in order to verify effectiveness of test procedures and the test hardware. Next, injection of the hypergolic propellants MMH and IRFNA are done in accordance to the test matrix, although ignition was not observed as expected. These injections are followed by simple drop-on-drop tests to investigate propellant quality and ignition delay. Drop tests are performed with propellants IRFNA/MMH, and again with H2O2/Block 0 as possible propellant replacements for the proposed test plan.

Recent Advancements in Propellant Densification

NASA Technical Reports Server (NTRS)

McNelis, Nancy B.; Tomsik, Thomas M.

1998-01-01

Next-generation launch vehicles demand several technological improvements to achieve lower cost and more reliable access to space. One technology area whose performance gains may far exceed others is densified propellants. The ideal rocket engine propellant is characterized by high specific impulse, high density, and low vapor pressure. A propellant combination of liquid hydrogen and liquid oxygen (LH2/LOX) is one of the highest performance propellants, but LH2 stored at standard conditions has a relatively low density and high vapor pressure. Propellant densification can significantly improve this propellant's properties relative to vehicle design and engine performance. Vehicle performance calculations based on an average of existing launch vehicles indicate that densified propellants may allow an increase in payload mass of up to 5 percent. Since the NASA Lewis Research Center became involved with the National Aerospace Plane program in the 1980's, it has been leading the way in making densified propellants a viable fuel for next-generation launch vehicles. Lewis researchers have been working to provide a method and critical data for continuous production of densified hydrogen and oxygen.

Injection Principles from Combustion Studies in a 200-Pound-Thrust Rocket Engine Using Liquid Oxygen and Heptane

NASA Technical Reports Server (NTRS)

Heidmann, M. F.; Auble, C. M.

1955-01-01

The importance of atomizing and mixing liquid oxygen and heptane was studied in a 200-pound-thrust rocket engine. Ten injector elements were used with both steel and transparent chambers. Characteristic velocity was measured over a range of mixture ratios. Combustion gas-flow and luminosity patterns within the chamber were obtained by photographic methods. The results show that, for efficient combustion, the propellants should be both atomized and mixed. Heptane atomization controlled the combustion rate to a much larger extent than oxygen atomization. Induced mixing, however, was required to complete combustion in the smallest volume. For stable, high-efficiency combustion and smooth engine starts, mixing after atomization was most promising.

High-Temperature Polymer Composites Tested for Hypersonic Rocket Combustor Backup Structure

NASA Technical Reports Server (NTRS)

Sutter, James K.; Shin, E. Eugene; Thesken, John C.; Fink, Jeffrey E.

2005-01-01

Significant component weight reductions are required to achieve the aggressive thrust-toweight goals for the Rocket Based Combined Cycle (RBCC) third-generation, reusable liquid propellant rocket engine, which is one possible engine for a future single-stage-toorbit vehicle. A collaboration between the NASA Glenn Research Center and Boeing Rocketdyne was formed under the Higher Operating Temperature Propulsion Components (HOTPC) program and, currently, the Ultra-Efficient Engine Technology (UEET) Project to develop carbon-fiber-reinforced high-temperature polymer matrix composites (HTPMCs). This program focused primarily on the combustor backup structure to replace all metallic support components with a much lighter polymer-matrixcomposite- (PMC-) titanium honeycomb sandwich structure.

Novel Techniques for Quantification of Correlation Between Primary Liquid Jet Breakup and Downstream Spray Characteristics

DTIC Science & Technology

2016-05-08

unlimited. 5 1. Introduction Several liquid -fuelled combustion systems, such as liquid propellant rocket engines and gas turbines...AFRL-AFOSR-JP-TR-2016-0084 Novel techniques for quantification of correlation between primary liquid jet breakup and downstream spray characteristics...to 17 Apr 2016 4.  TITLE AND SUBTITLE Novel techniques for quantification of correlation between primary liquid jet breakup and downstream spray

Novel Techniques for Quantification of Correlation Between Primary Liquid Jet Breakup and Downstream Spray Characteristics

DTIC Science & Technology

2016-10-05

unlimited. 5 1. Introduction Several liquid -fuelled combustion systems, such as liquid propellant rocket engines and gas turbines...AFRL-AFOSR-JP-TR-2016-0084 Novel techniques for quantification of correlation between primary liquid jet breakup and downstream spray characteristics...to 17 Apr 2016 4.  TITLE AND SUBTITLE Novel techniques for quantification of correlation between primary liquid jet breakup and downstream spray

The performance of a piezoelectric-sensor-based SHM system under a combined cryogenic temperature and vibration environment

NASA Astrophysics Data System (ADS)

Qing, Xinlin P.; Beard, Shawn J.; Kumar, Amrita; Sullivan, Kevin; Aguilar, Robert; Merchant, Munir; Taniguchi, Mike

2008-10-01

A series of tests have been conducted to determine the survivability and functionality of a piezoelectric-sensor-based active structural health monitoring (SHM) SMART Tape system under the operating conditions of typical liquid rocket engines such as cryogenic temperature and vibration loads. The performance of different piezoelectric sensors and a low temperature adhesive under cryogenic temperature was first investigated. The active SHM system for liquid rocket engines was exposed to flight vibration and shock environments on a simulated large booster LOX-H2 engine propellant duct conditioned to cryogenic temperatures to evaluate the physical robustness of the built-in sensor network as well as operational survivability and functionality. Test results demonstrated that the developed SMART Tape system can withstand operational levels of vibration and shock energy on a representative rocket engine duct assembly, and is functional under the combined cryogenic temperature and vibration environment.

Deimos Methane-Oxygen Rocket Engine Test Results

NASA Astrophysics Data System (ADS)

Engelen, S.; Souverein, L. J.; Twigt, D. J.

This paper presents the results of the first DEIMOS Liquid Methane/Oxygen rocket engine test campaign. DEIMOS is an acronym for `Delft Experimental Methane Oxygen propulsion System'. It is a project performed by students under the auspices of DARE (Delft Aerospace Rocket Engineering). The engine provides a theoretical design thrust of 1800 N and specific impulse of 287 s at a chamber pressure of 40 bar with a total mass flow of 637 g/s. It has links to sustainable development, as the propellants used are one of the most promising so-called `green propellants'-combinations, currently under scrutiny by the industry, and the engine is designed to be reusable. This paper reports results from the provisional tests, which had the aim of verifying the engine's ability to fire, and confirming some of the design assumptions to give confidence for further engine designs. Measurements before and after the tests are used to determine first estimates on feed pressures, propellant mass flows and achieved thrust. These results were rather disappointing from a performance point of view, with an average thrust of a mere 3.8% of the design thrust, but nonetheless were very helpful. The reliability of ignition and stability of combustion are discussed as well. An initial assessment as to the reusability, the flexibility and the adaptability of the engine was made. The data provides insight into (methane/oxygen) engine designs, leading to new ideas for a subsequent design. The ultimate goal of this project is to have an operational rocket and to attempt to set an amateur altitude record.

The techniques of quality operations computational and experimental researches of the launch vehicles in the drawing-board stage

NASA Astrophysics Data System (ADS)

Rozhaeva, K.

2018-01-01

The aim of the researchis the quality operations of the design process at the stage of research works on the development of active on-Board system of the launch vehicles spent stages descent with liquid propellant rocket engines by simulating the gasification process of undeveloped residues of fuel in the tanks. The design techniques of the gasification process of liquid rocket propellant components residues in the tank to the expense of finding and fixing errors in the algorithm calculation to increase the accuracy of calculation results is proposed. Experimental modelling of the model liquid evaporation in a limited reservoir of the experimental stand, allowing due to the false measurements rejection based on given criteria and detected faults to enhance the results reliability of the experimental studies; to reduce the experiments cost.

Photographic Study of Combustion in a Rocket Engine I : Variation in Combustion of Liquid Oxygen and Gasoline with Seven Methods of Propellant Injection

NASA Technical Reports Server (NTRS)

Bellman, Donald R; Humphrey, Jack C

1948-01-01

Motion pictures at camera speeds up to 3000 frames per second were taken of the combustion of liquid oxygen and gasoline in a 100-pound-thrust rocket engine. The engine consisted of thin contour and injection plates clamped between two clear plastic sheets forming a two-dimensional engine with a view of the entire combustion chamber and nozzle. A photographic investigation was made of the effect of seven methods of propellant injection on the uniformity of combustion. From the photographs, it was found that the flame front extended almost to the faces of the injectors with most of the injection methods, all the injection systems resulted in a considerable nonuniformity of combustion, and luminosity rapidly decreased in the divergent part of the nozzle. Pressure vibration records indicated combustion vibrations that approximately corresponded to the resonant frequencies of the length and the thickness of the chamber. The combustion temperature divided by the molecular weight of the combustion gases as determined from the combustion photographs was about 50 to 70 percent of the theoretical value.

Russian Rocket Engine Test

NASA Technical Reports Server (NTRS)

1998-01-01

NASA engineers successfully tested a Russian-built rocket engine on November 4, 1998 at the Marshall Space Flight Center (MSFC) Advanced Engine Test Facility, which had been used for testing the Saturn V F-1 engines and Space Shuttle Main engines. The MSFC was under a Space Act Agreement with Lockheed Martin Astronautics of Denver to provide a series of test firings of the Atlas III propulsion system configured with the Russian-designed RD-180 engine. The tests were designed to measure the performance of the Atlas III propulsion system, which included avionics and propellant tanks and lines, and how these components interacted with the RD-180 engine. The RD-180 is powered by kerosene and liquid oxygen, the same fuel mix used in Saturn rockets. The RD-180, the most powerful rocket engine tested at the MSFC since Saturn rocket tests in the 1960s, generated 860,000 pounds of thrust. The test was the first test ever anywhere outside Russia of a Russian designed and built engine.

Around Marshall

NASA Image and Video Library

1998-11-04

NASA engineers successfully tested a Russian-built rocket engine on November 4, 1998 at the Marshall Space Flight Center (MSFC) Advanced Engine Test Facility, which had been used for testing the Saturn V F-1 engines and Space Shuttle Main engines. The MSFC was under a Space Act Agreement with Lockheed Martin Astronautics of Denver to provide a series of test firings of the Atlas III propulsion system configured with the Russian-designed RD-180 engine. The tests were designed to measure the performance of the Atlas III propulsion system, which included avionics and propellant tanks and lines, and how these components interacted with the RD-180 engine. The RD-180 is powered by kerosene and liquid oxygen, the same fuel mix used in Saturn rockets. The RD-180, the most powerful rocket engine tested at the MSFC since Saturn rocket tests in the 1960s, generated 860,000 pounds of thrust.

Modelling of evaporation of a dispersed liquid component in a chemically active gas flow

NASA Astrophysics Data System (ADS)

Kryukov, V. G.; Naumov, V. I.; Kotov, V. Yu.

1994-01-01

A model has been developed to investigate evaporation of dispersed liquids in chemically active gas flow. Major efforts have been directed at the development of algorithms for implementing this model. The numerical experiments demonstrate that, in the boundary layer, significant changes in the composition and temperature of combustion products take place. This gives the opportunity to more correctly model energy release processes in combustion chambers of liquid-propellant rocket engines, gas-turbine engines, and other power devices.

Propellant Feed Subsystem for the X-34 Main Propulsion System

NASA Technical Reports Server (NTRS)

McDonald, J. P.; Minor, R. B.; Knight, K. C.; Champion, R. H., Jr.; Russell, F. J., Jr.

1998-01-01

The Orbital Sciences Corporation X-34 vehicle demonstrates technologies and operations key to future reusable launch vehicles. The general flight performance goal of this unmanned rocket plane is Mach 8 flight at an altitude of 250,000 feet. The Main Propulsion System supplies liquid propellants to the main engine, which provides the primary thrust for attaining mission goals. Major NMS design and operational goals are aircraft-like ground operations, quick turnaround between missions, and low initial/operational costs. This paper reviews major design and analysis aspects of the X-34 propellant feed subsystem of the X-34 Main Propulsion System. Topics include system requirements, system design, the integration of flight and feed system performance, propellant acquisition at engine start, and propellant tank terminal drain.

JANNAF "Test and Evaluation Guidelines for Liquid Rocket Engines": Status and Application

NASA Technical Reports Server (NTRS)

Parkinson, Douglas; VanLerberghe, Wayne M.; Rahman, Shamim A.

2017-01-01

For many decades, the U.S. rocket propulsion industrial base has performed remarkably in developing complex liquid rocket engines that can propel critical payloads into service for the nation, as well as transport people and hardware for missions that open the frontiers of space exploration for humanity. This has been possible only at considerable expense given the lack of detailed guidance that captures the essence of successful practices and knowledge accumulated over five decades of liquid rocket engine development. In an effort to provide benchmarks and guidance for the next generation of rocket engineers, the Joint Army Navy NASA Air Force (JANNAF) Interagency Propulsion Committee published a liquid rocket engine (LRE) test and evaluation (T&E) guideline document in 2012 focusing on the development challenges and test verification considerations for liquid rocket engine systems. This document has been well received and applied by many current LRE developers as a benchmark and guidance tool, both for government-driven applications as well as for fully commercial ventures. The USAF Space and Missile Systems Center (SMC) has taken an additional near-term step and is directing activity to adapt and augment the content from the JANNAF LRE T&E guideline into a standard for potential application to future USAF requests for proposals for LRE development initiatives and launch vehicles for national security missions. A draft of this standard was already sent out for review and comment, and is intended to be formally approved and released towards the end of 2017. The acceptance and use of the LRE T&E guideline is possible through broad government and industry participation in the JANNAF liquid propulsion committee and associated panels. The sponsoring JANNAF community is expanding upon this initial baseline version and delving into further critical development aspects of liquid rocket propulsion testing at the integrated stage level as well as engine component level, in order to advance the state of the practice. The full participation of the entire U.S. rocket propulsion industrial base is invited and expected at this opportune moment in the continuing advancement of spaceflight technology.

The SKYLON Spaceplane

NASA Astrophysics Data System (ADS)

Varvill, R.; Bond, A.

SKYLON is a single stage to orbit (SSTO) winged spaceplane designed to give routine low cost access to space. At a gross takeoff weight of 275 tonnes of which 220 tonnes is propellant the vehicle is capable of placing 12 tonnes into an equatorial low Earth orbit. The vehicle configuration consists of a slender fuselage containing the propellant tankage and payload bay with delta wings located midway along the fuselage carrying the SABRE engines in axisymmetric nacelles on the wingtips. The vehicle takes off and lands horizontally on it's own undercarriage. The fuselage is constructed as a multilayer structure consisting of aeroshell, insulation, structure and tankage. SKYLON employs extant or near term materials technology in order to minimise development cost and risk. The SABRE engines have a dual mode capability. In rocket mode the engine operates as a closed cycle liquid oxygen/liquid hydrogen high specific impulse rocket engine. In airbreathing mode (from takeoff to Mach 5) the liquid oxygen flow is replaced by atmospheric air, increasing the installed specific impulse 3-6 fold. The airflow is drawn into the engine via a 2 shock axisymmetric intake and cooled to cryogenic temperatures prior to compression. The hydrogen fuel flow acts as a heat sink for the closed cycle helium loop before entering the main combustion chamber.

Advanced high pressure engine study for mixed-mode vehicle applications

NASA Technical Reports Server (NTRS)

Luscher, W. P.; Mellish, J. A.

1977-01-01

High pressure liquid rocket engine design, performance, weight, envelope, and operational characteristics were evaluated for a variety of candidate engines for use in mixed-mode, single-stage-to-orbit applications. Propellant property and performance data were obtained for candidate Mode 1 fuels which included: RP-1, RJ-5, hydrazine, monomethyl-hydrazine, and methane. The common oxidizer was liquid oxygen. Oxygen, the candidate Mode 1 fuels, and hydrogen were evaluated as thrust chamber coolants. Oxygen, methane, and hydrogen were found to be the most viable cooling candidates. Water, lithium, and sodium-potassium were also evaluated as auxiliary coolant systems. Water proved to be the best of these, but the system was heavier than those systems which cooled with the engine propellants. Engine weight and envelope parametric data were established for candidate Mode 1, Mode 2, and dual-fuel engines. Delivered engine performance data were also calculated for all candidate Mode 1 and dual-fuel engines.

Rocket Engine Innovations Advance Clean Energy

NASA Technical Reports Server (NTRS)

2012-01-01

During launch countdown, at approximately T-7 seconds, the Space Shuttle Main Engines (SSMEs) roar to life. When the controllers indicate normal operation, the solid rocket boosters ignite and the shuttle blasts off. Initially, the SSMEs throttle down to reduce stress during the period of maximum dynamic pressure, but soon after, they throttle up to propel the orbiter to 17,500 miles per hour. In just under 9 minutes, the three SSMEs burn over 1.6 million pounds of propellant, and temperatures inside the main combustion chamber reach 6,000 F. To cool the engines, liquid hydrogen circulates through miles of tubing at -423 F. From 1981to 2011, the Space Shuttle fleet carried crew and cargo into orbit to perform a myriad of unprecedented tasks. After 30 years and 135 missions, the feat of engineering known as the SSME boasted a 100-percent flight success rate.

Hybrid rocket propellants from lunar material

NASA Astrophysics Data System (ADS)

Sparks, Douglas R.

This paper examines the use of lunar material for hybrid rocket propellants. Liquid oxygen is identified as the primary oxidizer and metals such as aluminum, magnesium, calcium, titanium and silicon are compared as possible fuels. Due to the reduced transportation costs, the use of lunar materials for both oxidizer and fuel will dramatically reduce the cost of a sustained space program. The advantage of hybrid rocket systems over liquid and solid rockets is discussed. It is pointed out that this type of hybrid rocket propellant could also be obtained from asteroidal and planetary soils, thereby facilitating the exploration and industrialization of the inner solar system.

Green Propellant Test Capabilities of the Altitude Combustion Stand at the NASA Glenn Research Center

NASA Technical Reports Server (NTRS)

Kubiak, Jonathan M.; Arnett, Lori A.

2016-01-01

The NASA Glenn Research Center (GRC) is committed to providing simulated altitude rocket test capabilities to NASA programs, other government agencies, private industry partners, and academic partners. A primary facility to support those needs is the Altitude Combustion Stand (ACS). ACS provides the capability to test combustion components at a simulated altitude up to 100,000 ft. (approx.0.2 psia/10 Torr) through a nitrogen-driven ejector system. The facility is equipped with an axial thrust stand, gaseous and cryogenic liquid propellant feed systems, data acquisition system with up to 1000 Hz recording, and automated facility control system. Propellant capabilities include gaseous and liquid hydrogen, gaseous and liquid oxygen, and liquid methane. A water-cooled diffuser, exhaust spray cooling chamber, and multi-stage ejector systems can enable run times up to 180 seconds to 16 minutes. The system can accommodate engines up to 2000-lbf thrust, liquid propellant supply pressures up to 1800 psia, and test at the component level. Engines can also be fired at sea level if needed. The NASA GRC is in the process of modifying ACS capabilities to enable the testing of green propellant (GP) thrusters and components. Green propellants are actively being explored throughout government and industry as a non-toxic replacement to hydrazine monopropellants for applications such as reaction control systems or small spacecraft main propulsion systems. These propellants offer increased performance and cost savings over hydrazine. The modification of ACS is intended to enable testing of a wide range of green propellant engines for research and qualification-like testing applications. Once complete, ACS will have the capability to test green propellant engines up to 880 N in thrust, thermally condition the green propellants, provide test durations up to 60 minutes depending on thrust class, provide high speed control and data acquisition, as well as provide advanced imaging and diagnostics such as infrared (IR) imaging.

Saturn Apollo Program

NASA Image and Video Library

1960-01-01

This chart provides the vital statistics for the F-1 rocket engine. Developed by Rocketdyne, under the direction of the Marshall Space Flight Center, the F-1 engine was utilized in a cluster of five engines to propel the Saturn V's first stage, the S-IC. Liquid oxygen and kerosene were used as its propellant. Initially rated at 1,500,000 pounds of thrust, the engine was later uprated to 1,522,000 pounds of thrust after the third Saturn V launch (Apollo 8, the first marned Saturn V mission) in December 1968. The cluster of five F-1 engines burned over 15 tons of propellant per second, during its two and one-half minutes of operation, to take the vehicle to a height of about 36 miles and to a speed of about 6,000 miles per hour.

Design of a Hybrid Propulsion System for Orbit Raising Applications

NASA Astrophysics Data System (ADS)

Boman, N.; Ford, M.

2004-10-01

A trade off between conventional liquid apogee engines used for orbit raising applications and hybrid rocket engines (HRE) has been performed using a case study approach. Current requirements for lower cost and enhanced safety places hybrid propulsion systems in the spotlight. For evaluating and design of a hybrid rocket engine a parametric engineering code is developed, based on the combustion chamber characteristics of selected propellants. A single port cylindrical section of fuel grain is considered. Polyethylene (PE) and hydroxyl-terminated polybutadiene (HTPB) represents the fuels investigated. The engine design is optimized to minimize the propulsion system volume and mass, while keeping the system as simple as possible. It is found that the fuel grain L/D ratio boundary condition has a major impact on the overall hybrid rocket engine design.

An overview of in-flight plume diagnostics for rocket engines

NASA Technical Reports Server (NTRS)

Madzsar, G. C.; Bickford, R. L.; Duncan, D. B.

1992-01-01

An overview and progress report of the work performed or sponsored by LeRC toward the development of in-flight plume spectroscopy technology for health and performance monitoring of liquid propellant rocket engines are presented. The primary objective of this effort is to develop technology that can be utilized on any flight engine. This technology will be validated by a hardware demonstration of a system capable of being retrofitted onto the Space Shuttle Main Engines for spectroscopic measurements during flight. The philosophy on system definition and status on the development of instrumentation, optics, and signal processing with respect to implementation on a flight engine are discussed.

Robust Exploration and Commercial Missions to the Moon Using Nuclear Thermal Rocket Propulsion and Lunar Liquid Oxygen Derived from FeO-Rich Pyroclasitc Deposits

NASA Technical Reports Server (NTRS)

Borowski, Stanley K.; Ryan, Stephen W.; Burke, Laura M.; McCurdy, David R.; Fittje, James E.; Joyner, Claude R.

2018-01-01

The nuclear thermal rocket (NTR) has frequently been identified as a key space asset required for the human exploration of Mars. This proven technology can also provide the affordable access through cislunar space necessary for commercial development and sustained human presence on the Moon. It is a demonstrated technology capable of generating both high thrust and high specific impulse (I(sub sp) approx. 900 s) twice that of today's best chemical rockets. Nuclear lunar transfer vehicles-consisting of a propulsion stage using three approx. 16.5-klb(sub f) small nuclear rocket engines (SNREs), an in-line propellant tank, plus the payload-are reusable, enabling a variety of lunar missions. These include cargo delivery and crewed lunar landing missions. Even weeklong ''tourism'' missions carrying passengers into lunar orbit for a day of sightseeing and picture taking are possible. The NTR can play an important role in the next phase of lunar exploration and development by providing a robust in-space lunar transportation system (LTS) that can allow initial outposts to evolve into settlements supported by a variety of commercial activities such as in-situ propellant production used to supply strategically located propellant depots and transportation nodes. The use of lunar liquid oxygen (LLO2) derived from iron oxide (FeO)-rich volcanic glass beads, found in numerous pyroclastic deposits on the Moon, can significantly reduce the launch mass requirements from Earth by enabling reusable, surface-based lunar landing vehicles (LLVs)that use liquid oxygen and hydrogen (LO2/LH2) chemical rocket engines. Afterwards, a LO2/LH2 propellant depot can be established in lunar equatorial orbit to supply the LTS. At this point a modified version of the conventional NTR-called the LO2-augmented NTR, or LANTR-is introduced into the LTS allowing bipropellant operation and leveraging the mission benefits of refueling with lunar-derived propellants for Earth return. The bipropellant LANTR engine utilizes the large divergent section of its nozzle as an ''afterburner'' into which oxygen is injected and supersonically combusted with nuclear preheated hydrogen emerging from the engine's choked sonic throat-essentially ''scramjet propulsion in reverse.'' By varying the oxygen-to-hydrogen mixture ratio, LANTR engines can operate over a range of thrust and I(sub sp) values while the reactor core power level remains relatively constant. A LANTR-based LTS offers unique mission capabilities including short-transit-time crewed cargo transports. Even a ''commuter'' shuttle service may be possible allowing ''one-way'' trip times to and from the Moon on the order of 36 hours or less. If only 1% of the extracted LLO2 propellant from identified resource sites were available for use in lunar orbit, such a supply could support daily commuter flights to the Moon for many thousands of years! This report outlines an evolutionary architecture and examines a variety of mission types and transfer vehicle designs, along with the increasing demands on LLO2 production as mission complexity and velocity change delta V requirements increase. A comparison of vehicle features and engine operating characteristics, for both NTR and LANTR engines, is also provided along with a discussion of the propellant production and mining requirements associated with using FeO-rich volcanic glass as source material.

Comparative analysis of the designs and implementation of vehicles based on reactive propulsion proposed during the nineteenth and beginning of the twentieth centuries

NASA Technical Reports Server (NTRS)

Sokolskiy, V. N.

1977-01-01

Examination of the presently known historical scientific literature related to the problem of reactive flight indicates that considerable attention had already been given to the idea of reactive propulsion in the nineteenth century; about thirty designs for reaction flying vehicles were proposed during this period. However, the authors of a majority of the designs limited themselves only to a presentation of a diagram of the engine or an account of the principle of its operation, giving neither plans for its structural development nor precise calculations of the amount of energy required for accomplishing reaction flight. None of these authors considered the reaction flying vehicle as an object of variable mass, their choice of energy sources was extremely random, and the theory of the flight of reaction flying vehicles remained completely undeveloped. Early rocket designs of Nezhdanovsky, Ganswindt, Goddard, Tsiolkovsky, and others are examined and the evolution of liquid-propellant rocket engines, solid-propellant rocket engines, and jet aircraft engines is reviewed.

Kerosene-Fuel Engine Testing Under Way

NASA Image and Video Library

2003-11-17

NASA Stennis Space Center engineers conducted a successful cold-flow test of an RS-84 engine component Sept. 24. The RS-84 is a reusable engine fueled by rocket propellant - a special blend of kerosene - designed to power future flight vehicles. Liquid oxygen was blown through the RS-84 subscale preburner to characterize the test facility's performance and the hardware's resistance. Engineers are now moving into the next phase, hot-fire testing, which is expected to continue into February 2004. The RS-84 engine prototype, developed by the Rocketdyne Propulsion and Power division of The Boeing Co. of Canoga Park, Calif., is one of two competing Rocket Engine Prototype technologies - a key element of NASA's Next Generation Launch Technology program.

Kerosene-Fuel Engine Testing Under Way

NASA Technical Reports Server (NTRS)

2003-01-01

NASA Stennis Space Center engineers conducted a successful cold-flow test of an RS-84 engine component Sept. 24. The RS-84 is a reusable engine fueled by rocket propellant - a special blend of kerosene - designed to power future flight vehicles. Liquid oxygen was blown through the RS-84 subscale preburner to characterize the test facility's performance and the hardware's resistance. Engineers are now moving into the next phase, hot-fire testing, which is expected to continue into February 2004. The RS-84 engine prototype, developed by the Rocketdyne Propulsion and Power division of The Boeing Co. of Canoga Park, Calif., is one of two competing Rocket Engine Prototype technologies - a key element of NASA's Next Generation Launch Technology program.

LOX tank installation

NASA Image and Video Library

2011-06-08

Construction of the A-3 Test Stand at Stennis Space Center continued June 8 with installation of a 35,000-gallon liquid oxygen tank atop the steel structure. The stand is being built to test next-generation rocket engines that will carry humans into deep space once more. The LOX tank and a liquid hydrogen tank to be installed atop the stand later will provide propellants for testing the engines. The A-3 Test Stand is scheduled for completion and activation in 2013.

Design criteria monograph on turbopump systems

NASA Technical Reports Server (NTRS)

1975-01-01

Turbopump assembly for modern liquid propellant rocket engine is complete system in itself. It consists of many components, some of which are themselves subsystems. Monograph deals with turbopump as system, covering selection of proper system type for each application and integration of components into working system.

Combustion Stability Analyses of Coaxial Element Injectors with Liquid Oxygen/Liquid Methane Propellants

NASA Technical Reports Server (NTRS)

Hulka, J. R.

2010-01-01

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 a flight-qualified engine system, so limited test data and analysis results are available at this stage of early development. NASA has funded several hardware-oriented activities 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, the NASA Marshall Space Flight Center has conducted combustion stability analyses of several of the configurations. This paper presents test data and analyses of combustion stability from the recent PCAD-funded test programs at the NASA MSFC. These test programs used swirl coaxial element injectors with liquid oxygen and liquid methane propellants. Oxygen was injected conventionally in the center of the coaxial element, and swirl was provided by tangential entry slots. Injectors with 28-element and 40-element patterns were tested with several configurations of combustion chambers, including ablative and calorimeter spool sections, and several configurations of fuel injection design. Low frequency combustion instability (chug) occurred with both injectors, and high-frequency combustion instability occurred at the first tangential (1T) transverse mode with the 40-element injector. In most tests, a transition between high-amplitude chug with gaseous methane flow and low-amplitude chug with liquid methane flow was readily observed. Chug analyses of both conditions were conducted using techniques from Wenzel and Szuch and from the Rocket Combustor Interactive Design and Analysis (ROCCID) code. The 1T mode instability occurred in several tests and was apparent by high-frequency pressure measurements as well as dramatic increases in calorimeter-measured heat flux throughout the chamber. Analyses of the transverse mode were conducted with ROCCID and empirical methods such as Hewitt d/V. This paper describes the test hardware configurations, test data, analysis methods, and presents results of the various analyses.

Ares First Stage "Systemology" - Combining Advanced Systems Engineering and Planning Tools to Assure Mission Success

NASA Technical Reports Server (NTRS)

Seiler, James; Brasfield, Fred; Cannon, Scott

2008-01-01

Ares is an integral part of NASA s Constellation architecture that will provide crew and cargo access to the International Space Station as well as low earth orbit support for lunar missions. Ares replaces the Space Shuttle in the post 2010 time frame. Ares I is an in-line, two-stage rocket topped by the Orion Crew Exploration Vehicle, its service module, and a launch abort system. The Ares I first stage is a single, five-segment reusable solid rocket booster derived from the Space Shuttle Program's reusable solid rocket motor. The Ares second or upper stage is propelled by a J-2X main engine fueled with liquid oxygen and liquid hydrogen. This paper describes the advanced systems engineering and planning tools being utilized for the design, test, and qualification of the Ares I first stage element. Included are descriptions of the current first stage design, the milestone schedule requirements, and the marriage of systems engineering, detailed planning efforts, and roadmapping employed to achieve these goals.

Bistable (latching) solenoid actuated propellant isolation valve

NASA Technical Reports Server (NTRS)

Wichmann, H.; Deboi, H. H.

1979-01-01

The design, fabrication, assembly and test of a development configuration bistable (latching) solenoid actuated propellant isolation valve suitable for the control hydrazine and liquid fluorine to an 800 pound thrust rocket engine is described. The valve features a balanced poppet, utilizing metal bellows, a hard poppet/seat interface and a flexure support system for the internal moving components. This support system eliminates sliding surfaces, thereby rendering the valve free of self generated particles.

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

NASA Technical Reports Server (NTRS)

Jassowski, Donald M.

1993-01-01

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.

Rocket propulsion research at Lewis Research Center

NASA Technical Reports Server (NTRS)

Dawson, Virginia P.

1992-01-01

A small contingent of engineers at NASA LeRC pioneered the basic research on liquid propellants for rockets shortly after World War 2. Carried on through the 1950s, this work influenced the important early decisions made by Abe Silverstein when he took charge of the Office of Space Flight Programs for NASA. He strongly supported the development of liquid hydrogen as a propulsion fuel in the face of resistance from Wernher von Braun. Members of the LeRC staff played an important role in bringing liquid hydrogen technology to the point of reliability through their management of the Centaur Program. This paper demonstrates how the personality and engineering intuition of Abe Silverstein shaped the Centaur program and left a lasting imprint on the laboratory research tradition. Many of the current leaders of LeRC received their first hands-on engineering experience when they worked on the Centaur program in the 1960s.

Rocket Propulsion Research at Lewis Research Center

NASA Technical Reports Server (NTRS)

Dawson, Virginia P.

1992-01-01

A small contingent of engineers at NASA Lewis Research Center pioneered in basic research on liquid propellants for rockets shortly after World War II. Carried on through the 1950s, this work influenced the important early decisions made by Abe Silverstein when he took charge of the Office of Space Flight Programs for NASA. He strongly supported the development of liquid hydrogen as a propulsion fuel in the face of resistance from Wernher von Braun. Members of the Lewis staff played an important role in bringing liquid hydrogen technology to the point of reliability through their management of the Centaur Program. This paper demonstrates how the personality and engineering intuition of Abe Silverstein shaped the Centaur program and left a lasting imprint on the laboratory research tradition. Many of the current leaders of Lewis Research Center received their first hands-on engineering experience when they worked on the Centaur program in the 1960s.

Nuclear Thermal Propulsion: Past, Present, and a Look Ahead

NASA Technical Reports Server (NTRS)

Borowski, Stanley K.

2014-01-01

NTR: High thrust high specific impulse (2 x LOXLH2 chemical) engine uses high power density fission reactor with enriched uranium fuel as thermal power source. Reactor heat is removed using H2 propellant which is then exhausted to produce thrust. Conventional chemical engine LH2 tanks, turbo pumps, regenerative nozzles and radiation-cooled shirt extensions used -- NTR is next evolutionary step in high performance liquid rocket engines.

Study of solid rocket motor for space shuttle booster, volume 2, book 2

NASA Technical Reports Server (NTRS)

1972-01-01

A technical analysis of the solid propellant rocket engines for use with the space shuttle is presented. The subjects discussed are: (1) solid rocket motor stage recovery, (2) environmental effects, (3) man rating of the solid propellant rocket engines, (4) system safety analysis, (5) ground support equipment, and (6) transportation, assembly, and checkout.

Liquid Oxygen/Liquid Methane Test Summary of the RS-18 Lunar Ascent Engine at Simulated Altitude Conditions at NASA White Sands Test Facility

NASA Technical Reports Server (NTRS)

Melcher, John C., IV; Allred, Jennifer K.

2009-01-01

Tests were conducted with the RS18 rocket engine using liquid oxygen (LO2) and liquid methane (LCH4) propellants under simulated altitude conditions at NASA Johnson Space Center White Sands Test Facility (WSTF). This project is part of NASA s Propulsion and Cryogenics Advanced Development (PCAD) project. "Green" propellants, such as LO2/LCH4, offer savings in both performance and safety over equivalently sized hypergolic propellant systems in spacecraft applications such as ascent engines or service module engines. Altitude simulation was achieved using the WSTF Large Altitude Simulation System, which provided altitude conditions equivalent up to approx.120,000 ft (approx.37 km). For specific impulse calculations, engine thrust and propellant mass flow rates were measured. Propellant flow rate was measured using a coriolis-style mass-flow meter and compared with a serial turbine-style flow meter. Results showed a significant performance measurement difference during ignition startup. LO2 flow ranged from 5.9-9.5 lbm/sec (2.7-4.3 kg/sec), and LCH4 flow varied from 3.0-4.4 lbm/sec (1.4-2.0 kg/sec) during the RS-18 hot-fire test series. Thrust was measured using three load cells in parallel. Ignition was demonstrated using a gaseous oxygen/methane spark torch igniter. Data was obtained at multiple chamber pressures, and calculations were performed for specific impulse, C* combustion efficiency, and thrust vector alignment. Test objectives for the RS-18 project are 1) conduct a shakedown of the test stand for LO2/methane lunar ascent engines, 2) obtain vacuum ignition data for the torch and pyrotechnic igniters, and 3) obtain nozzle kinetics data to anchor two-dimensional kinetics codes.

Ablative material testing for low-pressure, low-cost rocket engines

NASA Technical Reports Server (NTRS)

Richter, G. Paul; Smith, Timothy D.

1995-01-01

The results of an experimental evaluation of ablative materials suitable for the production of light weight, low cost rocket engine combustion chambers and nozzles are presented. Ten individual specimens of four different compositions of silica cloth-reinforced phenolic resin materials were evaluated for comparative erosion in a subscale rocket engine combustion chamber. Gaseous hydrogen and gaseous oxygen were used as propellants, operating at a nominal chamber pressure of 1138 kPa (165 psi) and a nominal mixture ratio (O/F) of 3.3. These conditions were used to thermally simulate operation with RP-1 and liquid oxygen, and achieved a specimen throat gas temperature of approximately 2456 K (4420 R). Two high-density composition materials exhibited high erosion resistance, while two low-density compositions exhibited approximately 6-75 times lower average erosion resistance. The results compare favorably with previous testing by NASA and provide adequate data for selection of ablatives for low pressure, low cost rocket engines.

Characterizing high-energy-density propellants for space propulsion applications

NASA Astrophysics Data System (ADS)

Kokan, Timothy

There exists wide ranging research interest in high-energy-density matter (HEDM) propellants as a potential replacement for existing industry standard fuels for liquid rocket engines. The U.S. Air Force Research Laboratory, the U.S. Army Research Lab, the NASA Marshall Space Flight Center, and the NASA Glenn Research Center each either recently concluded or currently has ongoing programs in the synthesis and development of these potential new propellants. In order to perform conceptual designs using these new propellants, most conceptual rocket engine powerhead design tools (e.g. NPSS, ROCETS, and REDTOP-2) require several thermophysical properties of a given propellant over a wide range of temperature and pressure. These properties include enthalpy, entropy, density, viscosity, and thermal conductivity. Very little thermophysical property data exists for most of these potential new HEDM propellants. Experimental testing of these properties is both expensive and time consuming and is impractical in a conceptual vehicle design environment. A new technique for determining these thermophysical properties of potential new rocket engine propellants is presented. The technique uses a combination of three different computational methods to determine these properties. Quantum mechanics and molecular dynamics are used to model new propellants at a molecular level in order to calculate density, enthalpy, and entropy. Additivity methods are used to calculate the kinematic viscosity and thermal conductivity of new propellants. This new technique is validated via a series of verification experiments of HEDM compounds. Results are provided for two HEDM propellants: quadricyclane and 2-azido-N,N-dimethylethanamine (DMAZ). In each case, the new technique does a better job than the best current computational methods at accurately matching the experimental data of the HEDM compounds of interest. A case study is provided to help quantify the vehicle level impacts of using HEDM propellants. The case study consists of the National Aeronautics and Space Administration's (NASA) Exploration Systems Architecture Study (ESAS) Lunar Surface Access Module (LSAM). The results of this study show that the use of HEDM propellants instead of hypergolic propellants can lower the gross weight of the LSAM and may be an attractive alternative to the current baseline hypergolic propellant choice.

[Progress in the protective medicine against [correction of aganist] rocket propellents].

PubMed

Hu, W X; Tan, C Y; Tan, S J; Jiang, J

1999-12-01

To review the progress in the major assignment, the organization and implementation of protection against liquid rocket propellent. The safety detection methods of the rocket [correction of rocked] propellent in the launching field were also discussed. Three steps of the sanitation and protection of the liquid propellent, the toxicity and the toxicology of hydrazine on central nervous system, blood circulatory system, assimilation system, respiratory system, immune system, liver, kidney, eye, skin and its hereditary toxicology were described. In addition, the clinical types of poisoning, the current principle and the common ways of the prevention and treatment of hydrazine and nitrogen oxides poisoning were summarized.

Multivariable optimization of liquid rocket engines using particle swarm algorithms

NASA Astrophysics Data System (ADS)

Jones, Daniel Ray

Liquid rocket engines are highly reliable, controllable, and efficient compared to other conventional forms of rocket propulsion. As such, they have seen wide use in the space industry and have become the standard propulsion system for launch vehicles, orbit insertion, and orbital maneuvering. Though these systems are well understood, historical optimization techniques are often inadequate due to the highly non-linear nature of the engine performance problem. In this thesis, a Particle Swarm Optimization (PSO) variant was applied to maximize the specific impulse of a finite-area combustion chamber (FAC) equilibrium flow rocket performance model by controlling the engine's oxidizer-to-fuel ratio and de Laval nozzle expansion and contraction ratios. In addition to the PSO-controlled parameters, engine performance was calculated based on propellant chemistry, combustion chamber pressure, and ambient pressure, which are provided as inputs to the program. The performance code was validated by comparison with NASA's Chemical Equilibrium with Applications (CEA) and the commercially available Rocket Propulsion Analysis (RPA) tool. Similarly, the PSO algorithm was validated by comparison with brute-force optimization, which calculates all possible solutions and subsequently determines which is the optimum. Particle Swarm Optimization was shown to be an effective optimizer capable of quick and reliable convergence for complex functions of multiple non-linear variables.

Study of solid rocket motors for a space shuttle booster. Volume 1: Executive summary

NASA Technical Reports Server (NTRS)

1972-01-01

An analysis of the solid propellant rocket engines for use with the space shuttle booster was conducted. A definition of the specific solid propellant rocket engine stage designs, development program requirements, production requirements, launch requirements, and cost data for each program phase were developed.

Method for providing real-time control of a gaseous propellant rocket propulsion system

NASA Technical Reports Server (NTRS)

Morris, Brian G. (Inventor)

1991-01-01

The new and improved methods and apparatus disclosed provide effective real-time management of a spacecraft rocket engine powered by gaseous propellants. Real-time measurements representative of the engine performance are compared with predetermined standards to selectively control the supply of propellants to the engine for optimizing its performance as well as efficiently managing the consumption of propellants. A priority system is provided for achieving effective real-time management of the propulsion system by first regulating the propellants to keep the engine operating at an efficient level and thereafter regulating the consumption ratio of the propellants. A lower priority level is provided to balance the consumption of the propellants so significant quantities of unexpended propellants will not be left over at the end of the scheduled mission of the engine.

Liquid rocket disconnects, couplings, fittings, fixed joints, and seals

NASA Technical Reports Server (NTRS)

1976-01-01

State of the art and design criteria for components used in liquid propellant rocket propulsion systems to contain and control the flow of fluids involved are discussed. Particular emphasis is placed on the design of components used in the engine systems of boosters and upper stages, and in spacecraft propulsion systems because of the high pressure and high vibration levels to which these components are exposed. A table for conversion of U.S. customary units to SI units is included with a glossary, and a list of NASA space vehicle design criteria monographs issued to September 1976.

X-34 Main Propulsion System-Selected Subsystem Analyses

NASA Technical Reports Server (NTRS)

Brown, T. M.; McDonald, J. P.; Knight, K. C.; Champion, R. H., Jr.

1998-01-01

The X-34 hypersonic flight vehicle is currently under development by Orbital Sciences Corporation (Orbital). The Main Propulsion System (MPS) has been designed around the liquid propellant Fastrac rocket engine currently under development at NASA Marshall Space Flight Center. This paper presents selected analyses of MPS subsystems and components. Topics include the integration of component and system level modeling of the LOX dump subsystem and a simple terminal bubble velocity analysis conducted to guide propellant feed line design.

Status on Technology Development of Optic Fiber-Coupled Laser Ignition System for Rocket Engine Applications

NASA Technical Reports Server (NTRS)

Trinh, Huu P.; Early, Jim; Osborne, Robin; Thomas, Matthew; Bossard, John

2003-01-01

To pursue technology developments for future launch vehicles, NASA/Marshall Space Flight Center (MSFC) is examining vortex chamber concepts for liquid rocket engine applications. Past studies indicated that the vortex chamber schemes potentially have a number of advantages over conventional chamber methods. Due to the nature of the vortex flow, relatively cooler propellant streams tend to flow along the chamber wall. Hence, the thruster chamber can be operated without the need of any cooling techniques. This vortex flow also creates strong turbulence, which promotes the propellant mixing process. Consequently, the subject chamber concept: not only offer system simplicity, but also enhance the combustion performance. Test results have shown that chamber performance is markedly high even at a low chamber length-to-diameter ratio. This incentive can be translated to a convenience in the thrust chamber packaging.

LOX/Hydrogen Coaxial Injector Atomization Test Program

NASA Technical Reports Server (NTRS)

Zaller, M.

1990-01-01

Quantitative information about the atomization of injector sprays is needed to improve the accuracy of computational models that predict the performance and stability margin of liquid propellant rocket engines. To obtain this data, a facility for the study of spray atomization is being established at NASA-Lewis to determine the drop size and velocity distributions occurring in vaporizing liquid sprays at supercritical pressures. Hardware configuration and test conditions are selected to make the cold flow simulant testing correspond as closely as possible to conditions in liquid oxygen (LOX)/gaseous H2 rocket engines. Drop size correlations from the literature, developed for liquid/gas coaxial injector geometries, are used to make drop size predictions for LOX/H2 coaxial injectors. The mean drop size predictions for a single element coaxial injector range from 0.1 to 2000 microns, emphasizing the need for additional studies of the atomization process in LOX/H2 engines. Selection of cold flow simulants, measured techniques, and hardware for LOX/H2 atomization simulations are discussed.

Design issues for lunar in situ aluminum/oxygen propellant rocket engines

NASA Technical Reports Server (NTRS)

Meyer, Michael L.

1992-01-01

Design issues for lunar ascent and lunar descent rocket engines fueled by aluminum/oxygen propellant produced in situ at the lunar surface were evaluated. Key issues are discussed which impact the design of these rockets: aluminum combustion, throat erosion, and thrust chamber cooling. Four engine concepts are presented, and the impact of combustion performance, throat erosion and thrust chamber cooling on overall engine design are discussed. The advantages and disadvantages of each engine concept are presented.

Evaluation of Impinging Stream Vortex Chamber Concepts for Liquid Rocket Engine Applications

NASA Technical Reports Server (NTRS)

Trinh, Huu P.; Bullard, Brad; Kopicz, Charles; Michaels, Scott; Turner, James (Technical Monitor)

2001-01-01

To pursue technology developments for future launch vehicles, NASA/Marshall Space Flight Center (MSFC) is examining vortex chamber concepts for liquid rocket engine applications. Past studies indicated that the vortex chamber schemes potentially have a number of advantages over conventional chamber methods. Due to the nature of the vortex flow, relatively cooler propellant streams tend to flow along the chamber wall. Hence, the thruster chamber can be operated without the need of any cooling techniques. This vortex flow also creates strong turbulence, which promotes the propellant mixing process. Consequently, the subject chamber concepts not only offer the system simplicity, but they also would enhance the combustion performance. The test results showed that the chamber performance was markedly high even at a low chamber length-to-diameter ratio (L/D). This incentive can be translated to a convenience in the thrust chamber packaging. Variations of the vortex chamber concepts have been introduced in the past few decades. These investigations include an ongoing work at Orbital Technologies Corporation (ORBITEC). By injecting the oxidizer tangentially at the chamber convergence and fuel axially at the chamber head end, Knuth et al. were able to keep the wall relatively cold. A recent investigation of the low L/D vortex chamber concept for gel propellants was conducted by Michaels. He used both triplet (two oxidizer and one fuel orifices) and unlike impinging schemes to inject propellants tangentially along the chamber wall. Michaels called the subject injection scheme as Impinging Stream Vortex Chamber (ISVC). His preliminary tests showed that high performance, with an Isp efficiency of 92%, can be obtained. MSFC and the U.S. Army are jointly investigating an application of the ISVC concept for the cryogenic oxygen/hydrocarbon propellant system. This vortex chamber concept is currently tested with gel propellants at AMCOM at Redstone Arsenal, Alabama. A version of this concept for the liquid oxygen (LOX)/hydrocarbon fuel (RPM) system has been derived from the one for the gel propellant.

Measurement of intact-core length of atomizing liquid jets by image deconvolution

NASA Technical Reports Server (NTRS)

Woodward, Roger; Burch, Robert; Kuo, Kenneth; Cheung, Fan-Bill

1993-01-01

The investigation of liquid jet breakup and spray development is critical to the understanding of combustion phenomena in liquid propellant rocket engines. Much work has been done to characterize low-speed liquid jet breakup and dilute sprays, but atomizing jets and dense sprays have yielded few quantitative measurements due to their high liquid load fractions and hence their optical opacity. Focus was on a characteristic of the primary breakup process of round liquid jets, namely the length of the intact-liquid core. The specific application considered is that of shear-coaxial-type rocket engine injectors in which liquid oxygen is injected through the center post while high velocity gaseous hydrogen is injected through a concentric annulus, providing a shear force to the liquid jet surface. Real-time x ray radiography, capable of imaging through the dense two-phase region surrounding the liquid core, is used to make the measurements. The intact-liquid-core length data were obtained and interpreted using two conceptually different methods to illustrate the effects of chamber pressure, gas-to-liquid momentum ratio, and cavitation.

Measurement of intact-core length of atomizing liquid jets by image deconvolution

NASA Astrophysics Data System (ADS)

Woodward, Roger; Burch, Robert; Kuo, Kenneth; Cheung, Fan-Bill

1993-11-01

The investigation of liquid jet breakup and spray development is critical to the understanding of combustion phenomena in liquid propellant rocket engines. Much work has been done to characterize low-speed liquid jet breakup and dilute sprays, but atomizing jets and dense sprays have yielded few quantitative measurements due to their high liquid load fractions and hence their optical opacity. Focus was on a characteristic of the primary breakup process of round liquid jets, namely the length of the intact-liquid core. The specific application considered is that of shear-coaxial-type rocket engine injectors in which liquid oxygen is injected through the center post while high velocity gaseous hydrogen is injected through a concentric annulus, providing a shear force to the liquid jet surface. Real-time x ray radiography, capable of imaging through the dense two-phase region surrounding the liquid core, is used to make the measurements. The intact-liquid-core length data were obtained and interpreted using two conceptually different methods to illustrate the effects of chamber pressure, gas-to-liquid momentum ratio, and cavitation.

Robust Exploration and Commercial Missions to the Moon Using LANTR Propulsion and Lunar Liquid Oxygen Derived from FeO-Rich Pyroclastic Deposits

NASA Technical Reports Server (NTRS)

Borowski, Stanley K.; Ryan, Stephen W.; Burke, Laura M.; McCurdy, David R.; Fittje, James E.; Joyner, Claude R.

2017-01-01

The nuclear thermal rocket (NTR) has frequently been identified as a key space asset required for the human exploration of Mars. This proven technology can also provide the affordable access through cislunar space necessary for commercial development and sustained human presence on the Moon. It is a demonstrated technology capable of generating both high thrust and high specific impulse (Isp approx.900 s) twice that of todays best chemical rockets. Nuclear lunar transfer vehicles consisting of a propulsion stage using three approx.16.5 klbf Small Nuclear Rocket Engines (SNREs), an in-line propellant tank, plus the payload can enable a variety of reusable lunar missions. These include cargo delivery and crewed lunar landing missions. Even weeklong tourism missions carrying passengers into lunar orbit for a day of sightseeing and picture taking are possible. The NTR can play an important role in the next phase of lunar exploration and development by providing a robust in-space lunar transportation system (LTS) that can allow initial outposts to evolve into settlements supported by a variety of commercial activities such as in-situ propellant production used to supply strategically located propellant depots and transportation nodes. The use of lunar liquid oxygen (LLO2) derived from iron oxide (FeO)-rich volcanic glass beads, found in numerous pyroclastic deposits on the Moon, can significantly reduce the launch mass requirements from Earth by enabling reusable, surface-based lunar landing vehicles (LLVs) using liquid oxygen/hydrogen (LO2/H2) chemical rocket engines. Afterwards, a LO2/H2 propellant depot can be established in lunar equatorial orbit to supply the LTS. At this point a modified version of the conventional NTR called the LOX-augmented NTR, or LANTR is introduced into the LTS allowing bipropellant operation and leveraging the mission benefits of refueling with lunar-derived propellants for Earth return. The bipropellant LANTR engine utilizes the large divergent section of its nozzle as an afterburner into which oxygen is injected and supersonically combusted with nuclear preheated hydrogen emerging from the engines choked sonic throat - essentially scramjet propulsion in reverse. By varying the oxygen-to-hydrogen mixture ratio, LANTR engines can operate over a range of thrust and Isp values while the reactor core power level remains relatively constant. A LANTR-based LTS offers unique mission capabilities including short transit time crewed cargo transports. Even a commuter shuttle service may be possible allowing one-way trip times to and from the Moon on the order of 36 hours or less. If only 1 of the extracted LLO2 propellant from identified resource sites were available for use in lunar orbit, such a supply could support daily commuter flights to the Moon for many thousands of years! The proposed paper outlines an evolutionary architecture and examines a variety of mission types and transfer vehicle designs, along with the increasing demands on LLO2 production as mission complexity and (Delta)V requirements increase. A comparison of vehicle features and engine operating characteristics, for both NTR and LANTR engines, is also provided along with a discussion of the propellant production and mining requirements associated with using FeO-rich volcanic glass as source material.

Early Rockets

NASA Image and Video Library

2004-04-15

This photograph is of the engine for the Redstone rocket. The Redstone ballistic missile was a high-accuracy, liquid-propelled, surface-to-surface missile developed by the Army Ballistic Missile Agency, Redstone Arsenal, in Huntsville, Alabama, under the direction of Dr. von Braun. The Redstone engine was a modified and improved version of the Air Force's Navaho cruise missile engine of the late forties. The A-series, as this would be known, utilized a cylindrical combustion chamber as compared with the bulky, spherical V-2 chamber. By 1951, the Army was moving rapidly toward the design of the Redstone missile, and the production was begun in 1952. Redstone rockets became the "reliable workhorse" for America's early space program. As an example of its versatility, the Redstone was utilized in the booster for Explorer 1, the first American satellite, with no major changes to the engine or missile.

Liquid Rocket Booster (LRB) for the Space Transportation System (STS) systems study. Appendix F: Performance and trajectory for ALS/LRB launch vehicles

NASA Technical Reports Server (NTRS)

1989-01-01

By simply combining two baseline pump-fed LOX/RP-1 Liquid Rocket Boosters (LRBs) with the Denver core, a launch vehicle (Option 1 Advanced Launch System (ALS)) is obtained that can perform both the 28.5 deg (ALS) mission and the polar orbit ALS mission. The Option 2 LRB was obtained by finding the optimum LOX/LH2 engine for the STS/LRB reference mission (70.5 K lb payload). Then this engine and booster were used to estimate ALS payload for the 28.5 deg inclination ALS mission. Previous studies indicated that the optimum number of STS/LRB engines is four. When the engine/booster sizing was performed, each engine had 478 K lb sea level thrust and the booster carried 625,000 lb of useable propellant. Two of these LRBs combined with the Denver core provided a launch vehicle that meets the payload requirements for both the ALS and STS reference missions. The Option 3 LRB uses common engines for the cores and boosters. The booster engines do not have the nozzle extension. These engines were sized as common ALS engines. An ALS launch vehicle that has six core engines and five engines per booster provides 109,100 lb payload for the 28.5 deg mission. Each of these LOX/LH2 LRBs carries 714,100 lb of useable propellant. It is estimated that the STS/LRB reference mission payload would be 75,900 lb.

Technology Challenges for Deep-Throttle Cryogenic Engines for Space Exploration

NASA Technical Reports Server (NTRS)

Brown, Kendall K.; Nelson, Karl W.

2005-01-01

Historically, cryogenic rocket engines have not been used for in-space applications due to their additional complexity, the mission need for high reliability, and the challenges of propellant boil-off. While the mission and vehicle architectures are not yet defined for the lunar and Martian robotic and human exploration objectives, cryogenic rocket engines offer the potential for higher performance and greater architecture/mission flexibility. In-situ cryogenic propellant production could enable a more robust exploration program by significantly reducing the propellant mass delivered to low earth orbit, thus warranting the evaluation of cryogenic rocket engines versus the hypergolic bi-propellant engines used in the Apollo program. A multi-use engine. one which can provide the functionality that separate engines provided in the Apollo mission architecture, is desirable for lunar and Mars exploration missions because it increases overall architecture effectiveness through commonality and modularity. The engine requirement derivation process must address each unique mission application and each unique phase within each mission. The resulting requirements, such as thrust level, performance, packaging, bum duration, number of operations; required impulses for each trajectory phase; operation after extended space or surface exposure; availability for inspection and maintenance; throttle range for planetary descent, ascent, acceleration limits and many more must be addressed. Within engine system studies, the system and component technology, capability, and risks must be evaluated and a balance between the appropriate amount of technology-push and technology-pull must be addressed. This paper will summarize many of the key technology challenges associated with using high-performance cryogenic liquid propellant rocket engine systems and components in the exploration program architectures. The paper is divided into two areas. The first area describes how the mission requirements affect the engine system requirements and create system level technology challenges. An engine system architecture for multiple applications or a family of engines based upon a set of core technologies, design, and fabrication approaches may reduce overall programmatic cost and risk. The engine system discussion will also address the characterization of engine cycle figures of merit, configurations, and design approaches for some in-space vehicle alternatives under consideration. The second area evaluates the component-level technology challenges induced from the system requirements. Component technology issues are discussed addressing injector, thrust chamber, ignition system, turbopump assembly, and valve design for the challenging requirements of high reliability, robustness, fault tolerance, deep throttling, reasonable performance (with respect to weight and specific impulse).

Gauging Systems Monitor Cryogenic Liquids

NASA Technical Reports Server (NTRS)

2009-01-01

Rocket fuel needs to stay cool - super cool, in fact. The ability to store gas propellants like liquid hydrogen and oxygen at cryogenic temperatures (below -243 F) is crucial for space missions in order to reduce their volumes and allow their storage in smaller (and therefore, less costly) tanks. The Agency has used these cryogenic fluids for vehicle propellants, reactants, and life support systems since 1962 with the Centaur upper stage rocket, which was powered with liquid oxygen and liquid hydrogen. During proposed long-duration missions, super-cooled fluids will also be used in space power systems, spaceports, and lunar habitation systems. In the next generation of launch vehicles, gaseous propellants will be cooled to and stored for extended periods at even colder temperatures than currently employed via a process called densification. Densification sub-cools liquids to temperatures even closer to absolute zero (-459 F), increasing the fluid s density and shrinking its volume beyond common cryogenics. Sub-cooling cryogenic liquid hydrogen, for instance, from 20 K (-423 F) to 15 K (-432.4 F) reduces its mass by 10 percent. These densified liquid gases can provide more cost savings from reduced payload volume. In order to benefit from this cost savings, the Agency is working with private industry to prevent evaporation, leakage, and other inadvertent loss of liquids and gases in payloads - requiring new cryogenic systems to prevent 98 percent (or more) of boil-off loss. Boil-off occurs when cryogenic or densified liquids evaporate, and is a concern during launch pad holds. Accurate sensing of propellants aboard space vehicles is also critical for proper engine shutdown and re-ignition after launch, and zero boil-off fuel systems are also in development for the Altair lunar lander.

The E-3 Test Facility at Stennis Space Center: Research and Development Testing for Cryogenic and Storable Propellant Combustion Systems

NASA Technical Reports Server (NTRS)

Pazos, John T.; Chandler, Craig A.; Raines, Nickey G.

2009-01-01

This paper will provide the reader a broad overview of the current upgraded capabilities of NASA's John C. Stennis Space Center E-3 Test Facility to perform testing for rocket engine combustion systems and components using liquid and gaseous oxygen, gaseous and liquid methane, gaseous hydrogen, hydrocarbon based fuels, hydrogen peroxide, high pressure water and various inert fluids. Details of propellant system capabilities will be highlighted as well as their application to recent test programs and accomplishments. Data acquisition and control, test monitoring, systems engineering and test processes will be discussed as part of the total capability of E-3 to provide affordable alternatives for subscale to full scale testing for many different requirements in the propulsion community.

Test Stand at the Rocket Engine Test Facility

NASA Image and Video Library

1973-02-21

The thrust stand in the Rocket Engine Test Facility at the National Aeronautics and Space Administration (NASA) Lewis Research Center in Cleveland, Ohio. The Rocket Engine Test Facility was constructed in the mid-1950s to expand upon the smaller test cells built a decade before at the Rocket Laboratory. The $2.5-million Rocket Engine Test Facility could test larger hydrogen-fluorine and hydrogen-oxygen rocket thrust chambers with thrust levels up to 20,000 pounds. Test Stand A, seen in this photograph, was designed to fire vertically mounted rocket engines downward. The exhaust passed through an exhaust gas scrubber and muffler before being vented into the atmosphere. Lewis researchers in the early 1970s used the Rocket Engine Test Facility to perform basic research that could be utilized by designers of the Space Shuttle Main Engines. A new electronic ignition system and timer were installed at the facility for these tests. Lewis researchers demonstrated the benefits of ceramic thermal coatings for the engine’s thrust chamber and determined the optimal composite material for the coatings. They compared the thermal-coated thrust chamber to traditional unlined high-temperature thrust chambers. There were more than 17,000 different configurations tested on this stand between 1973 and 1976. The Rocket Engine Test Facility was later designated a National Historic Landmark for its role in the development of liquid hydrogen as a propellant.

Program user's manual for optimizing the design of a liquid or gaseous propellant rocket engine with the automated combustor design code AUTOCOM

NASA Technical Reports Server (NTRS)

Reichel, R. H.; Hague, D. S.; Jones, R. T.; Glatt, C. R.

1973-01-01

This computer program manual describes in two parts the automated combustor design optimization code AUTOCOM. The program code is written in the FORTRAN 4 language. The input data setup and the program outputs are described, and a sample engine case is discussed. The program structure and programming techniques are also described, along with AUTOCOM program analysis.

Two-step rocket engine bipropellant valve concept

NASA Technical Reports Server (NTRS)

Capps, J. E.; Ferguson, R. E.; Pohl, H. O.

1969-01-01

Initiating combustion of altitude control rocket engines in a precombustion chamber of ductile material reduces high pressure surges generated by hypergolic propellants. Two-step bipropellant valve concepts control initial propellant flow into precombustion chamber and subsequent full flow into main chamber.

Parallelization of Rocket Engine System Software (Press)

NASA Technical Reports Server (NTRS)

Cezzar, Ruknet

1996-01-01

The main goal is to assess parallelization requirements for the Rocket Engine Numeric Simulator (RENS) project which, aside from gathering information on liquid-propelled rocket engines and setting forth requirements, involve a large FORTRAN based package at NASA Lewis Research Center and TDK software developed by SUBR/UWF. The ultimate aim is to develop, test, integrate, and suitably deploy a family of software packages on various aspects and facets of rocket engines using liquid-propellants. At present, all project efforts by the funding agency, NASA Lewis Research Center, and the HBCU participants are disseminated over the internet using world wide web home pages. Considering obviously expensive methods of actual field trails, the benefits of software simulators are potentially enormous. When realized, these benefits will be analogous to those provided by numerous CAD/CAM packages and flight-training simulators. According to the overall task assignments, Hampton University's role is to collect all available software, place them in a common format, assess and evaluate, define interfaces, and provide integration. Most importantly, the HU's mission is to see to it that the real-time performance is assured. This involves source code translations, porting, and distribution. The porting will be done in two phases: First, place all software on Cray XMP platform using FORTRAN. After testing and evaluation on the Cray X-MP, the code will be translated to C + + and ported to the parallel nCUBE platform. At present, we are evaluating another option of distributed processing over local area networks using Sun NFS, Ethernet, TCP/IP. Considering the heterogeneous nature of the present software (e.g., first started as an expert system using LISP machines) which now involve FORTRAN code, the effort is expected to be quite challenging.

Marshall Space Flight Center Autumn 2005

NASA Technical Reports Server (NTRS)

Allen, Mike; Clar, Harry E.

2006-01-01

The East Test Area at Marshall Space Flight Center has five major test stands, each of which has two or more test positions, not counting the SSME and RD-180 engine test facilities in the West Test Area. These research and development facilities are capable of testing high pressure pumps, both fuel and oxidizer, injectors, chambers and sea-level engine assemblies, as well as simulating deep space environments in the 12, 15 and 20 foot vacuum chambers. Liquid propellant capabilities are high pressure hydrogen (liquid and gas), methane (liquid and gas), and RP-1 and high pressure LOX. Solid propellant capability includes thrust measurement and firing capability up to 1/6 scale Shuttle SRB segment. In the past six months MSFC supported multiple space access and exploration programs in the previous six months. Major programs were Space Exploration, Shuttle External Tank research, Reusable Solid Rocket Motor (RSRM) development, as well as research programs for NASA and other customers. At Test Stand 115 monopropellant ignition testing was conducted on one position. At the second position multiple ignition/variable burn time cycles were conducted on Vacuum Plasma Spatter (VPS) coated injectors. Each injector received fifty cycles; the propellants were LOX Hydrogen and the ignition source was TEA. Following completion of the monopropellant test series the stand was reconfigured to support ignition testing on a LOX Methane injector system. At TS 116 a thrust stand used to test Booster Separation Motors from the Shuttle SRB system was disassembled and moved from Chemical Systems Division s Coyote Canyon plant to MSFC. The stand was reassembled and readied for BSM testing. Also, a series of tests was run on a Pratt & Whitney Rocketdyne Low Element Density (LED) injector engine. The propellants for this engine are LOX and LH2. At TS 300 the 20 foot vacuum chamber was configured to support hydrogen testing in the Multipurpose Hydrogen Test Bed (MHTB) test article. This testing, which went 24/7 for fourteen consecutive days, demonstrated long duration storage methods intended to minimize losses of propellant in support of the Space Exploration Initiative. The facility is being converted to support similar research using liquid methane. The 12 foot chamber at TS 300 was used to create ascent profiles (both heat and altitude effects) for foam panel testing in support of the Shuttle External Tank program. At TS 500, one position was in build-up to support ATK Thiokol research into the gas dynamics associated with high pressure flow across the propellant joint in segmented solid rocket motors. The testing involves flowing high pressure gas through a 24 motor case. Initial tests will be conducted with simulated aluminum grain, followed by tests using actual propellant. The second position at TS 500 has been in build-up for testing a LOX methane thruster manufactured by KT Engineering. At the Solid Propulsion Test Area (SPTA), the first dual segment 24 solid rocket motor was fired for ATK Thiokol in support of the RSRM program. A new axial thrust measurement stand was designed and fabricated for this testing. Real Time Radiography (RTR) will be deployed to examine nozzle erosion on the next dual segment motor.

Scaled Rocket Testing in Hypersonic Flow

NASA Technical Reports Server (NTRS)

Dufrene, Aaron; MacLean, Matthew; Carr, Zakary; Parker, Ron; Holden, Michael; Mehta, Manish

2015-01-01

NASA's Space Launch System (SLS) uses four clustered liquid rocket engines along with two solid rocket boosters. The interaction between all six rocket exhaust plumes will produce a complex and severe thermal environment in the base of the vehicle. This work focuses on a recent 2% scale, hot-fire SLS base heating test. These base heating tests are short-duration tests executed with chamber pressures near the full-scale values with gaseous hydrogen/oxygen engines and RSRMV analogous solid propellant motors. The LENS II shock tunnel/Ludwieg tube tunnel was used at or near flight duplicated conditions up to Mach 5. Model development was strongly based on the Space Shuttle base heating tests with several improvements including doubling of the maximum chamber pressures and duplication of freestream conditions. Detailed base heating results are outside of the scope of the current work, rather test methodology and techniques are presented along with broader applicability toward scaled rocket testing in supersonic and hypersonic flow.

Marshall Team Recreates Goddard Rocket

NASA Technical Reports Server (NTRS)

2003-01-01

In honor of the Centernial of Flight celebration and commissioned by the American Institute of Aeronautics and Astronautics (AIAA), a team of engineers from Marshall Space Flight Center (MSFC) built a replica of the first liquid-fueled rocket. The original rocket, designed and built by rocket engineering pioneer Robert H. Goddard in 1926, opened the door to modern rocketry. Goddard's rocket reached an altitude of 41 feet while its flight lasted only 2.5 seconds. The Marshall design team's plan was to stay as close as possible to an authentic reconstruction of Goddard's rocket. The same propellants were used - liquid oxygen and gasoline - as available during Goddard's initial testing and firing. The team also tried to construct the replica using the original materials and design to the greatest extent possible. By purposely using less advanced techniques and materials than many that are available today, the team encountered numerous technical challenges in testing the functional hardware. There were no original blueprints or drawings, only photographs and notes. However, this faithful adherence to historical accuracy has also allowed the team to experience many of the same challenges Goddard faced 77 years ago, and more fully appreciate the genius of this extraordinary man. The replica will undergo ground tests at MSFC this summer.

An air-breathing ballistic space transporter for Europe

NASA Technical Reports Server (NTRS)

Kramer, P. A.; Buehler, R. D.

1985-01-01

With increasing transport requirements, reusable space transporters again receive serious consideration in Europe as successors to the Ariane family. The paper deals with a hydrogen-ramjet-propelled, 1-1/2-stage reusable ballistic space transporter with vertical take-off and landing and using liquid hydrogen/oxygen rockets. This novel concept was developed in a theoretical study at the University of Stuttgart. The results are compared with recently published studies of several other European space transporter concepts. The data derived for the Istra - concept are: 15.4 Mg payload into low Earth-orbit, 155 Mg gross lift-off mass, 10% payload ratio, which represents a 57% propellant saving, and 44% reduction in dry mass (structure and engines) compared with comparable two-stage pure rocket concepts.

Development of the platelet micro-orifice injector. [for liquid propellant rocket engines

NASA Technical Reports Server (NTRS)

La Botz, R. J.

1984-01-01

For some time to come, liquid rocket engines will continue to provide the primary means of propulsion for space transportation. The injector represents a key to the optimization of engine and system performance. The present investigation is concerned with a unique injector design and fabrication process which has demonstrated performance capabilities beyond that achieved with more conventional approaches. This process, which is called the 'platelet process', makes it feasible to fabricate injectors with a pattern an order of magnitude finer than that obtainable by drilling. The fine pattern leads to an achievement of high combustion efficiencies. Platelet injectors have been identified as one of the significant technology advances contributing to the feasibility of advanced dual-fuel booster engines. Platelet injectors are employed in the Space Shuttle Orbit Maneuvering System (OMS) engines. Attention is given to injector design theory as it relates to pattern fineness, a description of platelet injectors, and test data obtained with three different platelet injectors.

Ignition of Hydrogen-Oxygen Rocket Combustor with Chlorine Trifluoride and Triethylaluminum

NASA Technical Reports Server (NTRS)

Gregory, John W.; Straight, David M.

1961-01-01

Ignition of a nominal-125-pound-thrust cold (2000 R) gaseous-hydrogen - liquid-oxygen rocket combustor with chlorine trifluoride (hypergolic with hydrogen) and triethylaluminum (hypergolic with oxygen) resulted in consistently smooth starting transients for a wide range of combustor operating conditions. The combustor exhaust nozzle discharged into air at ambient conditions. Each starting transient consisted of the following sequence of events: injection of the lead main propellant, injection of the igniter chemical, ignition of these two chemicals, injection of the second main propellant, ignition of the two main propellants, increase in chamber pressure to its terminal value, and cutoff of igniter-chemical flow. Smooth ignition was obtained with an ignition delay of less than 100 milliseconds for the reaction of the lead propellant with the igniter chemical using approximately 0.5 cubic inch (0-038 lb) of chlorine trifluoride or 1.0 cubic inch (0-031 lb) of triethylaluminum. These quantities of igniter chemical were sufficient to ignite a 20-percent-fuel hydrogen-oxygen mixture with a delay time of less than 15 milliseconds. Test results indicated that a simple, light weight chemical ignition system for hydrogen-oxygen rocket engines may be possible.

Performance of a transpiration-regenerative cooled rocket thrust chamber

NASA Technical Reports Server (NTRS)

Valler, H. W.

1979-01-01

The analysis, design, fabrication, and testing of a liquid rocket engine thrust chamber which is gas transpiration cooled in the high heat flux convergent portion of the chamber and water jacket cooled (simulated regenerative) in the barrel and divergent sections of the chamber are described. The engine burns LOX-hydrogen propellants at a chamber pressure of 600 psia. Various transpiration coolant flow rates were tested with resultant local hot gas wall temperatures in the 800 F to 1400 F range. The feasibility of transpiration cooling with hydrogen and helium, and the use of photo-etched copper platelets for heat transfer and coolant metering was successfully demonstrated.

Feasibility of rocket propellant production on Mars

NASA Technical Reports Server (NTRS)

Ash, R. L.; Dowler, W. L.; Varsi, G.

1978-01-01

In situ production of rocket propellant to reduce landed mass requirements for Mars return missions has been investigated. The analysis has shown that a system which utilizes atmospheric carbon dioxide and soil moisture to produce liquid methane-oxygen propellant requires a landed mass which is less than half the mass of the ascent vehicle it produces.

VIABILITY OF BACILLUS SUBTILIS SPORES IN ROCKET PROPELLANTS.

PubMed

GODDING, R M; LYNCH, V H

1965-01-01

The sporicidal activity of components used in liquid and solid rocket propellants was tested by use of spores of Bacillus subtilis dried on powdered glass. Liquid propellant ingredients tested were N(2)O(4), monomethylhydrazine and 1,1-dimethylhydrazine. N(2)O(4) was immediately sporicidal; the hydrazines were effective within several days. Solid propellants consisted of ammonium perchlorate in combination with epoxy resin (EPON 828), tris-1-(2-methyl) aziridinyl phosphine oxide, bis-1-(2-methyl) aziridinyl phenylphosphine oxide, and three modified polybutadiene polymers. There was no indication of appreciable sporicidal activity of these components.

Viability of Bacillus subtilis Spores in Rocket Propellants

PubMed Central

Godding, Rogene M.; Lynch, Victoria H.

1965-01-01

The sporicidal activity of components used in liquid and solid rocket propellants was tested by use of spores of Bacillus subtilis dried on powdered glass. Liquid propellant ingredients tested were N2O4, monomethylhydrazine and 1,1-dimethylhydrazine. N2O4 was immediately sporicidal; the hydrazines were effective within several days. Solid propellants consisted of ammonium perchlorate in combination with epoxy resin (EPON 828), tris-1-(2-methyl) aziridinyl phosphine oxide, bis-1-(2-methyl) aziridinyl phenylphosphine oxide, and three modified polybutadiene polymers. There was no indication of appreciable sporicidal activity of these components. PMID:14264838

Small Liquid Hydrogen Tank for Drop Tower Tests

NASA Image and Video Library

1964-11-21

A researcher fills a small container used to represent a liquid hydrogen tank in preparation for a microgravity test in the 2.2-Second Drop Tower at the National Aeronautics and Space Administration (NASA) Lewis Research Center. For over a decade, NASA Lewis endeavored to make liquid hydrogen a viable propellant. Hydrogen’s light weight and high energy made it very appealing for rocket propulsion. One of the unknowns at the time was the behavior of fluids in the microgravity of space. Rocket designers needed to know where the propellant would be inside the fuel tank in order to pump it to the engine. NASA Lewis utilized sounding rockets, research aircraft, and the 2.2 Second Drop Tower to study liquids in microgravity. The drop tower, originally built as a fuel distillation tower in 1948, descended into a steep ravine. By early 1961 the facility was converted into an eight-floor, 100-foot tower connected to a shop and laboratory space. Small glass tanks, like this one, were installed in experiment carts with cameras to film the liquid’s behavior during freefall. Thousands of drop tower tests in the early 1960s provided an increased understanding of low-gravity processes and phenomena. The tower only afforded a relatively short experiment time but was sufficient enough that the research could be expanded upon using longer duration freefalls on sounding rockets or aircraft. The results of the early experimental fluid studies verified predictions made by Lewis researchers that the total surface energy would be minimized in microgravity.

Coal-Fired Rocket Engine

NASA Technical Reports Server (NTRS)

Anderson, Floyd A.

1987-01-01

Brief report describes concept for coal-burning hybrid rocket engine. Proposed engine carries larger payload, burns more cleanly, and safer to manufacture and handle than conventional solid-propellant rockets. Thrust changeable in flight, and stops and starts on demand.

A Design Tool for Liquid Rocket Engine Injectors

NASA Technical Reports Server (NTRS)

Farmer, R.; Cheng, G.; Trinh, H.; Tucker, K.

2000-01-01

A practical design tool which emphasizes the analysis of flowfields near the injector face of liquid rocket engines has been developed and used to simulate preliminary configurations of NASA's Fastrac and vortex engines. This computational design tool is sufficiently detailed to predict the interactive effects of injector element impingement angles and points and the momenta of the individual orifice flows and the combusting flow which results. In order to simulate a significant number of individual orifices, a homogeneous computational fluid dynamics model was developed. To describe sub- and supercritical liquid and vapor flows, the model utilized thermal and caloric equations of state which were valid over a wide range of pressures and temperatures. The model was constructed such that the local quality of the flow was determined directly. Since both the Fastrac and vortex engines utilize RP-1/LOX propellants, a simplified hydrocarbon combustion model was devised in order to accomplish three-dimensional, multiphase flow simulations. Such a model does not identify drops or their distribution, but it does allow the recirculating flow along the injector face and into the acoustic cavity and the film coolant flow to be accurately predicted.

Analysis of a Radioisotope Thermal Rocket Engine

NASA Technical Reports Server (NTRS)

Machado-Rodriguez, Jonathan P.; Landis, Geoffrey A.

2017-01-01

The Triton Hopper is a concept for a vehicle to explore the surface of Neptunes moon Triton, which uses a radioisotope heated rocket engine and in-situ propellant acquisition. The initial Triton Hopper conceptual design stores pressurized Nitrogen in a spherical tank to be used as the propellant. The aim of the research was to investigate the benefits of storing propellant at ambient temperature and heating it through a thermal block during engine operation, as opposed to storing gas at a high temperature.

Ozone Depletion Caused by Rocket Engine Emissions: A Fundamental Limit on the Scale and Viability of Space-Based Geoengineering Schemes

NASA Astrophysics Data System (ADS)

Ross, M. N.; Toohey, D.

2008-12-01

Emissions from solid and liquid propellant rocket engines reduce global stratospheric ozone levels. Currently ~ one kiloton of payloads are launched into earth orbit annually by the global space industry. Stratospheric ozone depletion from present day launches is a small fraction of the ~ 4% globally averaged ozone loss caused by halogen gases. Thus rocket engine emissions are currently considered a minor, if poorly understood, contributor to ozone depletion. Proposed space-based geoengineering projects designed to mitigate climate change would require order of magnitude increases in the amount of material launched into earth orbit. The increased launches would result in comparable increases in the global ozone depletion caused by rocket emissions. We estimate global ozone loss caused by three space-based geoengineering proposals to mitigate climate change: (1) mirrors, (2) sunshade, and (3) space-based solar power (SSP). The SSP concept does not directly engineer climate, but is touted as a mitigation strategy in that SSP would reduce CO2 emissions. We show that launching the mirrors or sunshade would cause global ozone loss between 2% and 20%. Ozone loss associated with an economically viable SSP system would be at least 0.4% and possibly as large as 3%. It is not clear which, if any, of these levels of ozone loss would be acceptable under the Montreal Protocol. The large uncertainties are mainly caused by a lack of data or validated models regarding liquid propellant rocket engine emissions. Our results offer four main conclusions. (1) The viability of space-based geoengineering schemes could well be undermined by the relatively large ozone depletion that would be caused by the required rocket launches. (2) Analysis of space- based geoengineering schemes should include the difficult tradeoff between the gain of long-term (~ decades) climate control and the loss of short-term (~ years) deep ozone loss. (3) The trade can be properly evaluated only if our understanding of the stratospheric impact of rocket emissions is significantly improved. (4) Such an improved understanding requires a concerted effort of research including new in situ measurements in a variety of rocket plumes and a multi-scale modeling program similar in scope to the effort required to address the climate and ozone impacts of aircraft emissions.

Space and Missile Systems Center Standard: Test Requirements for Launch, Upper-Stage and Space Vehicles

DTIC Science & Technology

2014-09-05

adiabatic expansion of a perfect gas ; b. Contains a gas or liquid that would endanger personnel or equipment or create a mis- hap if released; or c...Guidelines for Liquid Rocket Engines 31. TOR-2013(3213)-6 Acoustic Testing on Production Space Vehicle (The Value of the Test and Deletion...materials used in space vehicles, interstages, payload adapters, payload fairings, motor cases, nozzles , propellant tanks, and over-wrapped pressure vessels

AXISYMMETRIC, THROTTLEABLE NON-GIMBALLED ROCKET ENGINE

NASA Technical Reports Server (NTRS)

Sackheim, Robert L. (Inventor); Hutt, John J. (Inventor); Anderson, William E. (Inventor); Dressler, Gordon A. (Inventor)

2005-01-01

A rocket engine assembly is provided for a vertically launched rocket vehicle. A rocket engine housing of the assembly includes two or more combustion chambers each including an outlet end defining a sonic throat area. A propellant supply for the combustion chambers includes a throttling injector, associated with each of the combustion chambers and located opposite to sonic throat area, which injects the propellant into the associated combustion chamber. A modulator, which may form part of the injector, and which is controlled by a controller, modulates the flow rate of the propellant to the combustion chambers so that the chambers provide a vectorable net thrust. An expansion nozzle or body located downstream of the throat area provides expansion of the combustion gases produced by the combustion chambers so as to increase the net thrust.

Analysis and Evaluation of German Attainments and Research in the Liquid Rocket Engine Field. Volume 4. Propellant Injectors

DTIC Science & Technology

1951-02-01

they were ob- served at a given pressure drop in "cold" testing with water or unreacted propellants. heat-transfer considerations and the location of... water as a coolant in the main chamber. The Winkler injector was used.on a test unit developing a thrust of 220 lb and an exhaust ve- locity of 6370 ft... water . Provision was made for an igniter in the center of the injector. The relatively high performance reported for this unit does not seem to be

Combustion stability with baffles, absorbers and velocity sensitive combustion. [liquid propellant rocket combustors

NASA Technical Reports Server (NTRS)

Mitchell, C. E.

1980-01-01

Analytical and computational techniques were developed to predict the stability behavior of liquid propellant rocket combustors using damping devices such as acoustic liners, slot absorbers, and injector face baffles. Models were developed to determine the frequency and decay rate of combustor oscillations, the spatial and temporal pressure waveforms, and the stability limits in terms of combustion response model parameters.

Conceptual Launch Vehicles Using Metallic Hydrogen Propellant

NASA Astrophysics Data System (ADS)

Cole, John W.; Silvera, Isaac F.; Foote, John P.

2008-01-01

Solid molecular hydrogen is predicted to transform into an atomic solid with metallic properties under pressures >4.5 Mbar. Atomic metallic hydrogen is predicted to be metastable, limited by some critical temperature and pressure, and to store very large amounts of energy. Experiments may soon determine the critical temperature, critical pressure, and specific energy availability. It is useful to consider the feasibility of using metastable atomic hydrogen as a rocket propellant. If one assumes that metallic hydrogen is stable at usable temperatures and pressures, and that it can be affordably produced, handled, and stored, then it may be a useful rocket propellant. Assuming further that the available specific energy can be determined from the recombination of the atoms into molecules (216 MJ/kg), then conceptual engines and launch vehicle concepts can be developed. Under these assumptions, metallic hydrogen would be a revolutionary new rocket fuel with a theoretical specific impulse of 1700 s at a chamber pressure of 100 atm. A practical problem that arises is that rocket chamber temperatures may be too high for the use of this pure fuel. This paper examines an engine concept that uses liquid hydrogen or water as a diluent coolant for the metallic hydrogen to reduce the chamber temperature to usable values. Several launch vehicles are then conceptually developed. Results indicate that if metallic hydrogen is experimentally found to have the properties assumed in this analysis, then there are significant benefits. These benefits become more attractive as the chamber temperatures increase.

Around Marshall

NASA Image and Video Library

2003-07-01

In honor of the Centernial of Flight celebration and commissioned by the American Institute of Aeronautics and Astronautics (AIAA), a team of engineers from Marshall Space Flight Center (MSFC) built a replica of the first liquid-fueled rocket. The original rocket, designed and built by rocket engineering pioneer Robert H. Goddard in 1926, opened the door to modern rocketry. Goddard's rocket reached an altitude of 41 feet while its flight lasted only 2.5 seconds. The Marshall design team's plan was to stay as close as possible to an authentic reconstruction of Goddard's rocket. The same propellants were used - liquid oxygen and gasoline - as available during Goddard's initial testing and firing. The team also tried to construct the replica using the original materials and design to the greatest extent possible. By purposely using less advanced techniques and materials than many that are available today, the team encountered numerous technical challenges in testing the functional hardware. There were no original blueprints or drawings, only photographs and notes. However, this faithful adherence to historical accuracy has also allowed the team to experience many of the same challenges Goddard faced 77 years ago, and more fully appreciate the genius of this extraordinary man. The replica will undergo ground tests at MSFC this summer.

Investigating the effects of critical phenomena in premixed methane-oxygen flames at cryogenic conditions

NASA Astrophysics Data System (ADS)

Gopal, Abishek; Yellapantula, Shashank; Larsson, Johan

2017-11-01

Methane is increasingly becoming viable as a rocket fuel in the latest generation of launch vehicles. In liquid rocket engines, fuel and oxidizer are injected under cryogenic conditions into the combustion chamber. At high pressures, typical of rocket combustion chambers, the propellants exist in supercritical states where the ideal gas thermodynamics are no longer valid. We investigate the effects of real-gas thermodynamics on transcritical laminar premixed methane-oxygen flames. The effect of the real-gas cubic equations of state and high-pressure transport properties on flame dynamics is presented. We also study real-gas effects on the extinction limits of the methane-oxygen flame.

Real-Time Simulation of the X-33 Aerospace Engine

NASA Technical Reports Server (NTRS)

Aguilar, Robert

1999-01-01

This paper discusses the development and performance of the X-33 Aerospike Engine RealTime Model. This model was developed for the purposes of control law development, six degree-of-freedom trajectory analysis, vehicle system integration testing, and hardware-in-the loop controller verification. The Real-Time Model uses time-step marching solution of non-linear differential equations representing the physical processes involved in the operation of a liquid propellant rocket engine, albeit in a simplified form. These processes include heat transfer, fluid dynamics, combustion, and turbomachine performance. Two engine models are typically employed in order to accurately model maneuvering and the powerpack-out condition where the power section of one engine is used to supply propellants to both engines if one engine malfunctions. The X-33 Real-Time Model is compared to actual hot fire test data and is been found to be in good agreement.

Flight Testing the Linear Aerospike SR-71 Experiment (LASRE)

NASA Technical Reports Server (NTRS)

Corda, Stephen; Neal, Bradford A.; Moes, Timothy R.; Cox, Timothy H.; Monaghan, Richard C.; Voelker, Leonard S.; Corpening, Griffin P.; Larson, Richard R.; Powers, Bruce G.

1998-01-01

The design of the next generation of space access vehicles has led to a unique flight test that blends the space and flight research worlds. The new space vehicle designs, such as the X-33 vehicle and Reusable Launch Vehicle (RLV), are powered by linear aerospike rocket engines. Conceived of in the 1960's, these aerospike engines have yet to be flown, and many questions remain regarding aerospike engine performance and efficiency in flight. To provide some of these data before flying on the X-33 vehicle and the RLV, a spacecraft rocket engine has been flight-tested atop the NASA SR-71 aircraft as the Linear Aerospike SR-71 Experiment (LASRE). A 20 percent-scale, semispan model of the X-33 vehicle, the aerospike engine, and all the required fuel and oxidizer tanks and propellant feed systems have been mounted atop the SR-71 airplane for this experiment. A major technical objective of the LASRE flight test is to obtain installed-engine performance flight data for comparison to wind-tunnel results and for the development of computational fluid dynamics-based design methodologies. The ultimate goal of firing the aerospike rocket engine in flight is still forthcoming. An extensive design and development phase of the experiment hardware has been completed, including approximately 40 ground tests. Five flights of the LASRE and firing the rocket engine using inert liquid nitrogen and helium in place of liquid oxygen and hydrogen have been successfully completed.

Rocket Research Presentation at the NACA's 1947 Inspection

NASA Image and Video Library

1947-10-21

Researcher John Sloop briefs visitors on his latest rocket engine research during the 1947 Inspection at the National Advisory Committee for Aeronautics (NACA) Lewis Flight Propulsion Laboratory. The NACA had been hosting annual Aircraft Engineering Conferences, better known as Inspections, since 1926. Individuals from the manufacturing industry, military, and university settings were invited to tour the NACA laboratories. There were a series of stops on the tour, mostly at test facilities, where researchers would brief the group on the latest efforts in their particular field. The Inspections grew in size and scope over the years and by the mid-1940s required multiple days. The three-day 1947 Inspection was the first time the event was held at NACA Lewis. Over 800 scientists, industrialists, and military leaders attended the three-day event. Talks were given at the Altitude Wind Tunnel, Four Burner Area, Engine Research Building, and other facilities. An array of topics were discussed, including full-scale engine testing, ramjets, axial-flow compressors, turbojets, fuels, icing, and materials. The NACA Lewis staff and their families were able to view the same presentations after the Inspection was over. Sloop, a researcher in the Fuels and Thermodynamics Division, briefed visitors on NACA Lewis’ early research in rocket engine propellants, combustion, and cooling. This early NACA Lewis work led to the development of liquid hydrogen as a viable propellant in the late 1950s.

Saturn Apollo Program

NASA Image and Video Library

1965-07-10

Marshall Space Flight Center's rocket development has always included component testing. Pictured here is a Cell 114-B burn stack. The C114-B is part of the gas generators used to test heat exchanges for the F-1 engine. On the initial firing of the C114-B the spark ignition would not light. The rocket propellant mixed with the liquid oxygen gelled creating a bomb. After several attempts at ignition, the spark ignited and blew up the stand. Subsequent testings were completed on newly constructed stands and no further mishaps were reported.

Uncertainty Quantification of Non-linear Oscillation Triggering in a Multi-injector Liquid-propellant Rocket Combustion Chamber

NASA Astrophysics Data System (ADS)

Popov, Pavel; Sideris, Athanasios; Sirignano, William

2014-11-01

We examine the non-linear dynamics of the transverse modes of combustion-driven acoustic instability in a liquid-propellant rocket engine. Triggering can occur, whereby small perturbations from mean conditions decay, while larger disturbances grow to a limit-cycle of amplitude that may compare to the mean pressure. For a deterministic perturbation, the system is also deterministic, computed by coupled finite-volume solvers at low computational cost for a single realization. The randomness of the triggering disturbance is captured by treating the injector flow rates, local pressure disturbances, and sudden acceleration of the entire combustion chamber as random variables. The combustor chamber with its many sub-fields resulting from many injector ports may be viewed as a multi-scale complex system wherein the developing acoustic oscillation is the emergent structure. Numerical simulation of the resulting stochastic PDE system is performed using the polynomial chaos expansion method. The overall probability of unstable growth is assessed in different regions of the parameter space. We address, in particular, the seven-injector, rectangular Purdue University experimental combustion chamber. In addition to the novel geometry, new features include disturbances caused by engine acceleration and unsteady thruster nozzle flow.

Technology Challenges for Deep-Throttle Cryogenic Engines for Space Exploration

NASA Astrophysics Data System (ADS)

Brown, Kendall K.; Nelson, Karl W.

2005-02-01

Historically, cryogenic rocket engines have not been used for in-space applications due to their additional complexity, the mission need for high reliability, and the challenges of propellant boil-off. While the mission and vehicle architectures are not yet defined for the lunar and Martian robotic and human exploration objectives, cryogenic rocket engines offer the potential for higher performance and greater architecture/mission flexibility. In-situ cryogenic propellant production could enable a more robust exploration program by significantly reducing the propellant mass delivered to low earth orbit, thus warranting the evaluation of cryogenic rocket engines versus the hypergolic bipropellant engines used in the Apollo program. A multi-use engine, one which can provide the functionality that separate engines provided in the Apollo mission architecture, is desirable for lunar and Mars exploration missions because it increases overall architecture effectiveness through commonality and modularity. The engine requirement derivation process must address each unique mission application and each unique phase within each mission. The resulting requirements, such as thrust level, performance, packaging, burn duration, number of operations; required impulses for each trajectory phase; operation after extended space or surface exposure; availability for inspection and maintenance; throttle range for planetary descent, ascent, acceleration limits and many more must be addressed. Within engine system studies, the system and component technology, capability, and risks must be evaluated and a balance between the appropriate amount of technology-push and technology-pull must be addressed. This paper will summarize many of the key technology challenges associated with using high-performance cryogenic liquid propellant rocket engine systems and components in the exploration program architectures. The paper is divided into two areas. The first area describes how the mission requirements affect the engine system requirements and create system level technology challenges. An engine system architecture for multiple applications or a family of engines based upon a set of core technologies, design, and fabrication approaches may reduce overall programmatic cost and risk. The engine system discussion will also address the characterization of engine cycle figures of merit, configurations, and design approaches for some in-space vehicle alternatives under consideration. The second area evaluates the component-level technology challenges induced from the system requirements. Component technology issues are discussed addressing injector, thrust chamber, ignition system, turbopump assembly, and valve design for the challenging requirements of high reliability, robustness, fault tolerance, deep throttling, reasonable performance (with respect to weight and specific impulse).

Extension of a simplified computer program for analysis of solid-propellant rocket motors

NASA Technical Reports Server (NTRS)

Sforzini, R. H.

1973-01-01

A research project to develop a computer program for the preliminary design and performance analysis of solid propellant rocket engines is discussed. The following capabilities are included as computer program options: (1) treatment of wagon wheel cross sectional propellant configurations alone or in combination with circular perforated grains, (2) calculation of ignition transients with the igniter treated as a small rocket engine, (3) representation of spherical circular perforated grain ends as an alternative to the conical end surface approximation used in the original program, and (4) graphical presentation of program results using a digital plotter.

Turbo Pump Fed Micro-Rocket Engine

NASA Astrophysics Data System (ADS)

Miotti, P.; Tajmar, M.; Seco, F.; Guraya, C.; Perennes, F.; Soldati, A.; Lang, M.

2004-10-01

Micro-satellites (from 10kg up to 100kg) have mass, volume, and electrical power constraints due to their low dimensions. These limitations lead to the lack in currently available active orbit control systems in micro-satellites. Therefore, a micro-propulsion system with a high thrust to mass ratio is required to increase the potential functionality of small satellites. Mechatronic is presently working on a liquid bipropellant micro-rocket engine under contract with ESA (Contract No.16914/NL/Sfe - Micro-turbo-machinery Based Bipropellant System Using MNT). The advances in Mechatronic's project are to realise a micro-rocket engine with propellants pressurised by micro-pumps. The energy for driving the pumps would be extracted from a micro-turbine. Cooling channels around the nozzle would be also used in order to maintain the wall material below its maximum operating temperature. A mass budget comparison with more traditional pressure-fed micro-rockets shows a real benefit from this system in terms of mass reduction. In the paper, an overview of the project status in Mechatronic is presented.

Examination of the liver in personnel working with liquid rocket propellant

PubMed Central

Petersen, Palle; Bredahl, Erik; Lauritsen, Ove; Laursen, Thomas

1970-01-01

Petersen, P., Bredahl, E., Lauritsen, O., and Laursen, T. (1970).Brit. J. industr. Med.,27, 141-146. Examination of the liver in personnel working with liquid rocket propellants. Personnel working with liquid rocket propellants were subjected to routine health examinations, including liver function tests, as the propellant, unsymmetrical dimethylhydrazine (UDMH) is potentially toxic to the liver. In 46 persons the concentrations of serum alanine aminotransferase (SGPT) were raised. Liver biopsy was performed in 26 of these men; 6 specimens were pathological (fatty degeneration), 5 were uncertain, and 15 were normal. All 6 pathological biopsies were from patients with a raised SGPT at the time of biopsy. Of the 15 persons with a normal liver biopsy, 14 had a normal SGPT, while one (who was an alcoholic) had a raised SGPT. The connection between SGPT and histology of the liver, as well as the possible causal relation between the pathological findings and exposure to UDMH, is discussed. Images PMID:5428632

State and prospects of solid propellant rocket development

NASA Astrophysics Data System (ADS)

Kukushkin, V. Kh.

1992-07-01

An overview is presented of aspects of solid-propellant rocket engine (SPRE) development with individual treatment given to sustainer and spacecraft SPRE technologies. The paper focuses on low-modulus fuels of composite solid propellant, requirements for adhesion stability, and enhancement of the power characteristics of solid propellants. R&D activities are described that relate to the use of SPREs with extending nozzles and to the design of ultradimensional nozzles for upper-stage engines. Other developments for the SPREs include engines with separate loading and pasty fuel applications, and progress is reported in the direction of detonation SPREs. The SPREs using pasty propellants provide good control over thrust characteristics and fuel qualities. A device is incorporated that assures fuel burning in the combustion region and reliable ignition during restarting of these engines.

Cooling Duct Analysis for Transpiration/Film Cooled Liquid Propellant Rocket Engines

NASA Technical Reports Server (NTRS)

Micklow, Gerald J.

1996-01-01

The development of a low cost space transportation system requires that the propulsion system be reusable, have long life, with good performance and use low cost propellants. Improved performance can be achieved by operating the engine at higher pressure and temperature levels than previous designs. Increasing the chamber pressure and temperature, however, will increase wall heating rates. This necessitates the need for active cooling methods such as film cooling or transpiration cooling. But active cooling can reduce the net thrust of the engine and add considerably to the design complexity. Recently, a metal drawing process has been patented where it is possible to fabricate plates with very small holes with high uniformity with a closely specified porosity. Such a metal plate could be used for an inexpensive transpiration/film cooled liner to meet the demands of advanced reusable rocket engines, if coolant mass flow rates could be controlled to satisfy wall cooling requirements and performance. The present study investigates the possibility of controlling the coolant mass flow rate through the porous material by simple non-active fluid dynamic means. The coolant will be supplied to the porous material by series of constant geometry slots machined on the exterior of the engine.

NASA Collaborative Design Processes

NASA Technical Reports Server (NTRS)

Jones, Davey

2017-01-01

This is Block 1, the first evolution of the world's most powerful and versatile rocket, the Space Launch System, built to return humans to the area around the moon. Eventually, larger and even more powerful and capable configurations will take astronauts and cargo to Mars. On the sides of the rocket are the twin solid rocket boosters that provide more than 75 percent during liftoff and burn for about two minutes, after which they are jettisoned, lightening the load for the rest of the space flight. Four RS-25 main engines provide thrust for the first stage of the rocket. These are the world's most reliable rocket engines. The core stage is the main body of the rocket and houses the fuel for the RS-25 engines, liquid hydrogen and liquid oxygen, and the avionics, or "brain" of the rocket. The core stage is all new and being manufactured at NASA's "rocket factory," Michoud Assembly Facility near New Orleans. The Launch Vehicle Stage Adapter, or LVSA, connects the core stage to the Interim Cryogenic Propulsion Stage. The Interim Cryogenic Propulsion Stage, or ICPS, uses one RL-10 rocket engine and will propel the Orion spacecraft on its deep-space journey after first-stage separation. Finally, the Orion human-rated spacecraft sits atop the massive Saturn V-sized launch vehicle. Managed out of Johnson Space Center in Houston, Orion is the first spacecraft in history capable of taking humans to multiple destinations within deep space. 2) Each element of the SLS utilizes collaborative design processes to achieve the incredible goal of sending human into deep space. Early phases are focused on feasibility and requirements development. Later phases are focused on detailed design, testing, and operations. There are 4 basic phases typically found in each phase of development.

Cooling of in-situ propellant rocket engines for Mars mission. M.S. Thesis - Cleveland State Univ.

NASA Technical Reports Server (NTRS)

Armstrong, Elizabeth S.

1991-01-01

One propulsion option of a Mars ascent/descent vehicle is multiple high-pressure, pump-fed rocket engines using in-situ propellants, which have been derived from substances available on the Martian surface. The chosen in-situ propellant combination for this analysis is carbon monoxide as the fuel and oxygen as the oxidizer. Both could be extracted from carbon dioxide, which makes up 96 percent of the Martian atmosphere. A pump-fed rocket engine allows for higher chamber pressure than a pressure-fed engine, which in turn results in higher thrust and in higher heat flux in the combustion chamber. The heat flowing through the wall cannot be sufficiently dissipated by radiation cooling and, therefore, a regenerative coolant may be necessary to avoid melting the rocket engine. The two possible fluids for this coolant scheme, carbon monoxide and oxygen, are compared analytically. To determine their heat transfer capability, they are evaluated based upon their heat transfer and fluid flow characteristics.

Materials for advanced rocket engine turbopump turbine blades

NASA Technical Reports Server (NTRS)

Chandler, W. T.

1985-01-01

A study program was conducted to identify those materials that will provide the greatest benefits as turbine blades for advanced liquid propellant rocket engine turbines and to prepare technology plans for the development of those materials for use in the 1990 through 1995 period. The candidate materials were selected from six classes of materials: single-crystal (SC) superalloys, oxide dispersion-strengthened (ODS) superalloys, rapid solidification processed (RSP) superalloys, directionally solidified eutectic (DSE) superalloys, fiber-reinforced superalloy (FRS) composites, and ceramics. Properties of materials from the six classes were compiled and evaluated and property improvements were projected approximately 5 years into the future for advanced versions of materials in each of the six classes.

Study of solid rocket motor for space shuttle booster, volume 2, book 1

NASA Technical Reports Server (NTRS)

1972-01-01

The technical requirements for the solid propellant rocket engine to be used with the space shuttle orbiter are presented. The subjects discussed are: (1) propulsion system definition, (2) solid rocket engine stage design, (3) solid rocket engine stage recovery, (4) environmental effects, (5) manrating of the solid rocket engine stage, (6) system safety analysis, and (7) ground support equipment.

Metallized Gelled Propellants: Oxygen/RP-1/Aluminum Rocket Heat Transfer and Combustion Measurements

NASA Technical Reports Server (NTRS)

Palaszewski, Bryan; Zakany, James S.

1996-01-01

A series of rocket engine heat transfer experiments using metallized gelled liquid propellants was conducted. These experiments used a small 20- to 40-lb/f thrust engine composed of a modular injector, igniter, chamber and nozzle. The fuels used were traditional liquid RP-1 and gelled RP-1 with 0-, 5-, and 55-percentage by weight loadings of aluminum particles. Gaseous oxygen was used as the oxidizer. Three different injectors were used during the testing: one for the baseline O(2)/RP-1 tests and two for the gelled and metallized gelled fuel firings. Heat transfer measurements were made with a rocket engine calorimeter chamber and nozzle with a total of 31 cooling channels. Each chamber used a water flow to carry heat away from the chamber and the attached thermocouples and flow meters allowed heat flux estimates at each of the 31 stations. The rocket engine Cstar efficiency for the RP-1 fuel was in the 65-69 percent range, while the gelled 0 percent by weight RP-1 and the 5-percent by weight RP-1 exhibited a Cstar efficiency range of 60 to 62% and 65 to 67%, respectively. The 55-percent by weight RP-1 fuel delivered a 42-47% Cstar efficiency. Comparisons of the heat flux and temperature profiles of the RP-1 and the metallized gelled RP-1/A1 fuels show that the peak nozzle heat fluxes with the metallized gelled O2/RP-1/A1 propellants are substantially higher than the baseline O2/RP-1: up to double the flux for the 55 percent by weight RP-1/A1 over the RP-1 fuel. Analyses showed that the heat transfer to the wall was significantly different for the RP-1/A1 at 55-percent by weight versus the RP-1 fuel. Also, a gellant and an aluminum combustion delay was inferred in the 0 percent and 5-percent by weight RP-1/A1 cases from the decrease in heat flux in the first part of the chamber. A large decrease in heat flux in the last half of the chamber was caused by fuel deposition in the chamber and nozzle. The engine combustion occurred well downstream of the injector face based on the heat flux estimates from the temperature measurements.

A study on various methods of supplying propellant to an orbit insertion rocket engine

NASA Technical Reports Server (NTRS)

Boretz, J. E.; Huniu, S.; Thompson, M.; Pagani, M.; Paulsen, B.; Lewis, J.; Paul, D.

1980-01-01

Various types of pumps and pump drives were evaluated to determine the lightest weight system for supplying propellants to a planetary orbit insertion rocket engine. From these analyses four candidate propellant feed systems were identified. Systems Nos. 1 and 2 were both battery powered (lithium-thionyl-chloride or silver-zinc) motor driven pumps. System 3 was a monopropellant gas generator powered turbopump. System 4 was a bipropellant gas generator powered turbopump. Parameters considered were pump break horsepower, weight, reliability, transient response and system stability. Figures of merit were established and the ranking of the candidate systems was determined. Conceptual designs were prepared for typical motor driven pumps and turbopump configurations for a 1000 lbf thrust rocket engine.

America's first long-range-missile and space exploration program: The ORDCIT project of the Jet Propulsion Laboratory, 1943 - 1946: A memoir

NASA Technical Reports Server (NTRS)

Malina, F. J.

1977-01-01

Research and achievements of the wartime Jet Propulsion Laboratory are outlined. Accomplishments included development of the solid-propellant Private A and private R rockets and the liquid-propellant nitric acid-aniline WAC Corporal rocket.

Method and apparatus to produce high specific impulse and moderate thrust from a fusion-powered rocket engine

DOE Office of Scientific and Technical Information (OSTI.GOV)

Cohen, Samuel A.; Pajer, Gary A.; Paluszek, Michael A.

A system and method for producing and controlling high thrust and desirable specific impulse from a continuous fusion reaction is disclosed. The resultant relatively small rocket engine will have lower cost to develop, test, and operate that the prior art, allowing spacecraft missions throughout the planetary system and beyond. The rocket engine method and system includes a reactor chamber and a heating system for heating a stable plasma to produce fusion reactions in the stable plasma. Magnets produce a magnetic field that confines the stable plasma. A fuel injection system and a propellant injection system are included. The propellant injectionmore » system injects cold propellant into a gas box at one end of the reactor chamber, where the propellant is ionized into a plasma. The propellant and fusion products are directed out of the reactor chamber through a magnetic nozzle and are detached from the magnetic field lines producing thrust.« less

Cold Flow Testing for Liquid Propellant Rocket Injector Scaling and Throttling

NASA Technical Reports Server (NTRS)

Kenny, Jeremy R.; Moser, Marlow D.; Hulka, James; Jones, Gregg

2006-01-01

Scaling and throttling of combustion devices are important capabilities to demonstrate in development of liquid rocket engines for NASA's Space Exploration Mission. Scaling provides the ability to design new injectors and injection elements with predictable performance on the basis of test experience with existing injectors and elements, and could be a key aspect of future development programs. Throttling is the reduction of thrust with fixed designs and is a critical requirement in lunar and other planetary landing missions. A task in the Constellation University Institutes Program (CUIP) has been designed to evaluate spray characteristics when liquid propellant rocket engine injectors are scaled and throttled. The specific objectives of the present study are to characterize injection and primary atomization using cold flow simulations of the reacting sprays. These simulations can provide relevant information because the injection and primary atomization are believed to be the spray processes least affected by the propellant reaction. Cold flow studies also provide acceptable test conditions for a university environment. Three geometric scales - 1/4- scale, 1/2-scale, and full-scale - of two different injector element types - swirl coaxial and shear coaxial - will be designed, fabricated, and tested. A literature review is currently being conducted to revisit and compile the previous scaling documentation. Because it is simple to perform, throttling will also be examined in the present work by measuring primary atomization characteristics as the mass flow rate and pressure drop of the six injector element concepts are reduced, with corresponding changes in chamber backpressure. Simulants will include water and gaseous nitrogen, and an optically accessible chamber will be used for visual and laser-based diagnostics. The chamber will include curtain flow capability to repress recirculation, and additional gas injection to provide independent control of the backpressure. This paper provides a short review of the appropriate literature, as well as descriptions of plans for experimental hardware, test chamber instrumentation, diagnostics, and testing.

Injection and swirl driven flowfields in solid and liquid rocket motors

NASA Astrophysics Data System (ADS)

Vyas, Anand B.

In this work, we seek approximate analytical solutions to describe the bulk flow motion in certain types of solid and liquid rocket motors. In the case of an idealized solid rocket motor, a cylindrical double base propellant grain with steady regression rate is considered. The well known inviscid profile determined by Culick is extended here to include the effects of viscosity and steady grain regression. The approximate analytical solution for the cold flow is obtained from similarity principles, perturbation methods and the method of variation of parameters. The velocity, vorticity, pressure gradient and the shear stress distributions are determined and interpreted for different rates of wall regression and injection Reynolds number. The liquid propellant rocket engine considered here is based on a novel design that gives rise to a cyclonic flow. The resulting bidirectional motion is triggered by the tangential injection of an oxidizer just upstream of the chamber nozzle. Velocity, vorticity and pressure gradient distributions are determined for the bulk gas dynamics using a non-reactive inviscid model. Viscous corrections are then incorporated to explain the formation of a forced vortex near the core. Our results compare favorably with numerical simulations and experimental measurements obtained by other researchers. They also indicate that the bidirectional vortex in a cylindrical chamber is a physical solution of the Euler equations. In closing, we investigate the possibility of multi-directional flow behavior as predicted by Euler's equation and as reported recently in laboratory experiments.

Control of Propellant Lead/Lag to the LAE in the AXAF Propulsion System

NASA Technical Reports Server (NTRS)

Casillas, A. R.; Eninger, J.; Joseph, G.; Kenney, J.; Trinidad, M.

1998-01-01

Control of the rate at which hypergolic propellants are supplied to a rocket engine prior to ignition is critically important. Potentially damaging explosions may result from excessive lead of either propellant into the combustion chamber. Because the injector fill process is governed by the engine as well as the propellant feed system design, proper management of this issue must take both into consideration. This was recognized early in the development of TRW's Advanced Columbium-Liquid Apogee Engine (LAE), which was flight-qualified in 1996 to maneuver the Advanced X-Ray Astrophysics Facility (AXAF) spacecraft into orbit. The LAE runs on hydrazine and nitrogen tetroxide (MON-3) at a nominal mixture ratio of 1.0. This paper describes the comprehensive test program conducted to ensure reliable startup operation of the LAE in the AYAF propulsion system. The most significant factors affecting chamber fuel lead were found to be: (1) engine location, (2) propellant saturation level, (3) amount of undissolved gas in the lines, and (4) off- nominal tank pressures. Hot-fire tests at a chamber fuel lead range over and above that expected for the LAEs in AXAF demonstrated extremely tolerant behavior of the engine. AY-AF is scheduled for launch on NASA's STS-93 in December 1998.

Active chlorine and nitric oxide formation from chemical rocket plume afterburning

NASA Astrophysics Data System (ADS)

Leone, D. M.; Turns, S. R.

Chlorine and oxides of nitrogen (NO(x)) released into the atmosphere contribute to acid rain (ground level or low-altitude sources) and ozone depletion from the stratosphere (high-altitude sources). Rocket engines have the potential for forming or activating these pollutants in the rocket plume. For instance, H2/O2 rockets can produce thermal NO(x) in their plumes. Emphasis, in the past, has been placed on determining the impact of chlorine release on the stratosphere. To date, very little, if any, information is available to understand what contribution NO(x) emissions from ground-based engine testing and actual rocket launches have on the atmosphere. The goal of this work is to estimate the afterburning emissions from chemical rocket plumes and determine their local stratospheric impact. Our study focuses on the space shuttle rocket motors, which include both the solid rocket boosters (SRB's) and the liquid propellant main engines (SSME's). Rocket plume afterburning is modeled employing a one-dimensional model incorporating two chemical kinetic systems: chemical and thermal equilibria with overlayed nitric oxide chemical kinetics (semi equilibrium) and full finite-rate chemical kinetics. Additionally, the local atmospheric impact immediately following a launch is modeled as the emissions diffuse and chemically react in the stratosphere.

Active chlorine and nitric oxide formation from chemical rocket plume afterburning

NASA Technical Reports Server (NTRS)

Leone, D. M.; Turns, S. R.

1994-01-01

Chlorine and oxides of nitrogen (NO(x)) released into the atmosphere contribute to acid rain (ground level or low-altitude sources) and ozone depletion from the stratosphere (high-altitude sources). Rocket engines have the potential for forming or activating these pollutants in the rocket plume. For instance, H2/O2 rockets can produce thermal NO(x) in their plumes. Emphasis, in the past, has been placed on determining the impact of chlorine release on the stratosphere. To date, very little, if any, information is available to understand what contribution NO(x) emissions from ground-based engine testing and actual rocket launches have on the atmosphere. The goal of this work is to estimate the afterburning emissions from chemical rocket plumes and determine their local stratospheric impact. Our study focuses on the space shuttle rocket motors, which include both the solid rocket boosters (SRB's) and the liquid propellant main engines (SSME's). Rocket plume afterburning is modeled employing a one-dimensional model incorporating two chemical kinetic systems: chemical and thermal equilibria with overlayed nitric oxide chemical kinetics (semi equilibrium) and full finite-rate chemical kinetics. Additionally, the local atmospheric impact immediately following a launch is modeled as the emissions diffuse and chemically react in the stratosphere.

A rapid method for optimization of the rocket propulsion system for single-stage-to-orbit vehicles

NASA Technical Reports Server (NTRS)

Eldred, C. H.; Gordon, S. V.

1976-01-01

A rapid analytical method for the optimization of rocket propulsion systems is presented for a vertical take-off, horizontal landing, single-stage-to-orbit launch vehicle. This method utilizes trade-offs between propulsion characteristics affecting flight performance and engine system mass. The performance results from a point-mass trajectory optimization program are combined with a linearized sizing program to establish vehicle sizing trends caused by propulsion system variations. The linearized sizing technique was developed for the class of vehicle systems studied herein. The specific examples treated are the optimization of nozzle expansion ratio and lift-off thrust-to-weight ratio to achieve either minimum gross mass or minimum dry mass. Assumed propulsion system characteristics are high chamber pressure, liquid oxygen and liquid hydrogen propellants, conventional bell nozzles, and the same fixed nozzle expansion ratio for all engines on a vehicle.

Theory and Design of Flight-Vehicle Engines

NASA Technical Reports Server (NTRS)

Zhdanov, V. T. (Editor); Kurziner, R. I. (Editor)

1987-01-01

Papers are presented on such topics as the testing of aircraft engines, errors in the experimental determination of the parameters of scramjet engines, the effect of the nonuniformity of supersonic flow with shocks on friction and heat transfer in the channel of a hypersonic ramjet engine, and the selection of the basic parameters of cooled GTE turbines. Consideration is also given to the choice of optimal total wedge angle for the acceleration of aerospace vehicles, the theory of an electromagnetic-resonator engine, the dynamic characteristics of the pumps and turbines of liquid propellant rocket engines in transition regimes, and a hierarchy of mathematical models for spacecraft control engines.

Handbook of recommended practices for the determination of liquid monopropellant rocket engine performance

NASA Technical Reports Server (NTRS)

Bjorklund, R. A.; Rogero, R. S.; Baerwald, R. K.

1979-01-01

The design, installation, and operation of systems to be used for directly measuring quantities of fundamental importance to the determination of monopropellant thruster performance is described. Areas covered include: (1) force and impulse measurement; (2) propellant mass usage and flow measurement; (3) pressure measurement; (4) temperature measurement; (5) exhaust gas composition measurement; and (6) data reduction and performance determination.

LOX/hydrocarbon rocket engine analytical design methodology development and validation. Volume 1: Executive summary and technical narrative

NASA Technical Reports Server (NTRS)

Pieper, Jerry L.; Walker, Richard E.

1993-01-01

During the past three decades, an enormous amount of resources were expended in the design and development of Liquid Oxygen/Hydrocarbon and Hydrogen (LOX/HC and LOX/H2) rocket engines. A significant portion of these resources were used to develop and demonstrate the performance and combustion stability for each new engine. During these efforts, many analytical and empirical models were developed that characterize design parameters and combustion processes that influence performance and stability. Many of these models are suitable as design tools, but they have not been assembled into an industry-wide usable analytical design methodology. The objective of this program was to assemble existing performance and combustion stability models into a usable methodology capable of producing high performing and stable LOX/hydrocarbon and LOX/hydrogen propellant booster engines.

Worldwide Space Launch Vehicles and Their Mainstage Liquid Rocket Propulsion

NASA Technical Reports Server (NTRS)

Rahman, Shamim A.

2010-01-01

Space launch vehicle begins with a basic propulsion stage, and serves as a missile or small launch vehicle; many are traceable to the 1945 German A-4. Increasing stage size, and increasingly energetic propulsion allows for heavier payloads and greater. Earth to Orbit lift capability. Liquid rocket propulsion began with use of storable (UDMH/N2O4) and evolved to high performing cryogenics (LOX/RP, and LOX/LH). Growth versions of SLV's rely on strap-on propulsive stages of either solid propellants or liquid propellants.

Cryogenic Fluid Management Technology and Nuclear Thermal Propulsion

NASA Technical Reports Server (NTRS)

Taylor, Brian D.; Caffrey, Jarvis; Hedayat, Ali; Stephens, Jonathan; Polsgrove, Robert

2016-01-01

Cryogenic fluid management (CFM) is critical to the success of future nuclear thermal propulsion powered vehicles. While this is an issue for any propulsion system utilizing cryogenic propellants, this is made more challenging by the radiation flux produced by the reactor in a nuclear thermal rocket (NTR). Managing the cryogenic fuel to prevent propellant loss to boil off and leakage is needed to limit the required quantity of propellant to a reasonable level. Analysis shows deposition of energy into liquid hydrogen fuel tanks in the vicinity of the nuclear thermal engine. This is on top of ambient environment sources of heat. Investments in cryogenic/thermal management systems (some of which are ongoing at various organizations) are needed in parallel to nuclear thermal engine development in order to one day see the successful operation of an entire stage. High durability, low thermal conductivity insulation is one developmental need. Light weight cryocoolers capable of removing heat from large fluid volumes at temperatures as low as approx. 20 K are needed to remove heat leak from the propellant of an NTR. Valve leakage is an additional CFM issue of great importance. Leakage rates of state of the art, launch vehicle size valves (which is approximately the size valves needed for a Mars transfer vehicle) are quite high and would result in large quantities of lost propellant over a long duration mission. Additionally, the liquid acquisition system inside the propellant tank must deliver properly conditioned propellant to the feed line for successful engine operation and avoid intake of warm or gaseous propellant. Analysis of the thermal environment and the CFM technology development are discussed in the accompanying presentation.

Additive Manufacturing a Liquid Hydrogen Rocket Engine

NASA Technical Reports Server (NTRS)

Jones, Carl P.; Robertson, Elizabeth H.; Koelbl, Mary Beth; Singer, Chris

2016-01-01

Space Propulsion is a 5 day event being held from 2nd May to the 6th May 2016 at the Rome Marriott Park Hotel in Rome, Italy. This event showcases products like Propulsion sub-systems and components, Production and manufacturing issues, Liquid, Solid, Hybrid and Air-breathing Propulsion Systems for Launcher and Upper Stages, Overview of current programmes, AIV issues and tools, Flight testing and experience, Technology building blocks for Future Space Transportation Propulsion Systems : Launchers, Exploration platforms & Space Tourism, Green Propulsion for Space Transportation, New propellants, Rocket propulsion & global environment, Cost related aspects of Space Transportation propulsion, Modelling, Pressure-Thrust oscillations issues, Impact of new requirements and regulations on design etc. in the Automotive, Manufacturing, Fabrication, Repair & Maintenance industries.

Evaluation of the Effect of Exhausts from Liquid and Solid Rockets on Ozone Layer

NASA Astrophysics Data System (ADS)

Yamagiwa, Yoshiki; Ishimaki, Tetsuya

This paper reports the analytical results of the influences of solid rocket and liquid rocket exhausts on ozone layer. It is worried about that the exhausts from solid propellant rockets cause the ozone depletion in the ozone layer. Some researchers try to develop the analytical model of ozone depletion by rocket exhausts to understand its physical phenomena and to find the effective design of rocket to minimize its effect. However, these models do not include the exhausts from liquid rocket although there are many cases to use solid rocket boosters with a liquid rocket at the same time in practical situations. We constructed combined analytical model include the solid rocket exhausts and liquid rocket exhausts to analyze their effects. From the analytical results, we find that the exhausts from liquid rocket suppress the ozone depletion by solid rocket exhausts.

Evaluation of undeveloped rocket engine cycle applications to advanced transportation

NASA Technical Reports Server (NTRS)

1990-01-01

Undeveloped pump-fed, liquid propellant rocket engine cycles were assessed and evaluated for application to Next Manned Transportation System (NMTS) vehicles, which would include the evolving Space Transportation System (STS Evolution), the Personnel Launch System (PLS), and the Advanced Manned Launch System (AMLS). Undeveloped engine cycles selected for further analysis had potential for increased reliability, more maintainability, reduced cost, and improved (or possibly level) performance when compared to the existing SSME and proposed STME engines. The split expander (SX) cycle, the full flow staged combustion (FFSC) cycle, and a hybrid version of the FFSC, which has a LOX expander drive for the LOX pump, were selected for definition and analysis. Technology requirements and issues were identified and analyses of vehicle systems weight deltas using the SX and FFSC cycles in AMLS vehicles were performed. A strawman schedule and cost estimate for FFSC subsystem technology developments and integrated engine system demonstration was also provided.

Astronautics

NASA Technical Reports Server (NTRS)

1977-01-01

Principles of rocket engineering, flight dynamics, and trajectories are discussed in this summary of Soviet rocket development and technology. Topics include rocket engine design, propellants, propulsive efficiency, and capabilities required for orbital launch. The design of the RD 107, 108, 119, and 214 rocket engines and their uses in various satellite launches are described. NASA's Saturn 5 and Atlas Agena launch vehicles are used to illustrate the requirements of multistage rockets.

Europa Sample Return Mission Utilizing High Specific Impulse Propulsion Refueled with Indigenous Resources

NASA Astrophysics Data System (ADS)

Paniagua, J.; Powell, J. R.; Maise, G.

2002-01-01

We have conducted studies of a revolutionary new concept for conducting a Europa Sample Return Mission. Robotic spacecraft exploration of the Solar System has been severely constrained by the large energy requirements of interplanetary trajectories and the inherent delta V limitations of chemical rockets. Current missions use gravitational assists from intermediate planets to achieve these high-energy trajectories restricting payload size and increasing flight times. We propose a 6-year Europa Sample Return mission with very modest launch requirements enabled by MITEE. A new nuclear thermal propulsion engine design, termed MITEE (MIniature reacTor EnginE), has over twice the delta V capability of H2/O2 rockets (and much greater when refueled with H2 propellant from indigenous extraterrestrial resources) enabling unique missions that are not feasible with chemical propulsion. The MITEE engine is a compact, ultra-lightweight, thermal nuclear rocket that uses hydrogen as the propellant. MITEE, with its small size (50 cm O.D.), low mass (200 kg), and high specific impulse (~1000 sec), can provide a quantum leap in the capability for space science and exploration missions. The Robotic Europa Explorer (REE) spacecraft has a two-year outbound direct trajectory and lands on the satellite surface for an approximate 9 month stay. During this time, the vehicle is refueled with H2 propellant derived from Europa ice by the Autonomous Propellant Producer (APP), while collecting samples and searching for life. A small nuclear-heated submarine probe, the Autonomous Submarine Vehicle (ASV), based on MITEE technology, would melt through the ice and explore the undersea realm. The spacecraft has approximately a three year return to Earth after departure from Europa with samples onboard. Spacecraft payload is 430 kg at the start of the mission and can be launched with a single, conventional medium-sized Delta III booster. The spacecraft can bring back 25 kg of samples from Europa. Europa, in the Jovian system, is a high priority target for an outer Solar System exploration mission. More than a decade ago the Voyager spacecraft revealed Europa as a world swathed in ice and geologically young. NASA's Galileo spacecraft passed approximately 500 miles above the surface and provided detailed images of Europa's terrain marked by a dynamic topology that appeared to be remnants of ice volcanoes or geysers. The surface temperature averages a chilly -200° C. The pictures appear to show a relatively young surface of ice, possibly only 1 km thick in some places. Internal heating of Europa from Jupiter's tidal pull could form an ocean of liquid water beneath the surface. More recently, Ganymede and Callisto are believed to be ocean-bearing Jovian moons based on magnetometer measurements from the Galileo spacecraft. If liquid water exists, life may also. NASA plans to send an orbiting spacecraft to Europa to measure the thickness of the ice and to detect if an underlying liquid ocean exists. This mission would precede the proposed Europa Sample Return mission, which includes dispatching an autonomous submarine-like vehicle that could melt through the ice and explore the undersea realm. Because of the large energy requirements typical of these ambitious solar system science missions, use of chemical rockets results in interplanetary spacecraft that are prohibitive in terms of Initial Mass in Low- Earth Orbit (IMLEO) and cost. For example, using chemical rockets to return samples from Europa appears to be technically impractical, as it would require large delta V and launch vehicle capabilities. On the other hand, use of nuclear thermal rockets will significantly reduce IMLEO and, subsequently, costs. Moreover, nuclear thermal rockets can utilize extraterrestrial resources as propellants, an option not practical with chemical rockets. This "refueling" capability would enable nuclear rockets to carry out very high-energy missions, such as the return of large amounts of extraterrestrial material to Earth. The Europa missions considered in this proposal will be restricted to starting from LEO only after being placed in a stable orbit by a launch vehicle. This simplifies and eases the safety issues and mitigates political concerns. High propulsive efficiency of the MITEE engine yields the benefits of reduced transit time and a smaller launch vehicle.

Liquid Rocket Booster (LRB) for the Space Transportation System (STS) systems study. Appendix A: Stress analysis report for the pump-fed and pressure-fed liquid rocket booster

NASA Technical Reports Server (NTRS)

1989-01-01

Pressure effects on the pump-fed Liquid Rocket Booster (LRB) of the Space Transportation System are examined. Results from the buckling tests; bending moments tests; barrel, propellant tanks, frame XB1513, nose cone, and intertank tests; and finite element examination of forward and aft skirts are presented.

Development of Ionic Liquid Monopropellants for In-Space Propulsion

NASA Technical Reports Server (NTRS)

Blevins, John A.; Osborne, Robin; Drake, Gregory W.

2005-01-01

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.

An experimental investigation of liquid methane convection and boiling in rocket engine cooling channels

NASA Astrophysics Data System (ADS)

Trujillo, Abraham Gerardo

In the past decades, interest in developing hydrocarbon-fueled rocket engines for deep spaceflight missions has continued to grow. In particular, liquid methane (LCH4) has been of interest due to the weight efficiency, storage, and handling advantages it offers over several currently used propellants. Deep space exploration requires reusable, long life rocket engines. Due to the high temperatures reached during combustion, the life of an engine is significantly impacted by the cooling system's efficiency. Regenerative (regen) cooling is presented as a viable alternative to common cooling methods such as film and dump cooling since it provides improved engine efficiency. Due to limited availability of experimental sub-critical liquid methane cooling data for regen engine design, there has been an interest in studying the heat transfer characteristics of the propellant. For this reason, recent experimental studies at the Center for Space Exploration Technology Research (cSETR) at the University of Texas at El Paso (UTEP) have focused on investigating the heat transfer characteristics of sub-critical CH4 flowing through sub-scale cooling channels. To conduct the experiments, the csETR developed a High Heat Flux Test Facility (HHFTF) where all the channels are heated using a conduction-based thermal concentrator. In this study, two smooth channels with cross sectional geometries of 1.8 mm x 4.1 mm and 3.2 mm x 3.2 mm were tested. In addition, three roughened channels all with a 3.2 mm x 3.2 mm square cross section were also tested. For the rectangular smooth channel, Reynolds numbers ranged between 68,000 and 131,000, while the Nusselt numbers were between 40 and 325. For the rough channels, Reynolds numbers ranged from 82,000 to 131,000, and Nusselt numbers were between 65 and 810. Sub-cooled film-boiling phenomena were confirmed for all the channels presented in this work. Film-boiling onset at Critical Heat Flux (CHF) was correlated to a Boiling Number (Bo) of approximately 0.1 for all channels. Convective Nusselt number follows predicted trends for Reynolds number with a wall temperature correction for both the boiling and non-boiling regimes.

KSC-2012-4455

NASA Image and Video Library

2012-08-14

CAPE CANAVERAL, Fla. – At NASA’s Kennedy Space Center in Florida, a crane is used to load the aft skirt for a space shuttle solid rocket booster on a truck. A twin set of space shuttle solid rocket boosters and an external fuel tank are being prepared for transport to separate museums. The solid rocket boosters, or SRBs, will be displayed at the California Science Center in Los Angeles. The external tank soon will be transported for display at the Wings of Dreams Aviation Museum at Keystone Heights Airport between Gainesville and Jacksonville, Fla. The 149-foot SRBs together provided six million pounds of thrust. The external fuel tank contained over 500,000 gallons of liquid hydrogen and liquid oxygen propellant for the shuttle orbiters' three main engines. The work is part of Transition and Retirement of the space shuttle. For more information, visit http://www.nasa.gov/transition Photo credit: NASA/ Dimitri Gerondidakis

Quick trips to Mars

NASA Technical Reports Server (NTRS)

Hornung, R.

1991-01-01

The design of a Mars Mission Vehicle that would have to be launched by two very heavy lift launch vehicles is described along with plans for a mission to Mars. The vehicle has three nuclear engine for rocket vehicle application (NERVA) boosters with a fourth in the center that acts as a dual mode system. The fourth generates electrical power while in route, but it also helps lift the vehicle out of earth orbit. A Mars Ascent Vehicle (MAV), a Mars transfer vehicle stage, and a Mars Excursion Vehicle (MEV) are located on the front end of this vehicle. Other aspects of this research including aerobraking, heat shielding, nuclear thermal rocket engines, a mars mission summary, closed Brayton cycle with and without regeneration, liquid hydrogen propellant storage, etc. are addressed.

Development of Ionic Liquid Monopropellants for In-Space Propulsion

NASA Technical Reports Server (NTRS)

Blevins, John A.; Drake, Gregory W.; Osborne, Robin J.

2005-01-01

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.

Combining MHD Airbreathing and Fusion Rocket Propulsion for Earth-to-Orbit Flight

NASA Astrophysics Data System (ADS)

Froning, H. D.; Miley, G. H.; Luo, Nie; Yang, Yang; Momota, H.; Burton, E.

2005-02-01

Previous studies have shown that Single-State-to-Orbit (SSTO) vehicle propellant can be reduced by Magnets-Hydro-Dynamic (MHD) processes that minimize airbreathing propulsion losses and propellant consumption during atmospheric flight. Similarly additional reduction in SSTO propellant is enabled by Inertial Electrostatic Confinement (IEC) fusion, whose more energetic reactions reduce rocket propellant needs. MHD airbreathing propulsion during an SSTO vehicle's initial atmospheric flight phase and IEC fusion propulsion during its final exo-atmospheric flight phase is therefore being explored. Accomplished work is not yet sufficient for claiming such a vehicle's feasibility. But takeoff and propellant mass for an MHD airbreathing and IEC fusion vehicle could be as much as 25 and 40 percent less than one with ordinary airbreathing and IEC fusion; and as much as 50 and 70 percent less than SSTO takeoff and propellant mass with MHD airbreathing and chemical rocket propulsion. Thus this unusual combined cycle engine shows great promise for performance gains beyond contemporary combined-cycle airbreathing engines.

X-ray Fluorescence Measurements of Turbulent Methane-Oxygen Shear Coaxial Flames

DTIC Science & Technology

2015-05-01

The shear coaxial jet injector is a typical injector design in liquid rocket engines, used as the main chamber element for Space Shuttle Main...current study. (b) Representation of the injector tip of the shear coaxial burner with propellant streams and dimensions labeled. (c) Picture of flame...integrated with the Air Force Research Laboratories’ (AFRL) Mobile Flow Laboratory (MFL). This facility is designed to allow aerospace-propulsion injector

Bleed cycle propellant pumping in a gas-core nuclear rocket engine system

NASA Technical Reports Server (NTRS)

Kascak, A. F.; Easley, A. J.

1972-01-01

The performance of ideal and real staged primary propellant pumps and bleed-powered turbines was calculated for gas-core nuclear rocket engines over a range of operating pressures from 500 to 5000 atm. This study showed that for a required engine operating pressure of 1000 atm the pump work was about 0.8 hp/(lb/sec), the specific impulse penalty resulting from the turbine propellant bleed flow as about 10 percent; and the heat required to preheat the propellant was about 7.8 MN/(lb/sec). For a specific impulse above 2400 sec, there is an excess of energy available in the moderator due to the gamma and neutron heating that occurs there. Possible alternative pumping cycles are the Rankine or Brayton cycles.

Development of Efficient Real-Fluid Model in Simulating Liquid Rocket Injector Flows

NASA Technical Reports Server (NTRS)

Cheng, Gary; Farmer, Richard

2003-01-01

The characteristics of propellant mixing near the injector have a profound effect on the liquid rocket engine performance. However, the flow features near the injector of liquid rocket engines are extremely complicated, for example supercritical-pressure spray, turbulent mixing, and chemical reactions are present. Previously, a homogeneous spray approach with a real-fluid property model was developed to account for the compressibility and evaporation effects such that thermodynamics properties of a mixture at a wide range of pressures and temperatures can be properly calculated, including liquid-phase, gas- phase, two-phase, and dense fluid regions. The developed homogeneous spray model demonstrated a good success in simulating uni- element shear coaxial injector spray combustion flows. However, the real-fluid model suffered a computational deficiency when applied to a pressure-based computational fluid dynamics (CFD) code. The deficiency is caused by the pressure and enthalpy being the independent variables in the solution procedure of a pressure-based code, whereas the real-fluid model utilizes density and temperature as independent variables. The objective of the present research work is to improve the computational efficiency of the real-fluid property model in computing thermal properties. The proposed approach is called an efficient real-fluid model, and the improvement of computational efficiency is achieved by using a combination of a liquid species and a gaseous species to represent a real-fluid species.

Marshall Team Fires Recreated Goddard Rocket

NASA Technical Reports Server (NTRS)

2003-01-01

In honor of the Centernial of Flight Celebration and commissioned by the American Institute of Aeronautics and Astronautics (AIAA), a team of engineers from Marshall Space Flight Center (MSFC) built a replica of the first liquid-fueled rocket. The original rocket, designed and built by rocket engineering pioneer Robert H. Goddard in 1926, opened the door to modern rocketry. Goddard's rocket reached an altitude of 41 feet while its flight lasted only 2.5 seconds. The Marshall design team's plan was to stay as close as possible to an authentic reconstruction of Goddard's rocket. The same propellants were used - liquid oxygen and gasoline - as available during Goddard's initial testing and firing. The team also tried to construct the replica using the original materials and design to the greatest extent possible. By purposely using less advanced techniques and materials than many that are available today, the team encountered numerous technical challenges in testing the functional hardware. There were no original blueprints or drawings, only photographs and notes. However, this faithful adherence to historical accuracy has allowed the team to experience many of the same challenges Goddard faced 77 years ago, and more fully appreciate the genius of this extraordinary man. In this photo, the replica is shown firing in the A-frame launch stand in near-flight configuration at MSFC's Test Area 116 during the American Institute of Aeronautics and Astronautics 39th Joint Propulsion Conference on July 23, 2003.

Liquid rocket booster study. Volume 2, book 4, appendices 6-8: Reports of Rocketdyne, Pratt and Whitney, and TRW

NASA Technical Reports Server (NTRS)

1988-01-01

For the pressure fed engines, detailed trade studies were conducted defining engine features such as thrust vector control methods, thrust chamber construction, etc. This was followed by engine design layouts and booster propulsion configuration layouts. For the pump fed engines parametric performance and weight data was generated for both O2/H2 and O2/RP-1 engines. Subsequent studies resulted in the selection of both LOX/RP-1 and O2/H2 propellants for the pump fed engines. More detailed analysis of the selected LOX/RP-1 and O2/H2 engines was conducted during the final phase of the study.

Early Rockets

NASA Image and Video Library

2004-04-15

The image depicts Redstone missile being erected. The Redstone ballistic missile was a high-accuracy, liquid-propelled, surface-to-surface missile developed by Army Ballistic Missile Agency, Redstone Arsenal, in Huntsville, Alabama, under the direction of Dr. von Braun. The Redstone engine was a modified and improved version of the Air Force's Navaho cruise missile engine of the late forties. The A-series, as this would be known, utilized a cylindrical combustion chamber as compared with the bulky, spherical V-2 chamber. By 1951, the Army was moving rapidly toward the design of the Redstone missile, and the production was begun in 1952. Redstone rockets became the "reliable workhorse" for America's early space program. As an example of the versatility, Redstone was utilized in the booster for Explorer 1, the first American satellite, with no major changes to the engine or missile

A History of Collapse Factor Modeling and Empirical Data for Cryogenic Propellant Tanks

NASA Technical Reports Server (NTRS)

deQuay, Laurence; Hodge, B. Keith

2010-01-01

One of the major technical problems associated with cryogenic liquid propellant systems used to supply rocket engines and their subassemblies and components is the phenomenon of propellant tank pressurant and ullage gas collapse. This collapse is mainly caused by heat transfer from ullage gas to tank walls and interfacing propellant, which are both at temperatures well below those of this gas. Mass transfer between ullage gas and cryogenic propellant can also occur and have minor to significant secondary effects that can increase or decrease ullage gas collapse. Pressurant gas is supplied into cryogenic propellant tanks in order to initially pressurize these tanks and then maintain required pressures as propellant is expelled from these tanks. The net effect of pressurant and ullage gas collapse is increased total mass and mass flow rate requirements of pressurant gases. For flight vehicles this leads to significant and undesirable weight penalties. For rocket engine component and subassembly ground test facilities this results in significantly increased facility hardware, construction, and operational costs. "Collapse Factor" is a parameter used to quantify the pressurant and ullage gas collapse. Accurate prediction of collapse factors, through analytical methods and modeling tools, and collection and evaluation of collapse factor data has evolved over the years since the start of space exploration programs in the 1950 s. Through the years, numerous documents have been published to preserve results of studies associated with the collapse factor phenomenon. This paper presents a summary and selected details of prior literature that document the aforementioned studies. Additionally other literature that present studies and results of heat and mass transfer processes, related to or providing important insights or analytical methods for the studies of collapse factor, are presented.

Prediction of explosive yield and other characteristics of liquid rocket propellant explosions

NASA Technical Reports Server (NTRS)

Farber, E. A.; Smith, J. H.; Watts, E. H.

1973-01-01

Work which has been done at the University of Florida in arriving at credible explosive yield values for liquid rocket propellants is presented. The results are based upon logical methods which have been well worked out theoretically and verified through experimental procedures. Three independent methods to predict explosive yield values for liquid rocket propellants are described. All three give the same end result even though they utilize different parameters and procedures. They are: (1) mathematical model; (2) seven chart approach; and (3) critical mass method. A brief description of the methods, how they were derived, how they were applied, and the results which they produced are given. The experimental work used to support and verify the above methods both in the laboratory and in the field with actually explosive mixtures are presented. The methods developed are used and their value demonstrated in analyzing real problems, among them the destruct system of the Saturn 5, and the early configurations of the space shuttle.

Ozone depletion caused by NO and H2O emissions from hydrazine-fueled rockets

NASA Astrophysics Data System (ADS)

Ross, M. N.; Danilin, M. Y.; Weisenstein, D. K.; Ko, M. K. W.

2004-11-01

Rockets using unsymmetrical dimethyl hydrazine (N(CH3)2NH2) and dinitrogen tetroxide (N2O4) propellants account for about one third of all stratospheric rocket engine emissions, comparable to the solid-fueled rocket emissions. We use plume and global atmosphere models to provide the first estimate of the local and global ozone depletion caused by NO and H2O emissions from the Proton rocket, the largest hydrazine-fueled launcher in use. NO and H2O emission indices are assumed to be 20 and 350 g/kg (propellant), respectively. Predicted maximum ozone loss in the plume of the Proton rocket is 21% at 44 km altitude. Plume ozone loss at 20 km equals 8% just after launch and steadily declines to 2% by model sunset. Predicted steady state global ozone loss from ten Proton launches annually is 1.2 × 10-4%, with nearly all of the loss due to the NO component of the emission. Normalized by stratospheric propellant consumption, the global ozone depletion efficiency of the Proton is approximately 66-90 times less than that of solid-fueled rockets. In situ Proton plume measurements are required to validate assumed emission indices and to assess the role of rocket emissions not considered in these calculations. Such future studies would help to establish a formalism to evaluate the relative ozone depletion caused by different rocket engines using different propellants.

NASA's Hydrogen Outpost: The Rocket Systems Area at Plum Brook Station

NASA Technical Reports Server (NTRS)

Arrighi, Robert S.

2016-01-01

"There was pretty much a general knowledge about hydrogen and its capabilities," recalled former researcher Robert Graham. "The question was, could you use it in a rocket engine? Do we have the technology to handle it? How will it cool? Will it produce so much heat release that we can't cool the engine? These were the questions that we had to address." The National Aeronautics and Space Administration's (NASA) Glenn Research Center, referred to historically as the Lewis Research Center, made a concerted effort to answer these and related questions in the 1950s and 1960s. The center played a critical role transforming hydrogen's theoretical potential into a flight-ready propellant. Since then NASA has utilized liquid hydrogen to send humans and robots to the Moon, propel dozens of spacecraft across the universe, orbit scores of satellite systems, and power 135 space shuttle flights. Rocket pioneers had recognized hydrogen's potential early on, but its extremely low boiling temperature and low density made it impracticable as a fuel. The Lewis laboratory first demonstrated that liquid hydrogen could be safely utilized in rocket and aircraft propulsion systems, then perfected techniques to store, pump, and cleanly burn the fuel, as well as use it to cool the engine. The Rocket Systems Area at Lewis's remote testing area, Plum Brook Station, played a little known, but important role in the center's hydrogen research efforts. This publication focuses on the activities at the Rocket Systems Area, but it also discusses hydrogen's role in NASA's space program and Lewis's overall hydrogen work. The Rocket Systems Area included nine physically modest test sites and three test stands dedicated to liquid-hydrogen-related research. In 1962 Cleveland Plain Dealer reporter Karl Abram claimed, "The rocket facility looks more like a petroleum refinery. Its test rigs sprout pipes, valves and tanks. During the night test runs, excess hydrogen is burned from special stacks in the best Oklahoma oil field tradition." Besides the Rocket Systems Area, Plum Brook Station also included a nuclear test reactor, a large vacuum tank, a hypersonic wind tunnel, and a full-scale upper-stage rocket stand. The Rocket Systems Area operated from 1961 until NASA shut down all of Plum Brook in 1974. The center reopened Plum Brook in the late 1980s and continues to use several test facilities. The Rocket Systems Area, however, was not restored. Today Plum Brook resembles a nature preserve more than an oil refinery. Lush fields and forests separate the large test facilities. Until recently, the abandoned Rocket Systems Area structures and equipment were visible amongst the greenery. These space-age ruins, particularly the three towers, stood as silent sentinels over the sparsely populated reservation. Few knew the story of these mysterious facilities when NASA removed them in the late 2000s.

Theoretical analysis of rotating two phase detonation in a rocket motor

NASA Technical Reports Server (NTRS)

Shen, I.; Adamson, T. C., Jr.

1973-01-01

Tangential mode, non-linear wave motion in a liquid propellant rocket engine is studied, using a two phase detonation wave as the reaction model. Because the detonation wave is followed immediately by expansion waves, due to the side relief in the axial direction, it is a Chapman-Jouguet wave. The strength of this wave, which may be characterized by the pressure ratio across the wave, as well as the wave speed and the local wave Mach number, are related to design parameters such as the contraction ratio, chamber speed of sound, chamber diameter, propellant injection density and velocity, and the specific heat ratio of the burned gases. In addition, the distribution of flow properties along the injector face can be computed. Numerical calculations show favorable comparison with experimental findings. Finally, the effects of drop size are discussed and a simple criterion is found to set the lower limit of validity of this strong wave analysis.

Combustion stability analysis of preburners in liquid propellant rocket engines during shutdown

NASA Technical Reports Server (NTRS)

Lim, Kair-Chuan; George, Paul E., II

1987-01-01

A linearized one-dimensional lumped-parameter model capable of predicting the occurrence of the low frequency combustion instability (chugging) experienced during preburner shutdown in the Space Shuttle Main Engines is discussed, and predictions are compared with NASA experimental results. Results from a parametric study of parameters including chamber pressure, fuel and oxygen temperatures, and the effective bulk modulus of the liquid oxidizer suggest that chugging is probably affected by conditions at shutdown through the fuel and oxidizer temperatures. It is suggested that chugging is initiated when the fuel, oxidizer, and helium temperature and flow rates pass into an unstable region, and that chugging may be terminated by decaying pressures.

Designing Liquid Rocket Engine Injectors for Performance, Stability, and Cost

NASA Technical Reports Server (NTRS)

Westra, Douglas G.; West, Jeffrey S.

2014-01-01

NASA is developing the Space Launch System (SLS) for crewed exploration missions beyond low Earth orbit. Marshall Space Flight Center (MSFC) is designing rocket engines for the SLS Advanced Booster (AB) concepts being developed to replace the Shuttle-derived solid rocket boosters. One AB concept uses large, Rocket-Propellant (RP)-fueled engines that pose significant design challenges. The injectors for these engines require high performance and stable operation while still meeting aggressive cost reduction goals for access to space. Historically, combustion stability problems have been a critical issue for such injector designs. Traditional, empirical injector design tools and methodologies, however, lack the ability to reliably predict complex injector dynamics that often lead to combustion stability. Reliance on these tools alone would likely result in an unaffordable test-fail-fix cycle for injector development. Recently at MSFC, a massively parallel computational fluid dynamics (CFD) program was successfully applied in the SLS AB injector design process. High-fidelity reacting flow simulations were conducted for both single-element and seven-element representations of the full-scale injector. Data from the CFD simulations was then used to significantly augment and improve the empirical design tools, resulting in a high-performance, stable injector design.

Research of the high performance low temperature vortex street flowmeter

NASA Astrophysics Data System (ADS)

Gao, Feng; Chen, Yang; Zhang, Zhen-peng; Geng, Wei-guo

2007-07-01

Flow measurement is the key method for R&D and operation monitoring of liquid rocket engine. Therefore, it is important to measure flux of low temperature liquid propellants for the liquid hydrogen/liquid oxygen or the liquid oxygen/kerosene rocket engine. Presently in China, the level meter and the turbine flowmeter are usually used in the experimentation of the liquid hydrogen/liquid oxygen rocket engine. The level meter can only scale average flux and the precision of the turbine flowmeter (the measuring wild point is 1.5%) can not be ensured due to the reason which there is not devices of low temperature real-time demarcation in China. Therefore, it is required to research the high performance low temperature flow measurement equipment and the vortex street flowmeter is selected because of its advantages. In the paper, some key techniques of low temperature vortex street flowmeter are researched from the design aspect. Firstly, the basic theoretical research of vortex street flowmeter includes signal detection method, shape of vortex producer and effects of dimension of vertex producer to vortex quality. Secondly, low temperature vortex street flowmeter adopts the method of piezoelectric components stress mode. As for the weakness of phase-change, lattice change and fragility for many piezoelectric materials in low temperature, it can not be fulfilled piezoelectric signal and mechanism performance under this condition. Some piezoelectric materials which can be used in low temperature are illustrated in the paper by lots of research in order for the farther research. The article places emphasis upon low temperature trait of piezoelectric materials, and the structure designs of signal detector and calculation of stress, electric charge quantity and heat transfer.

Small Launch Vehicle Concept Development for Affordable Multi-Stage Inline Configurations

NASA Technical Reports Server (NTRS)

Beers, Benjamin R.; Waters, Eric D.; Philips, Alan D.; Threet, Grady E., Jr.

2014-01-01

The Advanced Concepts Office at NASA's George C. Marshall Space Flight Center conducted a study of two configurations of a three stage, inline, liquid propellant small launch vehicle concept developed on the premise of maximizing affordability by targeting a specific payload capability range based on current industry demand. The initial configuration, NESC-1, employed liquid oxygen as the oxidizer and rocket propellant grade kerosene as the fuel in all three stages. The second and more heavily studied configuration, NESC-4, employed liquid oxygen and rocket propellant grade kerosene on the first and second stages and liquid oxygen and liquid methane fuel on the third stage. On both vehicles, sensitivity studies were first conducted on specific impulse and stage propellant mass fraction in order to baseline gear ratios and drive the focus of concept development. Subsequent sensitivity and trade studies on the NESC-4 configuration investigated potential impacts to affordability due to changes in gross liftoff weight and/or vehicle complexity. Results are discussed at a high level to understand the severity of certain sensitivities and how those trade studies conducted can either affect cost, performance or both.

KSC-2012-4454

NASA Image and Video Library

2012-08-14

CAPE CANAVERAL, Fla. – At NASA’s Kennedy Space Center in Florida, a crane is used to load a space shuttle solid rocket booster and an external fuel tank on trucks for transport to separate museums. The solid rocket boosters, or SRBs, will be displayed at the California Science Center in Los Angeles. The external tank soon will be transported for display at the Wings of Dreams Aviation Museum at Keystone Heights Airport between Gainesville and Jacksonville, Fla. The 149-foot SRBs together provided six million pounds of thrust. The external fuel tank contained over 500,000 gallons of liquid hydrogen and liquid oxygen propellant for the shuttle orbiters' three main engines. The work is part of Transition and Retirement of the space shuttle. For more information, visit http://www.nasa.gov/transition Photo credit: NASA/ Dimitri Gerondidakis

KSC-2012-4453

NASA Image and Video Library

2012-08-14

CAPE CANAVERAL, Fla. – At NASA’s Kennedy Space Center in Florida, a crane is used to load a space shuttle solid rocket booster and an external fuel tank on trucks for transport to separate museums. The solid rocket boosters, or SRBs, will be displayed at the California Science Center in Los Angeles. The external tank soon will be transported for display at the Wings of Dreams Aviation Museum at Keystone Heights Airport between Gainesville and Jacksonville, Fla. The 149-foot SRBs together provided six million pounds of thrust. The external fuel tank contained over 500,000 gallons of liquid hydrogen and liquid oxygen propellant for the shuttle orbiters' three main engines. The work is part of Transition and Retirement of the space shuttle. For more information, visit http://www.nasa.gov/transition Photo credit: NASA/ Dimitri Gerondidakis

KSC-2012-4448

NASA Image and Video Library

2012-08-14

CAPE CANAVERAL, Fla. – At NASA’s Kennedy Space Center in Florida, a crane is used to load a space shuttle solid rocket booster and an external fuel tank on to trucks for transport to separate museums. The solid rocket boosters, or SRBs, will be displayed at the California Science Center in Los Angeles. The external tank soon will be transported for display at the Wings of Dreams Aviation Museum at Keystone Heights Airport between Gainesville and Jacksonville, Fla. The 149-foot SRBs together provided six million pounds of thrust. The external fuel tank contained over 500,000 gallons of liquid hydrogen and liquid oxygen propellant for the shuttle orbiters' three main engines. The work is part of Transition and Retirement of the space shuttle. For more information, visit http://www.nasa.gov/transition Photo credit: NASA/ Dimitri Gerondidakis

KSC-2012-4449

NASA Image and Video Library

2012-08-14

CAPE CANAVERAL, Fla. – At NASA’s Kennedy Space Center in Florida, a crane is used to load a space shuttle solid rocket booster and an external fuel tank on trucks for transport to separate museums. The solid rocket boosters, or SRBs, will be displayed at the California Science Center in Los Angeles. The external tank soon will be transported for display at the Wings of Dreams Aviation Museum at Keystone Heights Airport between Gainesville and Jacksonville, Fla. The 149-foot SRBs together provided six million pounds of thrust. The external fuel tank contained over 500,000 gallons of liquid hydrogen and liquid oxygen propellant for the shuttle orbiters' three main engines. The work is part of Transition and Retirement of the space shuttle. For more information, visit http://www.nasa.gov/transition Photo credit: NASA/ Dimitri Gerondidakis

KSC-2012-4450

NASA Image and Video Library

2012-08-14

CAPE CANAVERAL, Fla. – At NASA’s Kennedy Space Center in Florida, a crane is used to load a space shuttle solid rocket booster and an external fuel tank on trucks for transport to separate museums. The solid rocket boosters, or SRBs, will be displayed at the California Science Center in Los Angeles. The external tank soon will be transported for display at the Wings of Dreams Aviation Museum at Keystone Heights Airport between Gainesville and Jacksonville, Fla. The 149-foot SRBs together provided six million pounds of thrust. The external fuel tank contained over 500,000 gallons of liquid hydrogen and liquid oxygen propellant for the shuttle orbiters' three main engines. The work is part of Transition and Retirement of the space shuttle. For more information, visit http://www.nasa.gov/transition Photo credit: NASA/ Dimitri Gerondidakis

KSC-2012-4452

NASA Image and Video Library

2012-08-14

CAPE CANAVERAL, Fla. – At NASA’s Kennedy Space Center in Florida, a crane is used to load a space shuttle solid rocket booster and an external fuel tank on trucks for transport to separate museums. The solid rocket boosters, or SRBs, will be displayed at the California Science Center in Los Angeles. The external tank soon will be transported for display at the Wings of Dreams Aviation Museum at Keystone Heights Airport between Gainesville and Jacksonville, Fla. The 149-foot SRBs together provided six million pounds of thrust. The external fuel tank contained over 500,000 gallons of liquid hydrogen and liquid oxygen propellant for the shuttle orbiters' three main engines. The work is part of Transition and Retirement of the space shuttle. For more information, visit http://www.nasa.gov/transition Photo credit: NASA/ Dimitri Gerondidakis

KSC-2012-4451

NASA Image and Video Library

2012-08-14

CAPE CANAVERAL, Fla. – At NASA’s Kennedy Space Center in Florida, a crane is used to load a space shuttle solid rocket booster and an external fuel tank on trucks for transport to separate museums. The solid rocket boosters, or SRBs, will be displayed at the California Science Center in Los Angeles. The external tank soon will be transported for display at the Wings of Dreams Aviation Museum at Keystone Heights Airport between Gainesville and Jacksonville, Fla. The 149-foot SRBs together provided six million pounds of thrust. The external fuel tank contained over 500,000 gallons of liquid hydrogen and liquid oxygen propellant for the shuttle orbiters' three main engines. The work is part of Transition and Retirement of the space shuttle. For more information, visit http://www.nasa.gov/transition Photo credit: NASA/ Dimitri Gerondidakis

KSC-2012-4444

NASA Image and Video Library

2012-08-14

CAPE CANAVERAL, Fla. – At NASA’s Kennedy Space Center in Florida, preparations are underway to load a twin set of space shuttle solid rocket boosters and an external fuel tank on trucks for transport to separate museums. The solid rocket boosters, or SRBs, will be displayed at the California Science Center in Los Angeles. The external tank soon will be transported for display at the Wings of Dreams Aviation Museum at Keystone Heights Airport between Gainesville and Jacksonville, Fla. The 149-foot SRBs together provided six million pounds of thrust. The external fuel tank contained over 500,000 gallons of liquid hydrogen and liquid oxygen propellant for the shuttle orbiters' three main engines. The work is part of Transition and Retirement of the space shuttle. For more information, visit http://www.nasa.gov/transition Photo credit: NASA/ Dimitri Gerondidakis

Experimental investigation of solid rocket motors for small sounding rockets

NASA Astrophysics Data System (ADS)

Suksila, Thada

2018-01-01

Experimentation and research of solid rocket motors are important subjects for aerospace engineering students. However, many institutes in Thailand rarely include experiments on solid rocket motors in research projects of aerospace engineering students, mainly because of the complexity of mixing the explosive propellants. This paper focuses on the design and construction of a solid rocket motor for total impulse in the class I-J that can be utilised as a small sounding rocket by researchers in the near future. Initially, the test stands intended for measuring the pressure in the combustion chamber and the thrust of the solid rocket motor were designed and constructed. The basic design of the propellant configuration was evaluated. Several formulas and ratios of solid propellants were compared for achieving the maximum thrust. The convenience of manufacturing and casting of the fabricated solid rocket motors were a critical consideration. The motor structural analysis such as the combustion chamber wall thickness was also discussed. Several types of nozzles were compared and evaluated for ensuring the maximum thrust of the solid rocket motors during the experiments. The theory of heat transfer analysis in the combustion chamber was discussed and compared with the experimental data.

NASA Engineer and Technician Instrument Zero Gravity Spheres

NASA Image and Video Library

1961-08-21

An engineer and technician at the National Aeronautics and Space Administration (NASA) Lewis Research Center install the instrumentation on spherical fuel tanks for an investigation of the behavior of liquids in microgravity. Lewis researchers were undertaking a broad effort to study the heat transfer properties of high energy propellants such as liquid hydrogen in microgravity. In the center’s 2.2-Second Drop Tower they investigated the wetting characteristics of liquid and the liquid-vapor configurations, and predicted the equilibrium state in microgravity conditions. Lewis was also conducting a series microgravity investigations which launched 9-inch diameter spherical dewars, seen here, on an Aerobee sounding rocket. A camera inside the rocket filmed the liquid hydrogen’s behavior during its 4 to 7 minutes of freefall. The researchers concluded, however, that they needed to extend the weightlessness period to obtain better results. So they designed an experiment to be launched on an Atlas missile that would provide 21 minutes of weightlessness. The experiment was flight qualified at Lewis. The 36-percent full liquid hydrogen stainless steel dewar was launched on the Atlas on February 25, 1964. The instrumentation measured temperature, pressure, vacuum, and liquid level. Temperature instrumentation indicated wall drying during the freefall. The resultant pressure-rise characteristics were similar to those used for the normal-gravity test.

Focused Rocket-Ejector RBCC Experiments

NASA Technical Reports Server (NTRS)

Santoro, Robert J.; Pal, Sibtosh

2003-01-01

This document reports the results of additional efforts for the Rocket Based Combined Cycle (RBCC) rocket-ejector mode research work carried out at the Perm State Propulsion Engineering Research Center in support of NASA s technology development efforts for enabling 3rd generation Reusable Launch Vehicles (RLV). The two tasks conducted under this program build on earlier NASA MSFC funded research program on rocket ejector investigations. The first task continued a systematic investigation of the improvements provided by a gaseous hydrogen (GHz)/oxygen (GO2) twin thruster RBCC rocket ejector system over a single rocket system. In a similar vein, the second task continued investigations into the performance of a hydrocarbon (liquid JP-7)/gaseous oxygen single thruster rocket-ejector system. To gain a systematic understanding of the rocket-ejector s internal fluid mechanic/combustion phenomena, experiments were conducted with both direct-connect and sea-level static diffusion and afterburning (DAB) configurations for a range of rocket operating conditions. For all experimental conditions, overall system performance was obtained through global measurements of wall static pressure profiles, heat flux profiles and engine thrust. For the GH2/GO2 propellant rocket ejector experiments, high frequency measurements of the pressure field within the system were also made to understand the unsteady behavior of the flowfield.

New Frontiers AO: Advanced Materials Bi-propellant Rocket (AMBR) Engine Information Summary

NASA Technical Reports Server (NTRS)

Liou, Larry C.

2008-01-01

The Advanced Material Bi-propellant Rocket (AMBR) engine is a high performance (I(sub sp)), higher thrust, radiation cooled, storable bi-propellant space engine of the same physical envelope as the High Performance Apogee Thruster (HiPAT(TradeMark)). To provide further information about the AMBR engine, this document provides details on performance, development, mission implementation, key spacecraft integration considerations, project participants and approach, contact information, system specifications, and a list of references. The In-Space Propulsion Technology (ISPT) project team at NASA Glenn Research Center (GRC) leads the technology development of the AMBR engine. Their NASA partners were Marshall Space Flight Center (MSFC) and Jet Propulsion Laboratory (JPL). Aerojet leads the industrial partners selected competitively for the technology development via the NASA Research Announcement (NRA) process.

Synthesis of Backfunctionalized Imidazolinium Salts and NHC Carbene Complexes

DTIC Science & Technology

2017-04-02

at American Chemical Society National Meeting; San Francisco, CA, USA (02 April 2017) Prepared in collaboration with California Institute of...AFRL – Tenant of Edwards AFB since late ‘50s – Full scale testing of the Atlas rockets (Gemini missions) – Initial testing of the F-1 engine (Apollo...Force has an interest in NHC carbene precursors for a variety applications – Ionic liquid propellants and additives – Ligands for Supercritical Chemical

High Speed Optical Diagnostics in a High Pressure, GOx/RP 2 Combustor

DTIC Science & Technology

2017-07-10

Similarly, the German Aerospace Center (DLR) at Lampoldshausen conducted research on LOX/LH2 propellants in the Combustion Chamber C experimental ...facility. This single element, optically accessible liquid rocket engine test article was designed to operate up to 1,450 psi, however most research was...significant result of this work was the acquisition of data at pressures up to 2400 psi. T 3 II. Experimental Facility This testing was

Theoretical performance of liquid hydrogen and liquid fluorine as a rocket propellant

NASA Technical Reports Server (NTRS)

Gordon, Sanford; Huff, Vearl N

1953-01-01

Theoretical values of performance parameters for liquid hydrogen and liquid fluorine as a rocket propellant were calculated on the assumption of equilibrium composition during the expansion process for a wide range of fuel-oxidant and expansion ratios. The parameters included were specific impulse, combustion-chamber temperature, nozzle-exit temperature, equilibrium composition, mean molecular weight, characteristic velocity, coefficient of thrust, ration of nozzle-exit area to throat area, specific heat at constant pressure, coefficient of viscosity, and coefficient of thermal conductivity. The maximum value of specific impulse was 364.6 pound-seconds per pound for a chamber pressure of 300 pounds per square inch absolute (20.41 atm) and an exit pressure of 1 atmosphere.

Theoretical performance of liquid ammonia and liquid fluorine as a rocket propellant

NASA Technical Reports Server (NTRS)

Gordon, Sanford; Huff, Vearl N

1953-01-01

Theoretical values of performance parameters for liquid ammonia and liquid fluorine as a rocket propellant were calculated on the assumption of equilibrium composition during the expansion process for a wide range of fuel-oxidant and expansion ratios. The parameters included were specific impulse, combustion chamber temperature, nozzle-exit temperature, equilibrium composition, mean molecular weight, characteristic velocity, coefficient of thrust, ratio of nozzle-exit area to throat area, specific heat at constant pressure, coefficient of viscosity, and coefficient of thermal conductivity. The maximum value of specific impulse was 311.5 pound-seconds per pound for a chamber pressure of 300 pounds per square inch absolute (20.41 atm) and an exit pressure of 1 atmosphere.

Early Rockets

NASA Image and Video Library

2004-04-15

By the end of the 19th Century, a Russian theorist, Konstantian Tsiolkovsky, was examining the fundamental scientific theories behind rocketry. He made some pioneering studies in liquid chemical rocket concepts and recommended liquid oxygen and liquid hydrogen as the optimum propellants. In the 1920's, Tsiolkovsky analyzed and mathematically formulated the technique for staged vehicles to reach escape velocities from Earth.

Liquid Rocket Propulsion for Atmospheric Flight in the Proposed ARES Mars Scout Mission

NASA Technical Reports Server (NTRS)

Kuhl, Christopher A.; Wright, Henry S.; Hunter, Craig A.; Guernsey, Carl S.; Colozza, Anthony J.

2004-01-01

Flying above the Mars Southern Highlands, an airplane will traverse over the terrain of Mars while conducting unique science measurements of the atmosphere, surface, and interior. This paper describes an overview of the ARES (Aerial Regional-scale Environmental Survey) mission with an emphasis on airplane propulsion needs. The process for selecting a propulsion system for the ARES airplane is also included. Details of the propulsion system, including system schematics, hardware and performance are provided. The airplane has a 6.25 m wingspan with a total mass of 149 kg and is propelled by a bi-propellant liquid rocket system capable of carrying roughly 48 kg of MMH/MON3 propellant.

Experimental Simulation of Turbine-Exhaust Oxygen Recovery

NASA Technical Reports Server (NTRS)

Clark, Jim A.; Branch, Ryan W.

2004-01-01

In some liquid-propellant rocket engines, the liquid-oxygen boost pump is driven by a turbine that is powered by high-pressure gaseous oxygen. Once it exits the turbine, this gaseous oxygen can be salvaged by injecting it into the subcooled liquid oxygen exiting the boost pump. If the main LOX pump is to function correctly under these circumstances, complete condensation of the gaseous oxygen must quickly follow its injection into the boost-pump discharge. The current investigation uses steam and water in a simple rig that allows the condensation process to be visualized and quantified. This paper offers dimensionless-parameter correlations of the data and trends observed.

Direct electrical arc ignition of hybrid rocket motors

NASA Astrophysics Data System (ADS)

Judson, Michael I., Jr.

Hybrid rockets motors provide distinct safety advantages when compared to traditional liquid or solid propellant systems, due to the inherent stability and relative inertness of the propellants prior to established combustion. As a result of this inherent propellant stability, hybrid motors have historically proven difficult to ignite. State of the art hybrid igniter designs continue to require solid or liquid reactants distinct from the main propellants. These ignition methods however, reintroduce to the hybrid propulsion system the safety and complexity disadvantages associated with traditional liquid or solid propellants. The results of this study demonstrate the feasibility of a novel direct electrostatic arc ignition method for hybrid motors. A series of small prototype stand-alone thrusters demonstrating this technology were successfully designed and tested using Acrylonitrile Butadiene Styrene (ABS) plastic and Gaseous Oxygen (GOX) as propellants. Measurements of input voltage and current demonstrated that arc-ignition will occur using as little as 10 watts peak power and less than 5 joules total energy. The motor developed for the stand-alone small thruster was adapted as a gas generator to ignite a medium-scale hybrid rocket motor using nitrous oxide /and HTPB as propellants. Multiple consecutive ignitions were performed. A large data set as well as a collection of development `lessons learned' were compiled to guide future development and research. Since the completion of this original groundwork research, the concept has been developed into a reliable, operational igniter system for a 75mm hybrid motor using both gaseous oxygen and liquid nitrous oxide as oxidizers. A development map of the direct spark ignition concept is presented showing the flow of key lessons learned between this original work and later follow on development.

Experimental Evaluation of a Subscale Gaseous Hydrogen/gaseous Oxygen Coaxial Rocket Injector

NASA Technical Reports Server (NTRS)

Smith, Timothy D.; Klem, Mark D.; Breisacher, Kevin J.; Farhangi, Shahram; Sutton, Robert

2002-01-01

The next generation reusable launch vehicle may utilize a Full-Flow Stage Combustion (FFSC) rocket engine cycle. One of the key technologies required is the development of an injector that uses gaseous oxygen and gaseous hydrogen as propellants. Gas-gas propellant injection provides an engine with increased stability margin over a range of throttle set points. This paper summarizes an injector design and testing effort that evaluated a coaxial rocket injector for use with gaseous oxygen and gaseous hydrogen propellants. A total of 19 hot-fire tests were conducted up to a chamber pressure of 1030 psia, over a range of 3.3 to 6.7 for injector element mixture ratio. Post-test condition of the hardware was also used to assess injector face cooling. Results show that high combustion performance levels could be achieved with gas-gas propellants and there were no problems with excessive face heating for the conditions tested.

Solid propellant processing factor in rocket motor design

NASA Technical Reports Server (NTRS)

1971-01-01

The ways are described by which propellant processing is affected by choices made in designing rocket engines. Tradeoff studies, design proof or scaleup studies, and special design features are presented that are required to obtain high product quality, and optimum processing costs. Processing is considered to include the operational steps involved with the lining and preparation of the motor case for the grain; the procurement of propellant raw materials; and propellant mixing, casting or extrusion, curing, machining, and finishing. The design criteria, recommended practices, and propellant formulations are included.

Nonlinear Longitudinal Mode Instability in Liquid Propellant Rocket Engine Preburners

NASA Technical Reports Server (NTRS)

Sims, J. D. (Technical Monitor); Flandro, Gary A.; Majdalani, Joseph; Sims, Joseph D.

2004-01-01

Nonlinear pressure oscillations have been observed in liquid propellant rocket instability preburner devices. Unlike the familiar transverse mode instabilities that characterize primary combustion chambers, these oscillations appear as longitudinal gas motions with frequencies that are typical of the chamber axial acoustic modes. In several respects, the phenomenon is similar to longitudinal mode combustion instability appearing in low-smoke solid propellant motors. An important feature is evidence of steep-fronted wave motions with very high amplitude. Clearly, gas motions of this type threaten the mechanical integrity of associated engine components and create unacceptably high vibration levels. This paper focuses on development of the analytical tools needed to predict, diagnose, and correct instabilities of this type. For this purpose, mechanisms that lead to steep-fronted, high-amplitude pressure waves are described in detail. It is shown that such gas motions are the outcome of the natural steepening process in which initially low amplitude standing acoustic waves grow into shock-like disturbances. The energy source that promotes this behavior is a combination of unsteady combustion energy release and interactions with the quasi-steady mean chamber flow. Since shock waves characterize the gas motions, detonation-like mechanisms may well control the unsteady combustion processes. When the energy gains exceed the losses (represented mainly by nozzle and viscous damping), the waves can rapidly grow to a finite amplitude limit cycle. Analytical tools are described that allow the prediction of the limit cycle amplitude and show the dependence of this wave amplitude on the system geometry and other design parameters. This information can be used to guide corrective procedures that mitigate or eliminate the oscillations.

The liquid rocket booster as an element of the U.S. national space transportation system

NASA Astrophysics Data System (ADS)

Bialla, Paul H.; Simon, Michael C.

Liquid rocket boosters (LRBs) were first considered for the U.S. Space Transportation System (STS) during the early conceptual phases of the Space Shuttle program. However, solid rocket boosters (SRBs) were ultimately selected for the STS, primarily due to near-term economics. Liquid rocket boosters are once again being considered as a possible future upgrade to the Shuttle. This paper addresses the findings of these studies to date, with emphasis on the feasibility, benefits, and implementation strategy for a LRB program. The principal issue relating to LRB feasibility is their ability to be integrated into the STS with minimal vehicle and facility impacts. Booster size has been shown to have a significant influence on compatibility with the STS. The physical dimensions of the Orbiter and STS support facilities place an inherent limitation on the size of any booster to be used with this system. In addition, excessively large diameter boosters can cause increased airloads to be induced on the Orbiter wings, requiring modification of STS launch trajectory and possible performance losses. However, trajectory and performance analyses have indicated that LRBs can be designed within these sizing constraints and still have sufficient performance to meet Space Shuttle mission requirements. In fact, several configurations have been developed to meet a design goal of providing a 20,000 lb performance improvement to low Earth-orbit (LEO), as compared with current SRBs. Several major system trade studies have been performed to establish a baseline design which is most compatible with the existing Space Transportation System. These trades include propellant selection (storable, hydrogen-oxygen, hydrocarbon-oxygen, and advanced propellants); pump-fed vs pressure-fed propellant feed system design; engine selection (Space Shuttle Main Engine, Titan LR-87, and advanced new engines); number of engines per booster; and reusability vs expendability. In general, it was determined through these trade studies that several options exist for designing a LRB that can be integrated into the STS with manageable impacts on STS facilities and operational procedures. While LRBs offer a potential 40% improvement in Shuttle performance, their most significant benefit is the potential improvements they offer in the area of Shuttle safety. This begins during ground handling operations, where LRBs eliminate the need for large quantities of hazardous solid propellants to be emplaced in the Kennedy Space Center Vehicle Assembly Building. In the pre-launch phase, all LRB engines can be ignited on the launch pad and verified prior to release of the STS. During flight, LRB engines can be shut down on command should the need arise. Further, missions could be aborted safely during the boost phase—an option not available with SRBs. A related benefit of LRBs is their ability to accomplish a mission even if one engine fails, assuming the LRB is designed with sufficient performance margin. An implementation plan has been developed which indicates that LRBs can be operational by 1997. The attractive features of the LRB have prompted NASA to include this booster as a principal element of the agency's long range plan for enhancing STS capabilities through an evolutionary program of block changes. The implementation of LRBs offers an attractive option for developing a safer, more reliable, and better performing STS.

Propellant vaporization as a criterion for rocket-engine design : experimental effect of fuel temperature on liquid-oxygen - heptane performance

NASA Technical Reports Server (NTRS)

Heidmann, M F

1957-01-01

Characteristic exhaust velocity of a 200-pound-thrust rocket engine was evaluated for fuel temperatures of -90 degrees, and 200 degrees f with a spray formed by two impinging heptane jets reacting in a highly atomized oxygen atmosphere. Tests covered a range of mixture ratios and chamber lengths. The characteristic exhaust-velocity efficiency increased 2 percent for a 290 degree f increase in fuel temperature. This increase in performance can be compared with that obtained by increasing chamber length by about 1/2 inch. The result agrees with the fuel-temperature effect predicted from an analysis based on droplet evaporation theory. Mixture ratio markedly affected characteristic exhaust velocity efficiency, but total flow rate and fuel temperature did not.

The prediction of three-dimensional liquid-propellant rocket nozzle admittances

NASA Technical Reports Server (NTRS)

Bell, W. A.; Zinn, B. T.

1973-01-01

Crocco's three-dimensional nozzle admittance theory is extended to be applicable when the amplitudes of the combustor and nozzle oscillations increase or decrease with time. An analytical procedure and a computer program for determining nozzle admittance values from the extended theory are presented and used to compute the admittances of a family of liquid-propellant rocket nozzles. The calculated results indicate that the nozzle geometry entrance Mach number and temporal decay coefficient significantly affect the nozzle admittance values. The theoretical predictions are shown to be in good agreement with available experimental data.

Studies on an aerial propellant transfer space plane (APTSP)

NASA Astrophysics Data System (ADS)

Jayan, N.; Biju Kumar, K. S.; Gupta, Anish Kumar; Kashyap, Akhilesh Kumar; Venkatraman, Kartik; Mathew, Joseph; Mukunda, H. S.

2004-04-01

This paper presents a study of a fully reusable earth-to-orbit launch vehicle concept with horizontal take-off and landing, employing a turbojet engine for low speed, and a rocket for high-speed acceleration and space operations. This concept uses existing technology to the maximum possible extent, thereby reducing development time, cost and effort. It uses the experience in aerial filling of military aircrafts for propellant filling at an altitude of 13 km at a flight speed of M=0.85. Aerial filling of propellant reduces the take-off weight significantly thereby minimizing the structural weight of the vehicle. The vehicle takes off horizontally and uses turbojet engines till the end of the propellant filling operation. The rocket engines provide thrust for the next phase till the injection of a satellite at LEO. A sensitivity analysis of the mission with respect to rocket engine specific impulse and overall vehicle structural factor is also presented in this paper. A conceptual design of space plane with a payload capability of 10 ton to LEO is carried out. The study shows that the realization of an aerial propellant transfer space plane is possible with limited development of new technology thus reducing the demands on the finances required for achieving the objectives.

Liquid Rocket Engine Testing Overview

NASA Technical Reports Server (NTRS)

Rahman, Shamim

2005-01-01

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.

Advanced Space Transportation Program (ASTP)

NASA Image and Video Library

2006-12-05

The NASA developed Ares rockets, named for the Greek god associated with Mars, will return humans to the moon and later take them to Mars and other destinations. This is an illustration of the Ares V with call outs. The Ares V is a heavy lift launch vehicle that will use five RS-68 liquid oxygen/liquid hydrogen engines mounted below a larger version of the space shuttle external tank, and two five-segment solid propellant rocket boosters for the first stage. The upper stage will use the same J-2X engine as the Ares I and past Apollo vehicles. The Ares V can lift more than 286,000 pounds to low Earth orbit and stands approximately 360 feet tall. This versatile system will be used to carry cargo and the components into orbit needed to go to the moon and later to Mars. Ares V is subject to configuration changes before it is actually launched. This illustration reflects the latest configuration as of January 2007.

Illustration of Ares V Launch Vehicle With Call Outs

NASA Technical Reports Server (NTRS)

2006-01-01

The NASA developed Ares rockets, named for the Greek god associated with Mars, will return humans to the moon and later take them to Mars and other destinations. This is an illustration of the Ares V with call outs. The Ares V is a heavy lift launch vehicle that will use five RS-68 liquid oxygen/liquid hydrogen engines mounted below a larger version of the space shuttle external tank, and two five-segment solid propellant rocket boosters for the first stage. The upper stage will use the same J-2X engine as the Ares I and past Apollo vehicles. The Ares V can lift more than 286,000 pounds to low Earth orbit and stands approximately 360 feet tall. This versatile system will be used to carry cargo and the components into orbit needed to go to the moon and later to Mars. Ares V is subject to configuration changes before it is actually launched. This illustration reflects the latest configuration as of January 2007.

Model of the Ares V Launch System

NASA Technical Reports Server (NTRS)

2006-01-01

This is a studio photograph of a model of the Ares V rocket. Named for the Greek god associated with Mars, Ares vehicles will return humans to the moon and later take them to Mars and other destinations. The Ares V is a heavy lift launch vehicle that will use five RS-68 liquid oxygen/liquid hydrogen engines mounted below a larger version of the space shuttle external tank, and two five-segment solid propellant rocket boosters for the first stage. The upper stage will use the same J-2X engine as the Ares I. The Ares V can lift more than 286,000 pounds to low Earth orbit and stands approximately 360 feet tall. This versatile system will be used to carry cargo and the components into orbit needed to go to the moon and later to Mars, while the Crew will be carried by the Ares I. Ares V is subject to configuration changes before it is actually launched. This illustration reflects the latest configuration as of September 2006.

Predicting performance of axial pump inducer of LOX booster turbo-pump of staged combustion cycle based rocket engine using CFD

NASA Astrophysics Data System (ADS)

Mishra, Arpit; Ghosh, Parthasarathi

2015-12-01

For low cost, high thrust, space missions with high specific impulse and high reliability, inert weight needs to be minimized and thereby increasing the delivered payload. Turbopump feed system for a liquid propellant rocket engine (LPRE) has the highest power to weight ratio. Turbopumps are primarily equipped with an axial flow inducer to achieve the high angular velocity and low suction pressure in combination with increased system reliability. Performance of the turbopump strongly depends on the performance of the inducer. Thus, for designing a LPRE turbopump, demands optimization of the inducer geometry based on the performance of different off-design operating regimes. In this paper, steady-state CFD analysis of the inducer of a liquid oxygen (LOX) axial pump used as a booster pump for an oxygen rich staged combustion cycle rocket engine has been presented using ANSYS® CFX. Attempts have been made to obtain the performance characteristic curves for the LOX pump inducer. The formalism has been used to predict the performance of the inducer for the throttling range varying from 80% to 113% of nominal thrust and for the different rotational velocities from 4500 to 7500 rpm. The results have been analysed to determine the region of cavitation inception for different inlet pressure.

A study of performance and cost improvement potential of the 120 inch (3.05 m) diameter solid rocket motor. Volume 1: Summary report

NASA Technical Reports Server (NTRS)

Backlund, S. J.; Rossen, J. N.

1971-01-01

A parametric study of ballistic modifications to the 120 inch diameter solid propellant rocket engine which forms part of the Air Force Titan 3 system is presented. 576 separate designs were defined and 24 were selected for detailed analysis. Detailed design descriptions, ballistic performance, and mass property data were prepared for each design. It was determined that a relatively simple change in design parameters could provide a wide range of solid propellant rocket engine ballistic characteristics for future launch vehicle applications.

Around Marshall

NASA Image and Video Library

2003-07-23

In honor of the Centernial of Flight Celebration and commissioned by the American Institute of Aeronautics and Astronautics (AIAA), a team of engineers from Marshall Space Flight Center (MSFC) built a replica of the first liquid-fueled rocket. The original rocket, designed and built by rocket engineering pioneer Robert H. Goddard in 1926, opened the door to modern rocketry. Goddard's rocket reached an altitude of 41 feet while its flight lasted only 2.5 seconds. The Marshall design team's plan was to stay as close as possible to an authentic reconstruction of Goddard's rocket. The same propellants were used - liquid oxygen and gasoline - as available during Goddard's initial testing and firing. The team also tried to construct the replica using the original materials and design to the greatest extent possible. By purposely using less advanced techniques and materials than many that are available today, the team encountered numerous technical challenges in testing the functional hardware. There were no original blueprints or drawings, only photographs and notes. However, this faithful adherence to historical accuracy has allowed the team to experience many of the same challenges Goddard faced 77 years ago, and more fully appreciate the genius of this extraordinary man. In this photo, the replica is shown firing in the A-frame launch stand in near-flight configuration at MSFC's Test Area 116 during the American Institute of Aeronautics and Astronautics 39th Joint Propulsion Conference on July 23, 2003.

Robust Exploration and Commercial Missions to the Moon Using NTR LANTR Propulsion and Lunar-Derived Propellants

NASA Technical Reports Server (NTRS)

Borowski, Stanley K.; Ryan, Stephen W.; Burke, Laura M.; McCurdy, David R.; Fittje, James E.; Joyner, Claude R.

2017-01-01

The nuclear thermal rocket (NTR) has frequently been identified as a key space asset required for the human exploration of Mars. This proven technology can also provide the affordable access through cislunar space necessary for commercial development and sustained human presence on the Moon. In his post-Apollo Integrated Space Program Plan (1970-1990), Wernher von Braun, proposed a reusable nuclear thermal propulsion stage (NTPS) to deliver cargo and crew to the Moon to establish a lunar base before undertaking human missions to Mars. The NTR option was selected by von Braun because it was a demonstrated technology capable of generating both high thrust and high specific impulse (Isp 900 s) twice that of todays best chemical rockets. In NASAs Mars Design Reference Architecture (DRA) 5.0 study, the crewed Mars transfer vehicle used three 25 klbf Pewee engines the smallest and highest performing engine tested in the Rover program along with graphite composite fuel. Smaller, lunar transfer vehicles consisting of a NTPS using three approximately 16.5 klbf Small Nuclear Rocket Engines (SNREs), an in-line propellant tank, plus the payload can enable a variety of reusable lunar missions. These include cargo delivery and crewed lunar landing missions. Even weeklong tourism missions carrying passengers into lunar orbit for a day of sightseeing and picture taking are possible. The NTR can play an important role in the next phase of lunar exploration and development by providing an affordable in-space lunar transportation system (LTS) that can allow initial outposts to evolve into settlements supported by a variety of commercial activities such as in-situ propellant production used to supply strategically located propellant depots and transportation nodes. The utilization of iron-rich volcanic glass or lunar polar ice (LPI) deposits (each estimated at billions of metric tons) for propellant production can significantly reduce the launch mass requirements from Earth and can enable reusable, surface-based lunar landing vehicles (LLVs) using liquid oxygen/hydrogen (LOX/LH2) chemical rocket engines. Afterwards, LOX/LH2 propellant depots can be established in lunar equatorial and polar orbits to supply the LTS. At this point a modified version of the conventional NTR called the LOX-augmented NTR, or LANTR would be introduced into the LTS allowing bipropellant operation and leveraging the mission benefits of refueling with lunar-derived propellants for Earth return. The bipropellant LANTR engine utilizes the large divergent section of its nozzle as an afterburner into which oxygen is injected and supersonically combusted with nuclear preheated hydrogen emerging from the engines choked sonic throat essentially scramjet propulsion in reverse. By varying the oxygen-to-hydrogen mixture ratio, LANTR engines can operate over a range of thrust and Isp values while the reactor core power level remains relatively constant. Eventually, a LANTR-based LTS can enable a rapid commuter shuttle with one-way trip times to and from the Moon ranging from 36 to 24 hours. Even if only 1 of the extracted propellant from identified volcanic glass and polar ice deposits were available for use in lunar orbit, such a supply could support daily commuter flights to the Moon for many thousands of years! An evolutionary mission architecture is outlined and a variety of lunar missions and transfer vehicle designs are examined, along with the increasing demands on propellant production as mission complexity increases. A comparison of vehicle features and engine operating characteristics, for both NTR and LANTR engines, is also provided along with a brief discussion on the propellant production issues associated with using volcanic glass and LPI as source material.

A-3 Test Stand work

NASA Image and Video Library

2011-07-29

Rocket engine propellant tanks and cell dome top the A-3 Test Stand under construction at Stennis Space Center. The stand will test next-generation rocket engines that could carry humans beyond low-Earth orbit into deep space once more.

A hybrid rocket engine design for simple low cost sounding rocket use

NASA Astrophysics Data System (ADS)

Grubelich, Mark; Rowland, John; Reese, Larry

1993-06-01

Preliminary test results on a nitrous oxide/HTPB hybrid rocket engine suitable for powering a small sounding rocket to altitudes of 50-100 K/ft are presented. It is concluded that the advantage of the N2O hybrid engine over conventional solid propellant rocket motors is the ability to obtain long burn times with core burning geometries due to the low regression rate of the fuel. Long burn times make it possible to reduce terminal velocity to minimize air drag losses.

Liquid Rocket Engine Testing

DTIC Science & Technology

2016-10-21

Briefing Charts 3. DATES COVERED (From - To) 17 October 2016 – 26 October 2016 4. TITLE AND SUBTITLE Liquid Rocket Engine Testing 5a. CONTRACT NUMBER...298 (Rev. 8-98) Prescribed by ANSI Std. 239.18 Liquid Rocket Engine Testing SFTE Symposium 21 October 2016 Jake Robertson, Capt USAF AFRL...Distribution Unlimited. PA Clearance 16493 Liquid Rocket Engine Testing • Engines and their components are extensively static-tested in development • This

Early Rockets

NASA Image and Video Library

1958-05-15

Redstone missile No. 1002 on the launch pad at Cape Canaveral, Florida, on May 16, 1958. The Redstone ballistic missile was a high-accuracy, liquid-propelled, surface-to-surface missile developed by the Army Ballistic Missile Agency, Redstone Arsenal, in Huntsville, Alabama, under the direction of Dr. von Braun. The Redstone engine was a modified and improved version of the Air Force's Navaho cruise missile engine of the late forties. The A-series, as this would be known, utilized a cylindrical combustion chamber as compared with the bulky, spherical V-2 chamber. By 1951, the Army was moving rapidly toward the design of the Redstone missile, and production was begun in 1952. Redstone rockets became the "reliable workhorse" for America's early space program. As an example of the versatility, Redstone was utilized in the booster for Explorer 1, the first American satellite, with no major changes to the engine or missile

Liquid Oxygen/Liquid Methane Test Results of the RS-18 Lunar Ascent Engine at Simulated Altitude Conditions at NASA White Sands Test Facility

NASA Technical Reports Server (NTRS)

Melcher, John C., IV; Allred, Jennifer K.

2009-01-01

Tests were conducted with the RS-18 rocket engine using liquid oxygen (LO2) and liquid methane (LCH4) propellants under simulated altitude conditions at NASA Johnson Space Center White Sands Test Facility (WSTF). This project is part of NASA's Propulsion and Cryogenics Advanced Development (PCAD) project. "Green" propellants, such as LO2/LCH4, offer savings in both performance and safety over equivalently sized hypergolic propulsion systems in spacecraft applications such as ascent engines or service module engines. Altitude simulation was achieved using the WSTF Large Altitude Simulation System, which provided altitude conditions equivalent up to 122,000 ft (37 km). For specific impulse calculations, engine thrust and propellant mass flow rates were measured. LO2 flow ranged from 5.9 - 9.5 lbm/sec (2.7 - 4.3 kg/sec), and LCH4 flow varied from 3.0 - 4.4 lbm/sec (1.4 - 2.0 kg/sec) during the RS-18 hot-fire test series. Propellant flow rate was measured using a coriolis mass-flow meter and compared with a serial turbine-style flow meter. Results showed a significant performance measurement difference during ignition startup due to two-phase flow effects. Subsequent cold-flow testing demonstrated that the propellant manifolds must be adequately flushed in order for the coriolis flow meters to give accurate data. The coriolis flow meters were later shown to provide accurate steady-state data, but the turbine flow meter data should be used in transient phases of operation. Thrust was measured using three load cells in parallel, which also provides the capability to calculate thrust vector alignment. Ignition was demonstrated using a gaseous oxygen/methane spark torch igniter. Test objectives for the RS-18 project are 1) conduct a shakedown of the test stand for LO2/methane lunar ascent engines, 2) obtain vacuum ignition data for the torch and pyrotechnic igniters, and 3) obtain nozzle kinetics data to anchor two-dimensional kinetics codes. All of these objectives were met with the RS-18 data and additional testing data from subsequent LO2/methane test programs in 2009 which included the first simulated-altitude pyrotechnic ignition demonstration of LO2/methane.

Status of Liquid Oxygen/Liquid Methane Injector Study for a Mars Ascent Engine

NASA Technical Reports Server (NTRS)

Trinh, Huu Ogyic; Cramer, John M.

1998-01-01

Preliminary mission studies for human exploration of Mars have been performed at Marshall Space Flight Center (MSFC). These studies indicate that for non-toxic chemical rockets only a cryogenic propulsion system would provide high enough performance to be considered for a Mars ascent vehicle. Although the mission is possible with Earth-supplied propellants for this vehicle, utilization of in-situ propellants is highly attractive. This option would significantly reduce the overall mass of the return vehicle. Consequently, the cost of the mission would be greatly reduced because the number and size of the Earth launch vehicle(s) needed for the mission decrease. NASA/Johnson Space Center has initiated several concept studies (2) of in-situ propellant production plants. Liquid oxygen (LOX) is the primary candidate for an in-situ oxidizer. In-situ fuel candidates include methane (CH4), ethylene (C2H4), and methanol (CH3OH). MSFC initiated a technology development program for a cryogenic propulsion system for the Mars human exploration mission in 1998. One part of this technology program is the effort described here: an evaluation of propellant injection concepts for a LOX/liquid methane Mars Ascent Engine (MAE) with an emphasis on light-weight, high efficiency, reliability, and thermal compatibility. In addition to the main objective, hot-fire tests of the subject injectors will be used to test other key technologies including light-weight combustion chamber materials and advanced ignition concepts. This state-of-the-art technology will then be applied to the development of a cryogenic propulsion system that will meet the requirements of the planned Mars sample return (MSR) mission. The current baseline propulsion system for the MSR mission uses a storable propellant combination [monomethyl hydrazine/mixed oxides of nitrogen-25(MMH/MON-25)]. However, a mission option that incorporates in-situ propellant production and utilization for the ascent stage is being carefully considered as a subscale precursor to a future human mission to Mars.

Theoretical Performance of Liquid Hydrogen with Liquid Oxygen as a Rocket Propellant

NASA Technical Reports Server (NTRS)

Gordon, Sanford; McBride, Bonnie J.

1959-01-01

Theoretical rocket performance for both equilibrium and frozen composition during expansion was calculated for the propellant combination liquid hydrogen and liquid oxygen at four chamber pressures (60, 150, 300, and 600 lb/sq in. abs) and a wide range of pressure ratios (1 to 4000) and oxidant-fuel ratios (1.190 to 39.683). Data are given to estimate performance parameters at chamber pressures other than those for which data are tabulated. The parameters included are specific impulse, specific impulse in vacuum, combustion-chamber temperature, nozzle-exit temperature, molecular weight, molecular-weight derivatives, characteristic velocity, coefficient of thrust, ratio of nozzle-exit area to throat area, specific heat at constant pressure, isentropic exponent, viscosity, thermal conductivity, Mach number, and equilibrium gas compositions.

KSC-2012-4442

NASA Image and Video Library

2012-08-14

CAPE CANAVERAL, Fla. – At NASA’s Kennedy Space Center in Florida, a crane is used to load a twin set of space shuttle solid rocket boosters and an external fuel tank on trucks for transport to separate museums. The solid rocket boosters, or SRBs, will be displayed at the California Science Center in Los Angeles. The external tank soon will be transported for display at the Wings of Dreams Aviation Museum at Keystone Heights Airport between Gainesville and Jacksonville, Fla. The 149-foot SRBs together provided six million pounds of thrust. The external fuel tank contained over 500,000 gallons of liquid hydrogen and liquid oxygen propellant for the shuttle orbiters' three main engines. The work is part of Transition and Retirement of the space shuttle. For more information, visit http://www.nasa.gov/transition Photo credit: NASA/ Dimitri Gerondidakis

KSC-2012-4438

NASA Image and Video Library

2012-08-14

CAPE CANAVERAL, Fla. – At NASA’s Kennedy Space Center in Florida, a crane is used to load a twin set of space shuttle solid rocket boosters and an external fuel tank on trucks for transport to separate museums. The solid rocket boosters, or SRBs, will be displayed at the California Science Center in Los Angeles. The external tank soon will be transported for display at the Wings of Dreams Aviation Museum at Keystone Heights Airport between Gainesville and Jacksonville, Fla. The 149-foot SRBs together provided six million pounds of thrust. The external fuel tank contained over 500,000 gallons of liquid hydrogen and liquid oxygen propellant for the shuttle orbiters' three main engines. The work is part of Transition and Retirement of the space shuttle. For more information, visit http://www.nasa.gov/transition Photo credit: NASA/ Dimitri Gerondidakis

JANNAF 35th Combustion Subcommittee Meeting. Volume 1

NASA Technical Reports Server (NTRS)

Fry, Ronald S. (Editor); Gannaway, Mary T. (Editor); Rognan, Melanie (Editor)

1998-01-01

Volume 1, the first of two volumes is a compilation of 63 unclassified/unlimited distribution technical papers presented at the 35th meeting of the Joint Army-Navy-NASA-Air Force (JANNAF) Combustion Subcommittee (CS) held jointly with the 17th Propulsion Systems Hazards Subcommittee (PSHS) and Airbreathing Propulsion Subcommittee (APS). The meeting was held on 7-11 December 1998 at Raytheon Systems Company and the Marriott Hotel, Tucson, AZ. Topics covered include solid gun propellant processing, ignition and combustion, charge concepts, barrel erosion and flash, gun interior ballistics, kinetics and molecular modeling, ETC gun modeling, simulation and diagnostics, and liquid gun propellant combustion; solid rocket motor propellant combustion, combustion instability fundamentals, motor instability, and measurement techniques; and liquid and hybrid rocket combustion.

Advanced Space Transportation Program (ASTP)

NASA Image and Video Library

2006-09-09

Named for the Greek god associated with Mars, the NASA developed Ares launch vehicles will return humans to the moon and later take them to Mars and other destinations. In this early illustration, the vehicle depicted on the left is the Ares I. Ares I is an inline, two-stage rocket configuration topped by the Orion crew vehicle and its launch abort system. In addition to its primary mission of carrying four to six member crews to Earth orbit, Ares I may also use its 25-ton payload capacity to deliver resources and supplies to the International Space Station (ISS), or to "park" payloads in orbit for retrieval by other spacecraft bound for the moon or other destinations. The Ares I employs a single five-segment solid rocket booster, a derivative of the space shuttle solid rocket booster, for the first stage. A liquid oxygen/liquid hydrogen J-2X engine derived from the J-2 engine used on the second stage of the Apollo vehicle will power the Ares V second stage. The Ares I can lift more than 55,000 pounds to low Earth orbit. The vehicle illustrated on the right is the Ares V, a heavy lift launch vehicle that will use five RS-68 liquid oxygen/liquid hydrogen engines mounted below a larger version of the space shuttle external tank, and two five-segment solid propellant rocket boosters for the first stage. The upper stage will use the same J-2X engine as the Ares I. The Ares V can lift more than 286,000 pounds to low Earth orbit and stands approximately 360 feet tall. This versatile system will be used to carry cargo and the components into orbit needed to go to the moon and later to Mars. Both vehicles are subject to configuration changes before they are actually launched. This illustration reflects the latest configuration as of September 2006.

Illustration of Ares I and Ares V Launch Vehicles

NASA Technical Reports Server (NTRS)

2006-01-01

Named for the Greek god associated with Mars, the NASA developed Ares launch vehicles will return humans to the moon and later take them to Mars and other destinations. In this early illustration, the vehicle depicted on the left is the Ares I. Ares I is an inline, two-stage rocket configuration topped by the Orion crew vehicle and its launch abort system. In addition to its primary mission of carrying four to six member crews to Earth orbit, Ares I may also use its 25-ton payload capacity to deliver resources and supplies to the International Space Station (ISS), or to 'park' payloads in orbit for retrieval by other spacecraft bound for the moon or other destinations. The Ares I employs a single five-segment solid rocket booster, a derivative of the space shuttle solid rocket booster, for the first stage. A liquid oxygen/liquid hydrogen J-2X engine derived from the J-2 engine used on the second stage of the Apollo vehicle will power the Ares V second stage. The Ares I can lift more than 55,000 pounds to low Earth orbit. The vehicle illustrated on the right is the Ares V, a heavy lift launch vehicle that will use five RS-68 liquid oxygen/liquid hydrogen engines mounted below a larger version of the space shuttle external tank, and two five-segment solid propellant rocket boosters for the first stage. The upper stage will use the same J-2X engine as the Ares I. The Ares V can lift more than 286,000 pounds to low Earth orbit and stands approximately 360 feet tall. This versatile system will be used to carry cargo and the components into orbit needed to go to the moon and later to Mars. Both vehicles are subject to configuration changes before they are actually launched. This illustration reflects the latest configuration as of September 2006.

Small Launch Vehicle Concept Development for Affordable Multi-Stage Inline Configurations

NASA Technical Reports Server (NTRS)

Beers, Benjamin R.; Waters, Eric D.; Philips, Alan D.; Threet, Grady E., Jr.

2014-01-01

The Advanced Concepts Office at NASA's George C. Marshall Space Flight Center conducted a study of two configurations of a three-stage, inline, liquid propellant small launch vehicle concept developed on the premise of maximizing affordability by targeting a specific payload capability range based on current and future industry demand. The initial configuration, NESC-1, employed liquid oxygen as the oxidizer and rocket propellant grade kerosene as the fuel in all three stages. The second and more heavily studied configuration, NESC-4, employed liquid oxygen and rocket propellant grade kerosene on the first and second stages and liquid oxygen and liquid methane fuel on the third stage. On both vehicles, sensitivity studies were first conducted on specific impulse and stage propellant mass fraction in order to baseline gear ratios and drive the focus of concept development. Subsequent sensitivity and trade studies on the NESC-4 concept investigated potential impacts to affordability due to changes in gross liftoff mass and/or vehicle complexity. Results are discussed at a high level to understand the impact severity of certain sensitivities and how those trade studies conducted can either affect cost, performance, or both.

Digital filtering of plume emission spectra

NASA Technical Reports Server (NTRS)

Madzsar, George C.

1990-01-01

Fourier transformation and digital filtering techniques were used to separate the superpositioned spectral phenomena observed in the exhaust plumes of liquid propellant rocket engines. Space shuttle main engine (SSME) spectral data were used to show that extraction of spectral lines in the spatial frequency domain does not introduce error, and extraction of the background continuum introduces only minimal error. Error introduced during band extraction could not be quantified due to poor spectrometer resolution. Based on the atomic and molecular species found in the SSME plume, it was determined that spectrometer resolution must be 0.03 nm for SSME plume spectral monitoring.

Large Eddy Simulations of Transverse Combustion Instability in a Multi-Element Injector

DTIC Science & Technology

2016-07-27

plagued the development of liquid rocket engines and remains a large riskin the development and acquisition of new liquid rocket engines. Combustion...simulations to better understand the physics that can lead combustion instability in liquid rocket engines. Simulations of this type are able to...instabilities found in liquid rocket engines are transverse. The motivating of the experiment behind the current work is to subject the CVRC injector

Solar-Thermal Engine Testing

NASA Technical Reports Server (NTRS)

Tucker, Stephen; Salvail, Pat; Haynes, Davy (Technical Monitor)

2001-01-01

A solar-thermal engine serves as a high-temperature solar-radiation absorber, heat exchanger, and rocket nozzle. collecting concentrated solar radiation into an absorber cavity and transferring this energy to a propellant as heat. Propellant gas can be heated to temperatures approaching 4,500 F and expanded in a rocket nozzle, creating low thrust with a high specific impulse (I(sub sp)). The Shooting Star Experiment (SSE) solar-thermal engine is made of 100 percent chemical vapor deposited (CVD) rhenium. The engine 'module' consists of an engine assembly, propellant feedline, engine support structure, thermal insulation, and instrumentation. Engine thermal performance tests consist of a series of high-temperature thermal cycles intended to characterize the propulsive performance of the engines and the thermal effectiveness of the engine support structure and insulation system. A silicone-carbide electrical resistance heater, placed inside the inner shell, substitutes for solar radiation and heats the engine. Although the preferred propellant is hydrogen, the propellant used in these tests is gaseous nitrogen. Because rhenium oxidizes at elevated temperatures, the tests are performed in a vacuum chamber. Test data will include transient and steady state temperatures on selected engine surfaces, propellant pressures and flow rates, and engine thrust levels. The engine propellant-feed system is designed to Supply GN2 to the engine at a constant inlet pressure of 60 psia, producing a near-constant thrust of 1.0 lb. Gaseous hydrogen will be used in subsequent tests. The propellant flow rate decreases with increasing propellant temperature, while maintaining constant thrust, increasing engine I(sub sp). In conjunction with analytical models of the heat exchanger, the temperature data will provide insight into the effectiveness of the insulation system, the structural support system, and the overall engine performance. These tests also provide experience on operational aspects of the engine and associated subsystems, and will include independent variation of both steady slate heat-exchanger temperature prior to thrust operation and nitrogen inlet pressure (flow rate) during thrust operation. Although the Shooting Star engines were designed as thermal-storage engines to accommodate mission parameters, they are fully capable of operating as scalable, direct-gain engines. Tests are conducted in both operational modes. Engine thrust and propellant flow rate will be measured and thereby I(sub sp). The objective of these tests is to investigate the effectiveness of the solar engine as a heat exchanger and a rocket. Of particular interest is the effectiveness of the support structure as a thermal insulator, the integrity of both the insulation system and the insulation containment system, the overall temperature distribution throughout the engine module, and the thermal power required to sustain steady state fluid temperatures at various flow rates.

Low thrust chemical rocket technology

NASA Technical Reports Server (NTRS)

Schneider, Steven J.

1992-01-01

An on-going technology program to improve the performance of low thrust chemical rockets for spacecraft on-board propulsion applications is reviewed. Improved performance and lifetime is sought by the development of new predictive tools to understand the combustion and flow physics, introduction of high temperature materials and improved component designs to optimize performance, and use of higher performance propellants. Improved predictive technology is sought through the comparison of both local and global predictions with experimental data. Predictions are based on both the RPLUS Navier-Stokes code with finite rate kinetics and the JANNAF methodology. Data were obtained with laser-based diagnostics along with global performance measurements. Results indicate that the modeling of the injector and the combustion process needs improvement in these codes and flow visualization with a technique such as 2-D laser induced fluorescence (LIF) would aid in resolving issues of flow symmetry and shear layer combustion processes. High temperature material fabrication processes are under development and small rockets are being designed, fabricated, and tested using these new materials. Rhenium coated with iridium for oxidation protection was produced by the Chemical Vapor Deposition (CVD) process and enabled an 800 K increase in rocket operating temperature. Performance gains with this material in rockets using Earth storable propellants (nitrogen tetroxide and monomethylhydrazine or hydrazine) were obtained through component redesign to eliminate fuel film cooling and its associated combustion inefficiency while managing head end thermal soakback. Material interdiffusion and oxidation characteristics indicated that the requisite lifetimes of tens of hours were available for thruster applications. Rockets were designed, fabricated, and tested with thrusts of 22, 62, 440 and 550 N. Performance improvements of 10 to 20 seconds specific impulse were demonstrated. Higher performance propellants were evaluated: Space storable propellants, including liquid oxygen (LOX) as the oxidizer with nitrogen hydrides or hydrocarbon as fuels. Specifically, a LOX/hydrazine engine was designed, fabricated, and shown to have a 95 pct theoretical c-star which translates into a projected vacuum specific impulse of 345 seconds at an area ratio of 204:1. Further performance improvment can be obtained by the use of LOX/hydrogen propellants, especially for manned spacecraft applications, and specific designs must be developed and advanced through flight qualification.

NASA Propulsion Engineering Research Center, volume 1

NASA Technical Reports Server (NTRS)

1993-01-01

Over the past year, the Propulsion Engineering Research Center at The Pennsylvania State University continued its progress toward meeting the goals of NASA's University Space Engineering Research Centers (USERC) program. The USERC program was initiated in 1988 by the Office of Aeronautics and Space Technology to provide an invigorating force to drive technology advancements in the U.S. space industry. The Propulsion Center's role in this effort is to provide a fundamental basis from which the technology advances in propulsion can be derived. To fulfill this role, an integrated program was developed that focuses research efforts on key technical areas, provides students with a broad education in traditional propulsion-related science and engineering disciplines, and provides minority and other under-represented students with opportunities to take their first step toward professional careers in propulsion engineering. The program is made efficient by incorporating government propulsion laboratories and the U.S. propulsion industry into the program through extensive interactions and research involvement. The Center is comprised of faculty, professional staff, and graduate and undergraduate students working on a broad spectrum of research issues related to propulsion. The Center's research focus encompasses both current and advanced propulsion concepts for space transportation, with a research emphasis on liquid propellant rocket engines. The liquid rocket engine research includes programs in combustion and turbomachinery. Other space transportation modes that are being addressed include anti-matter, electric, nuclear, and solid propellant propulsion. Outside funding supports a significant fraction of Center research, with the major portion of the basic USERC grant being used for graduate student support and recruitment. The remainder of the USERC funds are used to support programs to increase minority student enrollment in engineering, to maintain Center infrastructure, and to develop research capability in key new areas. Significant research programs in propulsion systems for air and land transportation complement the space propulsion focus. The primary mission of the Center is student education. The student program emphasizes formal class work and research in classical engineering and science disciplines with applications to propulsion.

Despin System for Hydrogen Tank in the Propulsion Systems Laboratory

NASA Image and Video Library

1962-04-21

Mechanic Howard Wine inspects the setup of a spin isolator in Cell 2 of the Propulsion Systems Laboratory at the National Aeronautics and Space Administration (NASA) Lewis Research Center. Photographer Al Jecko filmed the proceedings. This test was unique in that the chamber’s altitude system was used, but not its inlet air flow. The test was in preparation for an upcoming launch of modified liquid hydrogen propellant tank on a sounding rocket. This Weightlessness Analysis Sounding Probe (WASP) was part of Lewis investigation into methods for controlling partially filled liquid hydrogen fuel tanks during flight. Second-stage rockets, the Centaur in particular, were designed to stop their engines and coast, then restart them when needed. During this coast period, the propellant often shifted inside the tank. This movement could throw the rocket off course or result in the sloshing of fuel away from the fuel pump. Wine was one of only three journeymen mechanics at Lewis when he was hired in January 1954. He spent his first decade in the Propulsion Systems Laboratory and was soon named a section head. Wine went on to serve as Assistant Division Chief and later served as an assistant to the director. Jecko joined the center in 1947 as a photographer and artist. He studied at the Cleveland School or Art and was known for his cartoon drawing. He worked at the center for 26 years.

Robust Exploration and Commercial Missions to the Moon Using LANTR Propulsion and In-Situ Propellants Derived from Lunar Polar Ice (LPI) Deposits

NASA Technical Reports Server (NTRS)

Borowski, Stanley K.; Ryan, Stephen W.; Burke, Laura M.; McCurdy, David R.; Fittje, James E.; Joyner, Claude R.

2017-01-01

The nuclear thermal rocket (NTR) has frequently been identified as a key space asset required for the human exploration of Mars. This proven technology can also provide the affordable access through cislunar space necessary for commercial development and sustained human presence on the Moon. It is a demonstrated technology capable of generating both high thrust and high specific impulse (Isp 900 s) twice that of todays best chemical rockets. Nuclear lunar transfer vehicles consisting of a propulsion stage using three approx.16.5 klbf "Small Nuclear Rocket Engines (SNREs)", an in-line propellant tank, plus the payload can enable a variety of reusable lunar missions. These include cargo delivery and crewed lunar landing missions. Even weeklong "tourism" missions carrying passengers into lunar orbit for a day of sightseeing and picture taking are possible. The NTR can play an important role in the next phase of lunar exploration and development by providing a robust in-space lunar transportation system (LTS) that can allow initial outposts to evolve into settlements supported by a variety of commercial activities such as in-situ propellant production used to supply strategically located propellant depots and transportation nodes. The processing of LPI deposits (estimated to be approx. 2 billion metric tons) for propellant production - specifically liquid oxygen (LO2) and hydrogen (LH2) can significantly reduce the launch mass requirements from Earth and can enable reusable, surface-based lunar landing vehicles (LLVs) using LO2/LH2 chemical rocket engines. Afterwards, LO2/LH2 propellant depots can be established in lunar polar and equatorial orbits to supply the LTS. At this point a modified version of the conventional NTR called the LO2-augmented NTR, or LANTR would be introduced into the LTS allowing bipropellant operation and leveraging the mission benefits of refueling with lunar-derived propellants (LDPs) for Earth return. The bipropellant LANTR engine utilizes the large divergent section of its nozzle as an afterburner into which oxygen is injected and supersonically combusted with nuclear preheated hydrogen emerging from the engines choked sonic throat essentially scramjet propulsion in reverse. By varying the oxygen-to-hydrogen mixture ratio, LANTR engines can operate over a range of thrust and Isp values while the reactor core power level remains relatively constant. A LANTR-based LTS offers unique mission capabilities including short transit time crewed cargo transports. Even a commuter shuttle service may be possible allowing one-way trip times to and from the Moon on the order of 36 hours or less. If only 1 of the postulated water ice trapped in deep shadowed craters at the lunar poles were available for use in lunar orbit, such a supply could support daily commuter flights to the Moon for many thousands of years! The proposed paper outlines an evolutionary mission architecture and examines a variety of mission types and transfer vehicle designs, along with the increasing demands on LDP production as mission complexity and delta V requirements increase. A comparison of vehicle features and engine operating characteristics are also provided together with a discussion of the propellant production and mining requirements, and issues, associated with using LPI as the source material.

Cryogenic Orbital Testbed (CRYOTE) Ground Test Article, Final Report

NASA Technical Reports Server (NTRS)

Johnson, Wesley L.; Rhys, Noah O.; Bradley, David E.; Wollen, Mark; Kutter, Bernard; Gravlee, Mari; Walls, Laurie K.

2015-01-01

Liquid propulsion has been used since Robert Goddard started developing a liquid oxygen (LO2) and gasoline powered rocket and fired it in 1923 (Ref. 1). In the following decades engineers settled on the combination of liquid hydrogen (LH2) and LO2 as the most efficient propellant combination for in-space travel. Due to their low temperatures (LH2 at 20 K and LO2 at 90 K), they require special handling and procedures. General Dynamics began developing LO2 and LH2 upper stages in 1956 in the form of Centaur, these efforts were soon funded by the Department of Defense in conjunction with NASA (beginning in 1958) (Ref. 2). Meanwhile NASA also worked with McDonnell Douglas to develop the SIV-B stage for the Saturn V rocket. In the subsequent years, the engineers were able to push the Centaur to up to 9 hr of orbital lifetime and the SIV-B to up to 6 hr. Due to venting the resultant boil-off from the high heat loads through the foam insulation on the upper stages, both vehicles remained in a settled configuration throughout the flights, thus the two phases of propellant (liquid and vapor) were separated at a known location. The one exception to this were the Titan/Centaur missions, which thanks to the lower boil-off using three layers of multilayer insulation (MLI), were able to coast unsettled for up to 5.25 hr during direct geosynchronous orbit insertion missions. In the years since there has been a continuous effort to extend the life of these upper stages from hours to days or even months.

Numerical and experimental analysis of heat transfer in injector plate of hydrogen peroxide hybrid rocket motor

NASA Astrophysics Data System (ADS)

Cai, Guobiao; Li, Chengen; Tian, Hui

2016-11-01

This paper is aimed to analyze heat transfer in injector plate of hydrogen peroxide hybrid rocket motor by two-dimensional axisymmetric numerical simulations and full-scale firing tests. Long-time working, which is an advantage of hybrid rocket motor over conventional solid rocket motor, puts forward new challenges for thermal protection. Thermal environments of full-scale hybrid rocket motors designed for long-time firing tests are studied through steady-state coupled numerical simulations of flow field and heat transfer in chamber head. The motor adopts 98% hydrogen peroxide (98HP) oxidizer and hydroxyl-terminated poly-butadiene (HTPB) based fuel as the propellants. Simulation results reveal that flowing liquid 98HP in head oxidizer chamber could cool the injector plate of the motor. The cooling of 98HP is similar to the regenerative cooling in liquid rocket engines. However, the temperature of the 98HP in periphery portion of the head oxidizer chamber is higher than its boiling point. In order to prevent the liquid 98HP from unexpected decomposition, a thermal protection method for chamber head utilizing silica-phenolics annular insulating board is proposed. The simulation results show that the annular insulating board could effectively decrease the temperature of the 98HP in head oxidizer chamber. Besides, the thermal protection method for long-time working hydrogen peroxide hybrid rocket motor is verified through full-scale firing tests. The ablation of the insulating board in oxygen-rich environment is also analyzed.

Historical perspective - Viking Mars Lander propulsion

NASA Technical Reports Server (NTRS)

Morrisey, Donald C.

1989-01-01

This paper discusses the Viking 1 and 2 missions to Mars in 1975-1976 and describes the design evolution of the Viking Terminal Descent Rocket Engines responsible for decelerating the Viking Mars Landers during the final portion of their descent from orbit. The Viking Terminal Descent Rocket Engines have twice the thrust of the largest monopropellant hydrazine engine developed previously but weigh considerably less. The engine has 18 nozzles, the capability of 10:1 throttling, is totally sealed until fired, employs no organic unsealed materials, is 100 percent germ free, utilized hydrazine STM-20 as the propellant, and starts at a temperature more than 45 F below the propellant's freezing point.

Space Launch System Base Heating Test: Experimental Operations & Results

NASA Technical Reports Server (NTRS)

Dufrene, Aaron; Mehta, Manish; MacLean, Matthew; Seaford, Mark; Holden, Michael

2016-01-01

NASA's Space Launch System (SLS) uses four clustered liquid rocket engines along with two solid rocket boosters. The interaction between all six rocket exhaust plumes will produce a complex and severe thermal environment in the base of the vehicle. This work focuses on a recent 2% scale, hot-fire SLS base heating test. These base heating tests are short-duration tests executed with chamber pressures near the full-scale values with gaseous hydrogen/oxygen engines and RSRMV analogous solid propellant motors. The LENS II shock tunnel/Ludwieg tube tunnel was used at or near flight duplicated conditions up to Mach 5. Model development was based on the Space Shuttle base heating tests with several improvements including doubling of the maximum chamber pressures and duplication of freestream conditions. Test methodology and conditions are presented, and base heating results from 76 runs are reported in non-dimensional form. Regions of high heating are identified and comparisons of various configuration and conditions are highlighted. Base pressure and radiometer results are also reported.

Preliminary study of a hydrogen peroxide rocket for use in moving source jet noise tests

NASA Technical Reports Server (NTRS)

Plencner, R. M.

1977-01-01

A preliminary investigation was made of using a hydrogen peroxide rocket to obtain pure moving source jet noise data. The thermodynamic cycle of the rocket was analyzed. It was found that the thermodynamic exhaust properties of the rocket could be made to match those of typical advanced commercial supersonic transport engines. The rocket thruster was then considered in combination with a streamlined ground car for moving source jet noise experiments. When a nonthrottlable hydrogen peroxide rocket was used to accelerate the vehicle, propellant masses and/or acceleration distances became too large. However, when a throttlable rocket or an auxiliary system was used to accelerate the vehicle, reasonable propellant masses could be obtained.

Rho-Isp Revisited and Basic Stage Mass Estimating for Launch Vehicle Conceptual Sizing Studies

NASA Technical Reports Server (NTRS)

Kibbey, Timothy P.

2015-01-01

The ideal rocket equation is manipulated to demonstrate the essential link between propellant density and specific impulse as the two primary stage performance drivers for a launch vehicle. This is illustrated by examining volume-limited stages such as first stages and boosters. This proves to be a good approximation for first-order or Phase A vehicle design studies for solid rocket motors and for liquid stages, except when comparing to hydrogen-fueled stages. A next-order mass model is developed that is able to model the mass differences between hydrogen-fueled and other stages. Propellants considered range in density from liquid methane to inhibited red fuming nitric acid. Calculated comparisons are shown for solid rocket boosters, liquid first stages, liquid upper stages, and a balloon-deployed single-stage-to-orbit concept. The derived relationships are ripe for inclusion in a multi-stage design space exploration and optimization algorithm, as well as for single-parameter comparisons such as those shown herein.

Combustion response to acoustic perturbation in liquid rocket engines

NASA Astrophysics Data System (ADS)

Ghafourian, Akbar

An experimental study of the effect of acoustic perturbations on combustion behavior of a model liquid propellant rocket engine has been carried out. A pair of compression drivers were used to excite transverse and longitudinal acoustic fields at strengths of up to 156.6 dB and 159.5 dB respectively in the combustion chamber of the experimental rocket engine. Propellant simulants were injected into the combustion chamber through a single element shear coaxial injector. Water and air were used in cold flow studies and ethanol and oxygen-enriched air were used as fuel and oxidizer in reacting hot flow studies. In cold flow studies an imposed transverse acoustic field had a more pronounced effect on the spray pattern than a longitudinal acoustic fields. A transverse acoustic field widened the spray by as much as 33 percent and the plane of impingement of the spray with chamber walls moved up closer to the injection plane. The behavior was strongly influenced by the gas phase velocity but was less sensitive to changes in the liquid phase velocity. In reacting hot flow studies the effects of changes in equivalence ratio, excitation amplitude, excitation frequency, liquid and gas phase velocity and chamber pressure on the response of the injector to imposed high frequency transverse acoustic excitation were measured. Reducing the equivalence ratio from 7.4 to 3.8 increased the chamber pressure response to the imposed excitation at 3000 Hz. Increasing the excitation amplitude from 147 dB to 155.6 dB at 3000 Hz increased the chamber pressure response to the excitation. In the frequency range of 1240 Hz to 3220 Hz, an excitation frequency of 3000 Hz resulted in the largest response of the chamber pressure indicating the importance of fluid dynamic coupling. Increasing the liquid phase velocity from 9.2 m/sec to 22.7 m/sec, did not change the amplitude of the chamber pressure response to excitation. This implied the importance of local equivalence ratio and not the overall equivalence ratio on chamber pressure response to excitation. Increasing the chamber pressure from 1.5 atm to 3.1 atm and gas phase velocity from 93.2 m/sec to 105.1 m/sec significantly increased the chamber pressure response to acoustic excitation. This emphasized the significance of the gas phase density and velocity. Measurements of the free radical C2 emission zone and Schlieren images indicated that transverse acoustic excitation moved the combustion zone closer to the injection plane and longitudinal acoustic excitation widened the combustion zone. The histogram of these images indicates that the area over which combustion takes place in the chamber increases under imposed acoustic excitation. This implied that more propellants combust prior to exiting from the exhaust nozzle under unsteady conditions.

Multi-Stage Hybrid Rocket Conceptual Design for Micro-Satellites Launch using Genetic Algorithm

NASA Astrophysics Data System (ADS)

Kitagawa, Yosuke; Kitagawa, Koki; Nakamiya, Masaki; Kanazaki, Masahiro; Shimada, Toru

The multi-objective genetic algorithm (MOGA) is applied to the multi-disciplinary conceptual design problem for a three-stage launch vehicle (LV) with a hybrid rocket engine (HRE). MOGA is an optimization tool used for multi-objective problems. The parallel coordinate plot (PCP), which is a data mining method, is employed in the post-process in MOGA for design knowledge discovery. A rocket that can deliver observing micro-satellites to the sun-synchronous orbit (SSO) is designed. It consists of an oxidizer tank containing liquid oxidizer, a combustion chamber containing solid fuel, a pressurizing tank and a nozzle. The objective functions considered in this study are to minimize the total mass of the rocket and to maximize the ratio of the payload mass to the total mass. To calculate the thrust and the engine size, the regression rate is estimated based on an empirical model for a paraffin (FT-0070) propellant. Several non-dominated solutions are obtained using MOGA, and design knowledge is discovered for the present hybrid rocket design problem using a PCP analysis. As a result, substantial knowledge on the design of an LV with an HRE is obtained for use in space transportation.

Advanced space engine preliminary design. [liquid hydrogen/liquid oxygen upper stage engine for space tug application

NASA Technical Reports Server (NTRS)

Zachary, A. T.

1973-01-01

Analysis and design of an optimum LO2/LH2, combustion topping cycle, 88,964 Newtons (20,000-pound) thrust, liquid rocket engine was conducted. The design selected is well suited to high-energy, upper-stage engine applications such as the Space Tug and embodies features directed toward optimization of vehicle performance. A configuration selection was conducted based on prior Air Force Contracts, and additional criteria for optimum stage performance. Following configuration selection, analyses and design of the major components and engine systems were conducted to sufficient depth to provide layout drawings suitable for subsequent detailing. In addition, engine packaging to a common interface and a retractable nozzle concept were defined. Alternative development plans and related costs were also established. The design embodies high-performance, low-weight, low NPSH requirements (saturated propellant inlet conditions at start), idle-mode operation, and autogenous pressurization. The design is the result of the significant past and current LO2/LH2 technology efforts of the NASA centers and the Air Force, as well as company-funded programs.

User's manual for rocket combustor interactive design (ROCCID) and analysis computer program. Volume 1: User's manual

NASA Technical Reports Server (NTRS)

Muss, J. A.; Nguyen, T. V.; Johnson, C. W.

1991-01-01

The user's manual for the rocket combustor interactive design (ROCCID) computer program is presented. The program, written in Fortran 77, provides a standardized methodology using state of the art codes and procedures for the analysis of a liquid rocket engine combustor's steady state combustion performance and combustion stability. The ROCCID is currently capable of analyzing mixed element injector patterns containing impinging like doublet or unlike triplet, showerhead, shear coaxial, and swirl coaxial elements as long as only one element type exists in each injector core, baffle, or barrier zone. Real propellant properties of oxygen, hydrogen, methane, propane, and RP-1 are included in ROCCID. The properties of other propellants can easily be added. The analysis model in ROCCID can account for the influence of acoustic cavities, helmholtz resonators, and radial thrust chamber baffles on combustion stability. ROCCID also contains the logic to interactively create a combustor design which meets input performance and stability goals. A preliminary design results from the application of historical correlations to the input design requirements. The steady state performance and combustion stability of this design is evaluated using the analysis models, and ROCCID guides the user as to the design changes required to satisfy the user's performance and stability goals, including the design of stability aids. Output from ROCCID includes a formatted input file for the standardized JANNAF engine performance prediction procedure.

Can a pile of scrap unmask a new high technology? The A4/V-2 No V89 Bäckebo-torpeden

NASA Astrophysics Data System (ADS)

Ingemar Skoog, A.

2013-04-01

Three months before the first V-2 rocket attack on London a test vehicle crashed in southern Sweden on June 13, 1944. At this time the Allied only had limited knowledge about the rocket (A4/V-2) from agent reports and information from the Polish resistance investigating some remains from a crashed test vehicle in Poland. London was confronted with a new weapon supposedly able to carry an explosive warhead of several tons some 250 km. The A4/V-2 rocket test vehicle number V89 broke apart shortly before impacting ground. In a short time 2 t of metal parts and electrical equipment was collected and transported to Stockholm for investigations. A first Swedish report was ready by July 21, 1944 and the rocket parts were then transported to England for further investigations. By August 18, 1944 the Royal Aircraft Establishment (RAE) had its preliminary report ready. But how close to reality can a complex vehicle be reconstructed and the performance calculated from a pile of scrap by investigators dealing with a technology not seen before? In the early 1940s the state of art of liquid propellant rocket technology outside Germany was limited and the size of a liquid rocket engine for the likely performance hardly imaginable. The Swedish and British reports, at that time classified as top secret, have since been released and permit a very detailed analysis of the task to reconstruct the rocket vehicle, the engine itself and its performance. An assessment of the occurrence at Peenemünde and how the rocket became astray and fell in southern Sweden, together with the analyses by Swedish and British military investigators give a unique insight into the true nature of the V89. It shows the real capabilities of early aeronautical accident investigation methods in combination with solid engineering knowledge to unmask a new high technology.

Observation of rocket pollution with overhead sensors

NASA Astrophysics Data System (ADS)

Fisher, Annette

2011-12-01

The objective of this thesis is to study the dispersal of rocket pollution through remote sensing techniques. Substantial research with remote sensors has been dedicated to observation of volcanic plumes, particulate dispersion, and aircraft contrails with less emphasis on observing rocket launches and the effects on the surrounding environment. This research focuses on observation of rocket exhaust constituents, particularly carbon soot, alumina, and water vapor. The sensors utilized in this thesis have unique capabilities that provide measurements that are likely capable of detecting the rocket exhaust constituents. Methodology and analysis included choosing an appropriate launch vehicle with obtainable launch data and various booster combinations of liquid propellant only or a combination of liquid and solid propellant, prioritizing the data based on launch time versus sensor passing, processing the data, and applying known constituent properties to the data sets where key areas of work in this endeavor. Results of this work demonstrate a unique capability in monitoring man-made pollution and the extent the pollution can spread to surrounding areas.

Analysis of quasi-hybrid solid rocket booster concepts for advanced earth-to-orbit vehicles

NASA Technical Reports Server (NTRS)

Zurawski, Robert L.; Rapp, Douglas C.

1987-01-01

A study was conducted to assess the feasibility of quasi-hybrid solid rocket boosters for advanced Earth-to-orbit vehicles. Thermochemical calculations were conducted to determine the effect of liquid hydrogen addition, solids composition change plus liquid hydrogen addition, and the addition of an aluminum/liquid hydrogen slurry on the theoretical performance of a PBAN solid propellant rocket. The space shuttle solid rocket booster was used as a reference point. All three quasi-hybrid systems theoretically offer higher specific impulse when compared with the space shuttle solid rocket boosters. However, based on operational and safety considerations, the quasi-hybrid rocket is not a practical choice for near-term Earth-to-orbit booster applications. Safety and technology issues pertinent to quasi-hybrid rocket systems are discussed.

Propellant Feed System Leak Detection: Lessons Learned From the Linear Aerospike SR-71 Experiment (LASRE)

NASA Technical Reports Server (NTRS)

Hass, Neal; Mizukami, Masashi; Neal, Bradford A.; St. John, Clinton; Beil, Robert J.; Griffin, Timothy P.

1999-01-01

This paper presents pertinent results and assessment of propellant feed system leak detection as applied to the Linear Aerospike SR-71 Experiment (LASRE) program flown at the NASA Dryden Flight Research Center, Edwards, California. The LASRE was a flight test of an aerospike rocket engine using liquid oxygen and high-pressure gaseous hydrogen as propellants. The flight safety of the crew and the experiment demanded proven technologies and techniques that could detect leaks and assess the integrity of hazardous propellant feed systems. Point source detection and systematic detection were used. Point source detection was adequate for catching gross leakage from components of the propellant feed systems, but insufficient for clearing LASRE to levels of acceptability. Systematic detection, which used high-resolution instrumentation to evaluate the health of the system within a closed volume, provided a better means for assessing leak hazards. Oxygen sensors detected a leak rate of approximately 0.04 cubic inches per second of liquid oxygen. Pressure sensor data revealed speculated cryogenic boiloff through the fittings of the oxygen system, but location of the source(s) was indeterminable. Ultimately, LASRE was cancelled because leak detection techniques were unable to verify that oxygen levels could be maintained below flammability limits.

Magnesium and Carbon Dioxide - A Rocket Propellant for Mars Missions

NASA Technical Reports Server (NTRS)

Shafirovich, E. IA.; Shiriaev, A. A.; Goldshleger, U. I.

1993-01-01

A rocket engine for Mars missions is proposed that could utilize CO2 accumulated from the Martian atmosphere as an oxidizer. For use as possible fuel, various metals, their hydrides, and mixtures with hydrogen compounds are considered. Thermodynamic calculations show that beryllium fuels ensure the most impulse but poor inflammability of Be and high toxicity of its compounds put obstacles to their applications. Analysis of the engine performance for other metals together with the parameters of ignition and combustion show that magnesium seems to be the most promising fuel. Ballistic estimates imply that a hopper with the chemical rocket engine on Mg + CO2 propellant could be readily developed. This vehicle would be able to carry out 2-3 ballistic flights on Mars before the final ascent to orbit.

Moving, Moving, Moving- A Giant Rocket Fuel Tank

NASA Image and Video Library

2016-10-07

Technicians moved a giant fuel tank from the Vertical Assembly Center where the tank recently completed friction stir welding to an adjacent work area at NASA's Michoud Assembly Facility in New Orleans. More than 1.7 miles of welds have been completed for core stage hardware at Michoud. This liquid hydrogen fuel tank is the largest piece of the core stage that will provide the fuel for the first flight of NASA's new rocket, the Space Launch System, with the Orion spacecraft in 2018. The tank is more than 130 feet long, and together with the liquid oxygen tank holds 733,000 gallons of propellant to feed the vehicle's four RS-25 engines to produce a total of 2 million pounds of thrust. SLS will have the power and capacity to carry humans to Mars. For more information on the core stage: http://www.nasa.gov/exploration/syste... Video Credit: NASA/MAF/Eric Bordelon

Extensions to the time lag models for practical application to rocket engine stability design

NASA Astrophysics Data System (ADS)

Casiano, Matthew J.

The combustion instability problem in liquid-propellant rocket engines (LREs) has remained a tremendous challenge since their discovery in the 1930s. Improvements are usually made in solving the combustion instability problem primarily using computational fluid dynamics (CFD) and also by testing demonstrator engines. Another approach is to use analytical models. Analytical models can be used such that design, redesign, or improvement of an engine system is feasible in a relatively short period of time. Improvements to the analytical models can greatly aid in design efforts. A thorough literature review is first conducted on liquid-propellant rocket engine (LRE) throttling. Throttling is usually studied in terms of vehicle descent or ballistic missile control however there are many other cases where throttling is important. It was found that combustion instabilities are one of a few major issues that occur during deep throttling (other major issues are heat transfer concerns, performance loss, and pump dynamics). In the past and again recently, gas injected into liquid propellants has shown to be a viable solution to throttle engines and to eliminate some forms of combustion instability. This review uncovered a clever solution that was used to eliminate a chug instability in the Common Extensible Cryogenic Engine (CECE), a modified RL10 engine. A separate review was also conducted on classic time lag combustion instability models. Several new stability models are developed by incorporating important features to the classic and contemporary models, which are commonly used in the aerospace rocket industry. The first two models are extensions of the original Crocco and Cheng concentrated combustion model with feed system contributions. A third new model is an extension to the Wenzel and Szuch double-time lag model also with feed system contributions. The first new model incorporates the appropriate injector acoustic boundary condition which is neglected in contemporary models. This new feature shows that the injector boundary can play a significant role for combustion stability, especially for gaseous injection systems or a system with an injector orifice on the order of the size of the chamber. The second new model additionally accounts for resistive effects. Advanced signal analysis techniques are used to extract frequency-dependent damping from a gas generator component data set. The damping values are then used in the new stability model to more accurately represent the chamber response of the component. The results show a more realistic representation of stability margin by incorporating the appropriate damping effects into the chamber response from data. The original Crocco model, a contemporary model, and the two new models are all compared and contrasted to a marginally stable test case showing their applicability. The model that incorporates resistive aspects shows the best comparison to the test data. Parametrics are also examined to show the influence of the new features and their applicability. The new features allow a more accurate representation of stability margin to be obtained. The third new model is an extension to the Wenzel and Szuch double-time lag chug model. The feed system chug model is extended to account for generic propellant flow rates. This model is also extended to incorporate aspects due to oxygen boiling and helium injection in the feed system. The solutions to the classic models, for the single-time lag and the double-time lag models, are often plotted on a practical engine operating map, however the models have presented some difficulties for numerical algorithms for several reasons. Closed-form solutions for use on these practical operating maps are formulated and developed. These models are incorporated in a graphical user interface tool and the new model is compared to an extensive data set. It correctly predicts the stability behavior at various operating conditions incorporating the influence of injected helium and boiling oxygen in the feed system.

Advanced Chemical Propulsion for Science Missions

NASA Technical Reports Server (NTRS)

Liou, Larry

2008-01-01

The advanced chemical propulsion technology area of NASA's In-Space Technology Project is investing in systems and components for increased performance and reduced cost of chemical propulsion technologies applicable to near-term science missions. Presently the primary investment in the advanced chemical propulsion technology area is in the AMBR high temperature storable bipropellant rocket engine. Scheduled to be available for flight development starting in year 2008, AMBR engine shows a 60 kg payload gain in an analysis for the Titan-Enceladus orbiter mission and a 33 percent manufacturing cost reduction over its baseline, state-of-the-art counterpart. Other technologies invested include the reliable lightweight tanks for propellant and the precision propellant management and mixture ratio control. Both technologies show significant mission benefit, can be applied to any liquid propulsion system, and upon completion of the efforts described in this paper, are at least in parts ready for flight infusion. Details of the technologies are discussed.

Metallized Gelled Propellants: Oxygen/RP-1/aluminum Rocket Combustion Experiments

NASA Technical Reports Server (NTRS)

Palaszewski, Bryan; Zakany, James S.

1995-01-01

A series of combustion experiments were conducted to measure the specific impulse, Cstar-, and specific-impulse efficiencies of a rocket engine using metallized gelled liquid propellants. These experiments used a small 20- to 40-1bf (89- to 178-N) thrust, modular engine consisting of an injector, igniter, chamber and nozzle. The fuels used were traditional liquid RP-1 and gelled RP-1 with 0-, 5-, and 55-wt% loadings of aluminum and gaseous oxygen was the oxidizer. Ten different injectors were used during the testing: 6 for the baseline 02/RP-1 tests and 4 for the gelled fuel tests which covered a wide range of mixture ratios. At the peak of the Isp versus oxidizer-to-fuel ratio (O/F) data, a range of 93 to 99% Cstar efficiency was reached with ungelled 02/RP-1. A Cstar efficiency range of 75 to 99% was obtained with gelled RP-l (0-wt% RP-1/Al) while the metallized 5-wt% RP-1/Al delivered a Cstar efficiency of 94 to 99% at the peak Isp in the O/F range tested. An 88 to 99% Cstar efficiency was obtained at the peak Isp of the gelled RP1/Al with 55-wt% Al. Specific impulse efficiencies for the 55-wt% RP-1/Al of 67%-83% were obtained at a 2.4:1 expansion ratio. Injector erosion was evident with the 55-wt% testing, while there was little or no erosion seen with the gelled RP-1 with 0- and 5-wt% Al. A protective layer of gelled fuel formed in the firings that minimized the damage to the rocket injector face. This effect may provide a useful technique for engine cooling. These experiments represent a first step in characterizing the performance of and operational issues with gelled RP-1 fuels.

Booster Main Engine Selection Criteria for the Liquid Fly-Back Booster

NASA Technical Reports Server (NTRS)

Ryan, Richard M.; Rothschild, William J.; Christensen, David L.

1998-01-01

The Liquid Fly-Back Booster (LFBB) Program seeks to enhance the Space Shuttle system safety performance and economy of operations through the use of an advanced, liquid propellant Booster Main Engine (BME). There are several viable BME candidates that could be suitable for this application. The objective of this study was to identify the key criteria to be applied in selecting among these BME candidates. This study involved an assessment of influences on the overall LFBB utility due to variations in the candidate rocket engines' characteristics. This includes BME impacts on vehicle system weight, perfortnance,design approaches, abort modes, margins of safety, engine-out operations, and maintenance and support concepts. Systems engineering analyses and trade studies were performed to identify the LFBB system level sensitivities to a wide variety of BME related parameters. This presentation summarizes these trade studies and the resulting findings of the LFBB design teams regarding the BME characteristics that most significantly affect the LFBB system. The resulting BME choice should offer the best combination of reliability, performance, reusability, robustness, cost, and risk for the LFBB program.

Booster Main Engine Selection Criteria for the Liquid Fly-Back Booster

NASA Technical Reports Server (NTRS)

Ryan, Richard M.; Rothschild, William J.; Christensen, David L.

1998-01-01

The Liquid Fly-Back Booster (LFBB) Program seeks to enhance the Space Shuttle system safety, performance and economy of operations through the use of an advanced, liquid propellant Booster Main Engine (BME). There are several viable BME candidates that could be suitable for this application. The objective of this study was to identify the key Criteria to be applied in selecting among these BME candidates. This study involved an assessment of influences on the overall LFBB utility due to variations in the candidate rocket-engines characteristics. This includes BME impacts on vehicle system weight, performance, design approaches, abort modes, margins of safety, engine-out operations, and maintenance and support concepts. Systems engineering analyses and trade studies were performed to identify the LFBB system level sensitivities to a wide variety of BME related parameters. This presentation summarizes these trade studies and the resulting findings of the LFBB design teams regarding the BME characteristics that most significantly affect the LFBB system. The resulting BME choice should offer the best combination of reliability, performance, reusability, robustness, cost, and risk for the LFBB program.

Study of solid rocket motor for space shuttle booster, Volume 3: Program acquisition planning

NASA Technical Reports Server (NTRS)

1972-01-01

The program planning acquisition functions for the development of the solid propellant rocket engine for the space shuttle booster is presented. The subjects discussed are: (1) program management, (2) contracts administration, (3) systems engineering, (4) configuration management, and (5) maintenance engineering. The plans for manufacturing, testing, and operations support are included.

Tail Service Mast Umbilical Arrival

NASA Image and Video Library

2016-08-02

A crane lowers the first Tail Service Mast Umbilical (TSMU) onto a test stand at the Launch Equipment Test Facility at NASA’s Kennedy Space Center in Florida. Two TSMUs will provide liquid propellants and power to the Space Launch System (SLS) rocket’s core stage engine. Both TSMUs will connect to the zero-level deck on the mobile launcher, providing fuel and electricity to the SLS rocket before it launches on Exploration Mission 1. The TSMU will undergo testing and validation at the LETF to verify it is functioning properly. The center’s Engineering Directorate and the Ground Systems Development and Operations Program are overseeing processing and testing of the umbilicals.

Advanced transportation system studies technical area 3: Alternate propulsion subsystem concepts, volume 2

NASA Technical Reports Server (NTRS)

Levak, Daniel

1993-01-01

The Alternate Propulsion Subsystem Concepts contract had five tasks defined for the first year. The tasks were: F-1A Restart Study, J-2S Restart Study, Propulsion Database Development, Space Shuttle Main Engine (SSME) Upper Stage Use, and CER's for Liquid Propellant Rocket Engines. The detailed study results, with the data to support the conclusions from various analyses, are being reported as a series of five separate Final Task Reports. Consequently, this volume only reports the required programmatic information concerning Computer Aided Design Documentation, and New Technology Reports. A detailed Executive Summary, covering all the tasks, is also available as Volume 1.

Metallic Hydrogen - Potentially a High Energy Rocket Propellant

NASA Technical Reports Server (NTRS)

Cole, John; Silvera, Ike

2007-01-01

Pure metallic hydrogen is predicted to have a specific impulse (Isp) of 1700 seconds, but the reaction temperature is too high for current engine materials. Diluting metallic hydrogen with liquid hydrogen can reduce the reaction temperature to levels compatible with current material limits and still provide an Isp greater than 900 s. Metallic hydrogen has not yet been produced on earth, but experimental techniques exist that may change this situation. This paper will provide a brief description of metallic hydrogen and the status of experiments that may soon produce detectable quantities of this material in the lab. Also provided are some characteristics for diluted metallic hydrogen engines and launch vehicles.

Design of a Mars Airplane Propulsion System for the Aerial Regional-Scale Environmental Survey (ARES) Mission Concept

NASA Technical Reports Server (NTRS)

Kuhl, Christopher A.

2008-01-01

The Aerial Regional-Scale Environmental Survey (ARES) is a Mars exploration mission concept that utilizes a rocket propelled airplane to take scientific measurements of atmospheric, surface, and subsurface phenomena. The liquid rocket propulsion system design has matured through several design cycles and trade studies since the inception of the ARES concept in 2002. This paper describes the process of selecting a bipropellant system over other propulsion system options, and provides details on the rocket system design, thrusters, propellant tank and PMD design, propellant isolation, and flow control hardware. The paper also summarizes computer model results of thruster plume interactions and simulated flight performance. The airplane has a 6.25 m wingspan with a total wet mass of 185 kg and has to ability to fly over 600 km through the atmosphere of Mars with 45 kg of MMH / MON3 propellant.

Investigation of Cleanliness Verification Techniques for Rocket Engine Hardware

NASA Technical Reports Server (NTRS)

Fritzemeier, Marilyn L.; Skowronski, Raymund P.

1994-01-01

Oxidizer propellant systems for liquid-fueled rocket engines must meet stringent cleanliness requirements for particulate and nonvolatile residue. These requirements were established to limit residual contaminants which could block small orifices or ignite in the oxidizer system during engine operation. Limiting organic residues in high pressure oxygen systems, such as in the Space Shuttle Main Engine (SSME), is particularly important. The current method of cleanliness verification for the SSME uses an organic solvent flush of the critical hardware surfaces. The solvent is filtered and analyzed for particulate matter followed by gravimetric determination of the nonvolatile residue (NVR) content of the filtered solvent. The organic solvents currently specified for use (1, 1, 1-trichloroethane and CFC-113) are ozone-depleting chemicals slated for elimination by December 1995. A test program is in progress to evaluate alternative methods for cleanliness verification that do not require the use of ozone-depleting chemicals and that minimize or eliminate the use of solvents regulated as hazardous air pollutants or smog precursors. Initial results from the laboratory test program to evaluate aqueous-based methods and organic solvent flush methods for NVR verification are provided and compared with results obtained using the current method. Evaluation of the alternative methods was conducted using a range of contaminants encountered in the manufacture of rocket engine hardware.

Subsystem Analysis/Optimization for the X-34 Main Propulsion System

NASA Technical Reports Server (NTRS)

McDonald, J. P.; Hedayat, A.; Brown, T. M.; Knight, K. C.; Champion, R. H., Jr.

1998-01-01

The Orbital Sciences Corporation X-34 vehicle demonstrates technologies and operations key to future reusable launch vehicles. The general flight performance goal of this unmanned rocket plane is Mach 8 flight at an altitude of 250,000 feet. The Main Propulsion System (MPS) supplies liquid propellants to the main engine, which provides the primary thrust for attaining mission goals. Major MPS design and operational goals are aircraft-like ground operations, quick turnaround between missions, and low initial/operational costs. Analyses related to optimal MPS subsystem design are reviewed in this paper. A pressurization system trade weighs maintenance/reliability concerns against those for safety in a comparison of designs using pressure regulators versus orifices to control pressurant flow. A propellant dump/feed system analysis weighs the issues of maximum allowable vehicle landing weight, trajectory, and MPS complexity to arrive at a final configuration for propellant dump/feed systems.

Photoignition Torch Applied to Cryogenic H2/O2 Coaxial Jet

DTIC Science & Technology

2016-12-06

suitable for certain thrusters and liquid rocket engines. This ignition system is scalable for applications in different combustion chambers such as gas ...turbines, gas generators, liquid rocket engines, and multi grain solid rocket motors. photoignition, fuel spray ignition, high pressure ignition...thrusters and liquid rocket engines. This ignition system is scalable for applications in different combustion chambers such as gas turbines, gas

The Swedish Rocket Corps, 1833 - 1845

NASA Technical Reports Server (NTRS)

Skoog, A. I.

1977-01-01

Rockets for pyrotechnic displays used in Sweden in the 19th century are examined in terms of their use in war situations. Work done by the Swedish chemist J. J. Berzelius, who analyzed and improved the propellants of such rockets, and the German engineer, Martin Westermaijer, who researched manufacturing techniques of these rockets is also included.

Raman Spectroscopy for Instantaneous Multipoint, Multispecies Gas Concentration and Temperature Measurements in Rocket Engine Propellant Injector Flows

NASA Technical Reports Server (NTRS)

Wehrmeyer, Joseph A.; Trinh, Huu Phuoc

2001-01-01

Propellant injector development at MSFC includes experimental analysis using optical techniques, such as Raman, fluorescence, or Mie scattering. For the application of spontaneous Raman scattering to hydrocarbon-fueled flows a technique needs to be developed to remove the interfering polycyclic aromatic hydrocarbon fluorescence from the relatively weak Raman signals. A current application of such a technique is to the analysis of the mixing and combustion performance of multijet, impinging-jet candidate fuel injectors for the baseline Mars ascent engine, which will burn methane and liquid oxygen produced in-situ on Mars to reduce the propellant mass transported to Mars for future manned Mars missions. The present technique takes advantage of the strongly polarized nature of Raman scattering. It is shown to be discernable from unpolarized fluorescence interference by subtracting one polarized image from another. Both of these polarized images are obtained from a single laser pulse by using a polarization-separating calcite rhomb mounted in the imaging spectrograph. A demonstration in a propane-air flame is presented.

Liquid Methane/Liquid Oxygen Injectors for Potential Future Mars Ascent Engines

NASA Technical Reports Server (NTRS)

Trinh, Huu Phuoc

1999-01-01

Preliminary mission studies for human exploration of Mars have been performed at Marshall Space Flight Center (MSFC). These studies indicate that for chemical rockets only a cryogenic propulsion system would provide high enough performance to be considered for a Mars ascent vehicle. Although the mission is possible with Earth-supplied propellants for this vehicle, utilization of in-situ propellants is highly attractive. This option would significantly reduce the overall mass of launch vehicles. Consequently, the cost of the mission would be greatly reduced because the number and size of the Earth launch vehicle(s) needed for the mission would decrease. NASA/Johnson Space Center has initiated several concept studies of in-situ propellant production plants. Liquid oxygen (LOX) is the primary candidate for an in-situ oxidizer. In-situ fuel candidates include methane (CH4), ethylene (C2H4), and methanol (CH3OH). MSFC initiated a technology development program for a cryogenic propulsion system for the Mars human exploration mission in 1998. One part of this technology program is the effort described here: an evaluation of propellant injection concepts for a LOX/liquid methane Mars Ascent Engine (MAE) with an emphasis on light-weight, high efficiency, reliability, and thermal compatibility. In addition to the main objective, hot-fire tests of the subject injectors will be used to test other key technologies including light-weight combustion chamber materials and advanced ignition concepts. This paper will address the results of the liquid methane/LOX injector study conducted at MSFC. A total of four impinging injector configurations were tested under combustion conditions in a modular combustor test article (MCTA), equipped with optically accessible windows. A series of forty hot-fire tests, which covered a wide range of engine operating conditions with the chamber pressure varied from 320 to 510 and the mixture ratio from 1.5 to 3.5, were performed. The test matrix also included a variation in the combustion chamber length for the purpose of investigating its effects on the combustion performance and stability.

Rocket noise - A review

NASA Astrophysics Data System (ADS)

McInerny, S. A.

1990-10-01

This paper reviews what is known about far-field rocket noise from the controlled studies of the late 1950s and 1960s and from launch data. The peak dimensionless frequency, the dependence of overall sound power on exhaust parameters, and the directivity of the overall sound power of rockets are compared to those of subsonic jets and turbo-jets. The location of the dominant sound source in the rocket exhaust plume and the mean flow velocity in this region are discussed and shown to provide a qualitative explanation for the low peak Strouhal number, fD(e)/V(e), and large angle of maximum directivity. Lastly, two empirical prediction methods are compared with data from launches of a Titan family vehicle (two, solid rocket motors of 5.7 x 10 to the 6th N thrust each) and the Saturn V (five, liquid oxygen/rocket propellant engines of 6.7 x 10 to the 6th N thrust, each). The agreement is favorable. In contrast, these methods appear to overpredict the far-field sound pressure levels generated by the Space Shuttle.

Development of the RFBB “Bargouzine” concept for Ariane-5 evolution

NASA Astrophysics Data System (ADS)

Sumin, Yuriy; Kostromin, Sergey F.; Panichkin, Nikolai; Prel, Yves; Osin, Mikhail; Iranzo-Greus, David; Prampolini, Marco

2009-10-01

This paper presents the study of a concept of Ariane-5 evolution by means of replacement of two solid-propellant boosters EAP with two liquid-propellant reusable fly-back boosters (RFBBs) called "Bargouzine". The main design feature of the reference RFBB is LOX/LH2 propellant, the canard aerodynamic configuration with delta wings and rocket engines derived from Vulcain-2 identical to that of the central core except for the nozzle length. After separation RFBBs return back by use of air breathing engines mounted in the aft part and then landing on a runway. The aim of the study is a more detailed investigation of critical technology issues concerning reliability, re-usability and maintenance requirements. The study was performed in three main phases: system trade-off, technical consolidation, and programmatic synthesis. The system trade-off includes comparative analysis of two systems with three and four engines on each RFBB and determination of the necessary thrust level taking into account thrust reservation for emergency situations. Besides, this phase contains trade-off on booster aerodynamic configurations and abort scenario analysis. The second phase includes studying of controllability during the ascent phase and separation, thermo-mechanical design, development of ground interfaces and attachment means, and turbojets engine analysis taking into account reusability.

Modal survey of the space shuttle solid rocket motor using multiple input methods

NASA Technical Reports Server (NTRS)

Brillhart, Ralph; Hunt, David L.; Jensen, Brent M.; Mason, Donald R.

1987-01-01

The ability to accurately characterize propellant in a finite element model is a concern of engineers tasked with studying the dynamic response of the Space Shuttle Solid Rocket Motor (SRM). THe uncertainties arising from propellant characterization through specimem testing led to the decision to perform a model survey and model correlation of a single segment of the Shuttle SRM. Multiple input methods were used to excite and define case/propellant modes of both an inert segment and, later, a live propellant segment. These tests were successful at defining highly damped, flexible modes, several pairs of which occured with frequency spacing of less than two percent.

Performance and technical feasibility comparison of reusable launch systems: A synthesis of the ESA winged launcher studies

NASA Astrophysics Data System (ADS)

Berry, W.; Grallert, H.

1996-02-01

The paper presents a synthesis of the performance and technical feasibility assessment of 7 reusable launcher types, comprising 13 different vehicles, studied by European Industry for ESA in the ESA Winged Launcher Study in the period January 1988 to May 1994. The vehicles comprised single-stage-to-orbit (SSTO) and two-stage-to-orbit (TSTO) vehicles, propelled by either air-breathing/rocket propulsion or entirely by rocket propulsion. The results showed that an SSTO vehicle of the HOTOL-type, propelled by subsonic combustion air-breathing/rocket engines could barely deliver the specified payload mass and was aerodynamically unstable; that a TSTO vehicle of the Saenger type, employing subsonic combustion airbreathing propulsion in its first stage and rocket propulsion in its second stage, could readily deliver the specified payload mass and was found to be technically feasible and versatile; that an SSTO vehicle of the NASP type, propelled by supersonic combustion airbreathing/rocket propulsion was able to deliver a reduced payload mass, was very complex and required very advanced technologies; that an air-launched rocket propelled vehicle of the Interim HOTOL type, although technically feasible, could deliver only a reduced payload mass, being constrained by the lifting capability of the carrier airplane; that three different, entirely rocket-propelled vehicles could deliver the specified payload mass, were technically feasible but required relatively advanced technologies.

Comparison of super-high-energy-propulsion-systems based on metallic hydrogen propellant for ES to LEO space transportation

NASA Technical Reports Server (NTRS)

Thierschmann, M.

1990-01-01

The application is studied of metallic H2 as a rocket propellant, which contains a specific energy of about 52 kcal/g in theory yielding a maximum specific impulse of 1700 s. With the convincing advantage of having a density 14 times that of conventional liquid H2/liquid O2 propellants, metallic H2 could satisfy the demands of advanced launch vehicle propulsion for the next millennium. Provided that there is an atomic metallic state of H2, and that this state is metastable at ambient pressure, which still is not proven, the results are given of the study of some important areas, which concern the production of metallic H2, the combustion, chamber cooling, and storage. The results show that the use of metallic H2 as rocket propellant could lead to revolutionary changes in space vehicle philosophy toward small size, small weight, and high performance single stage to orbit systems. The use of high metallic H2 mass fractions results in a dramatic reduction of required propellant volume, while gas temperatures in the combustion chamber exceed 5000 K. Furthermore, it follows, that H2 (liquid or slush) is the most favorable candidate as working fluid. Jet generated noise due to high exhaust velocities could be a problem.

Combustion Instability in an Acid-Heptane Rocket with a Pressurized-Gas Propellant Pumping System

NASA Technical Reports Server (NTRS)

Tischler, Adelbert O.; Bellman, Donald R.

1951-01-01

Results of experimental measurements of low-frequency combustion instability of a 300-pound thrust acid-heptane rocket engine were compared to the trends predicted by an analysis of combustion instability in a rocket engine with a pressurized-gas propellant pumping system. The simplified analysis, which assumes a monopropellant model, was based on the concept of a combustion the delay occurring from the moment of propellant injection to the moment of propellant combustion. This combustion time delay was experimentally measured; the experimental values were of approximately half the magnitude predicted by the analysis. The pressure-fluctuation frequency for a rocket engine with a characteristic length of 100 inches and operated at a combustion-chamber pressure of 280 pounds per square inch absolute was 38 cycles per second; the analysis indicated. a frequency of 37 cycles per second. Increasing combustion-chamber characteristic length decreased the pressure-fluctuation frequency, in conformity to the analysis. Increasing the chamber operating pressure or increasing the injector pressure drop increased the frequency. These latter two effects are contrary to the analysis; the discrepancies are attributed to the conflict between the assumptions made to simplify the analysis and the experimental conditions. Oxidant-fuel ratio had no apparent effect on the experimentally measured pressure-fluctuation frequency for acid-heptane ratios from 3.0 to 7.0. The frequencies decreased with increased amplitude of the combustion-chamber pressure variations. The analysis indicated that if the combustion time delay were sufficiently short, low-frequency combustion instability would be eliminated.

Status of flow separation prediction in liquid propellant rocket nozzles

NASA Technical Reports Server (NTRS)

Schmucker, R. H.

1974-01-01

Flow separation which plays an important role in the design of a rocket engine nozzle is discussed. For a given ambient pressure, the condition of no flow separation limits the area ratio and, therefore, the vacuum performance. Avoidance of performance loss due to area ratio limitation requires a correct prediction of the flow separation conditions. To provide a better understanding of the flow separation process, the principal behavior of flow separation in a supersonic overexpanded rocket nozzle is described. The hot firing separation tests from various sources are summarized, and the applicability and accuracy of the measurements are described. A comparison of the different data points allows an evaluation of the parameters that affect flow separation. The pertinent flow separation predicting methods, which are divided into theoretical and empirical correlations, are summarized and the numerical results are compared with the experimental points.

Fiber-optic sensing in cryogenic environments. [for rocket propellant tank monitoring

NASA Technical Reports Server (NTRS)

Sharma, M.; Brooks, R. E.

1980-01-01

Passive optical sensors using fiber-optic signal transmission to a remote monitoring station are explored as an alternative to electrical sensors used to monitor the status of explosive propellants. The designs of passive optical sensors measuring liquid level, pressure, and temperature in cryogenic propellant tanks are discussed. Test results for an experimental system incorporating these sensors and operating in liquid nitrogen demonstrate the feasibility of passive sensor techniques and indicate that they can serve as non-hazardous replacements for more conventional measuring equipment in explosive environments.

Low-Cost Propellant Launch to LEO from a Tethered Balloon - 'Propulsion Depots' Not 'Propellant Depots'

NASA Technical Reports Server (NTRS)

Wilcox, Brian H.; Schneider, Evan G.; Vaughan, David A.; Hall, Jeffrey L.; Yu, Chi Yau

2011-01-01

As we have previously reported, it may be possible to launch payloads into low-Earth orbit (LEO) at a per-kilogram cost that is one to two orders of magnitude lower than current launch systems, using only a relatively small capital investment (comparable to a single large present-day launch). An attractive payload would be large quantities of high-performance chemical rocket propellant (e.g. Liquid Oxygen/Liquid Hydrogen (LO2/LH2)) that would greatly facilitate, if not enable, extensive exploration of the moon, Mars, and beyond.

Altitude Starting Tests of a 1000-Pound-Thrust Solid-Propellant Rocket

NASA Technical Reports Server (NTRS)

Sloop, John L.; Rollbuhler, R. James; Krawczonek, Eugene M.

1957-01-01

Four solid-propellant rocket engines of nominal 1000-pound-thrust were tested for starting characteristics at pressure altitudes ranging from 112,500 to 123,000 feet and at a temperature of -75 F. All engines ignited and operated successfully. Average chamber pressures ranged from 1060 to ll90 pounds per square inch absolute with action times from 1.51 to 1.64 seconds and ignition delays from 0.070 t o approximately 0.088 second. The chamber pressures and action times were near the specifications, but the ignition delay was almost twice the specified value of 0.040 second.

NMR imaging and hydrodynamic analysis of neutrally buoyant non-Newtonian slurry flows

NASA Astrophysics Data System (ADS)

Bouillard, J. X.; Sinton, S. W.

The flow of solids loaded suspension in cylindrical pipes has been the object of intense experimental and theoretical investigations in recent years. These types of flows are of great interest in chemical engineering because of their important use in many industrial manufacturing processes. Such flows are for example encountered in the manufacture of solid-rocket propellants, advanced ceramics, reinforced polymer composites, in heterogeneous catalytic reactors, and in the pipeline transport of liquid-solids suspensions. In most cases, the suspension microstructure and the degree of solids dispersion greatly affect the final performance of the manufactured product. For example, solid propellant pellets need to be extremely-well dispersed in gel matrices for use as rocket engine solid fuels. The homogeneity of pellet dispersion is critical to allow good uniformity of the burn rate, which in turn affects the final mechanical performance of the engine. Today's manufacturing of such fuels uses continuous flow processes rather than batch processes. Unfortunately, the hydrodynamics of such flow processes is poorly understood and is difficult to assess because it requires the simultaneous measurements of liquid/solids phase velocities and volume fractions. Due to the recent development in pulsed Fourier Transform NMR imaging, NMR imaging is now becoming a powerful technique for the non intrusive investigation of multi-phase flows. This paper reports and exposes a state-of-the-art experimental and theoretical methodology that can be used to study such flows. The hydrodynamic model developed for this study is a two-phase flow shear thinning model with standard constitutive fluid/solids interphase drag and solids compaction stresses. this model shows good agreement with experimental data and the limitations of this model are discussed.

Evaluation and Improvement of Liquid Propellant Rocket Chugging Analysis Techniques. Part 2: a Study of Low Frequency Combustion Instability in Rocket Engine Preburners Using a Heterogeneous Stirred Tank Reactor Model. Final Report M.S. Thesis - Aug. 1987

NASA Technical Reports Server (NTRS)

Bartrand, Timothy A.

1988-01-01

During the shutdown of the space shuttle main engine, oxygen flow is shut off from the fuel preburner and helium is used to push the residual oxygen into the combustion chamber. During this process a low frequency combustion instability, or chug, occurs. This chug has resulted in damage to the engine's augmented spark igniter due to backflow of the contents of the preburner combustion chamber into the oxidizer feed system. To determine possible causes and fixes for the chug, the fuel preburner was modeled as a heterogeneous stirred tank combustion chamber, a variable mass flow rate oxidizer feed system, a constant mass flow rate fuel feed system and an exit turbine. Within the combustion chamber gases were assumed perfectly mixed. To account for liquid in the combustion chamber, a uniform droplet distribution was assumed to exist in the chamber, with mean droplet diameter determined from an empirical relation. A computer program was written to integrate the resulting differential equations. Because chamber contents were assumed perfectly mixed, the fuel preburner model erroneously predicted that combustion would not take place during shutdown. The combustion rate model was modified to assume that all liquid oxygen that vaporized instantaneously combusted with fuel. Using this combustion model, the effect of engine parameters on chamber pressure oscillations during the SSME shutdown was calculated.

Test Planning Approach and Lessons

NASA Technical Reports Server (NTRS)

Parkinson, Douglas A.; Brown, Kendall K.

2004-01-01

As NASA began technology risk reduction activities and planning for the next generation launch vehicle under the Space Launch Initiative (SLI), now the Next Generation Launch Technology (NGLT) Program, a review of past large liquid rocket engine development programs was performed. The intent of the review was to identify any significant lessons from the development testing programs that could be applied to current and future engine development programs. Because the primary prototype engine in design at the time of this study was the Boeing-Rocketdyne RS-84, the study was slightly biased towards LOX/RP-1 liquid propellant engines. However, the significant lessons identified are universal. It is anticipated that these lessons will serve as a reference for test planning in the Engine Systems Group at Marshall Space Flight Center (MSFC). Towards the end of F-1 and J-2 engine development testing, NASA/MSFC asked Rocketdyne to review those test programs. The result was a document titled, Study to Accelerate Development by Test of a Rocket Engine (R-8099). The "intent (of this study) is to apply this thinking and learning to more efficiently develop rocket engines to high reliability with improved cost effectivenes" Additionally, several other engine programs were reviewed - such as SSME, NSTS, STME, MC-1, and RS-83- to support or refute the R-8099. R-8099 revealed two primary lessons for test planning, which were supported by the other engine development programs. First, engine development programs can benefit from arranging the test program for engine system testing as early as feasible. The best test for determining environments is at the system level, the closest to the operational flight environment. Secondly, the component testing, which tends to be elaborate, should instead be geared towards reducing risk to enable system test. Technical risk can be reduced at the component level, but the design can only be truly verified and validated after engine system testing.

User's manual for rocket combustor interactive design (ROCCID) and analysis computer program. Volume 2: Appendixes A-K

NASA Technical Reports Server (NTRS)

Muss, J. A.; Nguyen, T. V.; Johnson, C. W.

1991-01-01

The appendices A-K to the user's manual for the rocket combustor interactive design (ROCCID) computer program are presented. This includes installation instructions, flow charts, subroutine model documentation, and sample output files. The ROCCID program, written in Fortran 77, provides a standardized methodology using state of the art codes and procedures for the analysis of a liquid rocket engine combustor's steady state combustion performance and combustion stability. The ROCCID is currently capable of analyzing mixed element injector patterns containing impinging like doublet or unlike triplet, showerhead, shear coaxial and swirl coaxial elements as long as only one element type exists in each injector core, baffle, or barrier zone. Real propellant properties of oxygen, hydrogen, methane, propane, and RP-1 are included in ROCCID. The properties of other propellants can be easily added. The analysis models in ROCCID can account for the influences of acoustic cavities, helmholtz resonators, and radial thrust chamber baffles on combustion stability. ROCCID also contains the logic to interactively create a combustor design which meets input performance and stability goals. A preliminary design results from the application of historical correlations to the input design requirements. The steady state performance and combustion stability of this design is evaluated using the analysis models, and ROCCID guides the user as to the design changes required to satisfy the user's performance and stability goals, including the design of stability aids. Output from ROCCID includes a formatted input file for the standardized JANNAF engine performance prediction procedure.

Engine throat/nozzle optics for plume spectroscopy

NASA Technical Reports Server (NTRS)

Bickford, R. L.; Duncan, D. B.

1991-01-01

The Task 2.0 Engine Throat/Nozzle Optics for Plume Spectroscopy, effort was performed under the NASA LeRC Development of Life Prediction Capabilities for Liquid Propellant Rocket Engines program. This Task produced the engineering design of an optical probe to enable spectroscopic measurements within the SSME main chamber. The probe mounts on the SSME nozzle aft manifold and collects light emitted from the throat plane and chamber. Light collected by the probe is transferred to a spectrometer through a fiber optic cable. The design analyses indicate that the probe will function throughout the engine operating cycle and is suitable for both test stand and flight operations. By detecting metallic emissions that are indicative of component degradation or incipient failure, engine shutdown can be initiated before catastrophic failure. This capability will protect valuable test stand hardware and provide enhanced mission safety.

Parametric Modeling for Fluid Systems

NASA Technical Reports Server (NTRS)

Pizarro, Yaritzmar Rosario; Martinez, Jonathan

2013-01-01

Fluid Systems involves different projects that require parametric modeling, which is a model that maintains consistent relationships between elements as is manipulated. One of these projects is the Neo Liquid Propellant Testbed, which is part of Rocket U. As part of Rocket U (Rocket University), engineers at NASA's Kennedy Space Center in Florida have the opportunity to develop critical flight skills as they design, build and launch high-powered rockets. To build the Neo testbed; hardware from the Space Shuttle Program was repurposed. Modeling for Neo, included: fittings, valves, frames and tubing, between others. These models help in the review process, to make sure regulations are being followed. Another fluid systems project that required modeling is Plant Habitat's TCUI test project. Plant Habitat is a plan to develop a large growth chamber to learn the effects of long-duration microgravity exposure to plants in space. Work for this project included the design and modeling of a duct vent for flow test. Parametric Modeling for these projects was done using Creo Parametric 2.0.

Development of Detonation Modeling Capabilities for Rocket Test Facilities: Hydrogen-Oxygen-Nitrogen Mixtures

NASA Technical Reports Server (NTRS)

Allgood, Daniel C.

2016-01-01

The objective of the presented work was to develop validated computational fluid dynamics (CFD) based methodologies for predicting propellant detonations and their associated blast environments. Applications of interest were scenarios relevant to rocket propulsion test and launch facilities. All model development was conducted within the framework of the Loci/CHEM CFD tool due to its reliability and robustness in predicting high-speed combusting flow-fields associated with rocket engines and plumes. During the course of the project, verification and validation studies were completed for hydrogen-fueled detonation phenomena such as shock-induced combustion, confined detonation waves, vapor cloud explosions, and deflagration-to-detonation transition (DDT) processes. The DDT validation cases included predicting flame acceleration mechanisms associated with turbulent flame-jets and flow-obstacles. Excellent comparison between test data and model predictions were observed. The proposed CFD methodology was then successfully applied to model a detonation event that occurred during liquid oxygen/gaseous hydrogen rocket diffuser testing at NASA Stennis Space Center.

Advanced transportation system studies. Alternate propulsion subsystem concepts: Propulsion database

NASA Technical Reports Server (NTRS)

Levack, Daniel

1993-01-01

The Advanced Transportation System Studies alternate propulsion subsystem concepts propulsion database interim report is presented. The objective of the database development task is to produce a propulsion database which is easy to use and modify while also being comprehensive in the level of detail available. The database is to be available on the Macintosh computer system. The task is to extend across all three years of the contract. Consequently, a significant fraction of the effort in this first year of the task was devoted to the development of the database structure to ensure a robust base for the following years' efforts. Nonetheless, significant point design propulsion system descriptions and parametric models were also produced. Each of the two propulsion databases, parametric propulsion database and propulsion system database, are described. The descriptions include a user's guide to each code, write-ups for models used, and sample output. The parametric database has models for LOX/H2 and LOX/RP liquid engines, solid rocket boosters using three different propellants, a hybrid rocket booster, and a NERVA derived nuclear thermal rocket engine.

Developments in REDES: The rocket engine design expert system

NASA Technical Reports Server (NTRS)

Davidian, Kenneth O.

1990-01-01

The Rocket Engine Design Expert System (REDES) is being developed at the NASA-Lewis to collect, automate, and perpetuate the existing expertise of performing a comprehensive rocket engine analysis and design. Currently, REDES uses the rigorous JANNAF methodology to analyze the performance of the thrust chamber and perform computational studies of liquid rocket engine problems. The following computer codes were included in REDES: a gas properties program named GASP, a nozzle design program named RAO, a regenerative cooling channel performance evaluation code named RTE, and the JANNAF standard liquid rocket engine performance prediction code TDK (including performance evaluation modules ODE, ODK, TDE, TDK, and BLM). Computational analyses are being conducted by REDES to provide solutions to liquid rocket engine thrust chamber problems. REDES is built in the Knowledge Engineering Environment (KEE) expert system shell and runs on a Sun 4/110 computer.

Developments in REDES: The Rocket Engine Design Expert System

NASA Technical Reports Server (NTRS)

Davidian, Kenneth O.

1990-01-01

The Rocket Engine Design Expert System (REDES) was developed at NASA-Lewis to collect, automate, and perpetuate the existing expertise of performing a comprehensive rocket engine analysis and design. Currently, REDES uses the rigorous JANNAF methodology to analyze the performance of the thrust chamber and perform computational studies of liquid rocket engine problems. The following computer codes were included in REDES: a gas properties program named GASP; a nozzle design program named RAO; a regenerative cooling channel performance evaluation code named RTE; and the JANNAF standard liquid rocket engine performance prediction code TDK (including performance evaluation modules ODE, ODK, TDE, TDK, and BLM). Computational analyses are being conducted by REDES to provide solutions to liquid rocket engine thrust chamber problems. REDES was built in the Knowledge Engineering Environment (KEE) expert system shell and runs on a Sun 4/110 computer.

16 CFR § 1500.85 - Exemptions from classification as banned hazardous substances.

Code of Federal Regulations, 2013 CFR

2013-01-01

... component has no hazards other than being in a self-pressurized container. (8) Model rocket propellant devices designed for use in light-weight, recoverable, and reflyable model rockets, provided such devices... recovery system activation devices intended for use with premanufactured model rocket engines wherein all...

16 CFR 1500.85 - Exemptions from classification as banned hazardous substances.

Code of Federal Regulations, 2014 CFR

2014-01-01

... component has no hazards other than being in a self-pressurized container. (8) Model rocket propellant devices designed for use in light-weight, recoverable, and reflyable model rockets, provided such devices... recovery system activation devices intended for use with premanufactured model rocket engines wherein all...

16 CFR 1500.85 - Exemptions from classification as banned hazardous substances.

Code of Federal Regulations, 2011 CFR

2011-01-01

... component has no hazards other than being in a self-pressurized container. (8) Model rocket propellant devices designed for use in light-weight, recoverable, and reflyable model rockets, provided such devices... recovery system activation devices intended for use with premanufactured model rocket engines wherein all...

A Theoretical Evaluation of Secondary Atomization Effects on Engine Performance for Aluminum Gel Propellants

NASA Technical Reports Server (NTRS)

Mueller, D. C.; Turns, S. R.

1994-01-01

A one-dimensional model of a gel-fueled rocket combustion chamber has been developed. This model includes the processes of liquid hydrocarbon burnout, secondary atomization. aluminum ignition, and aluminum combustion. Also included is a model of radiative heat transfer from the solid combustion products to the chamber walls. Calculations indicate that only modest secondary atomization is required to significantly reduce propellant burnout distances, aluminum oxide residual size and radiation heat wall losses. Radiation losses equal to approximately 2-13 percent of the energy released during combustion were estimated. A two-dimensional, two-phase nozzle code was employed to estimate radiation and nozzle two-phase flow effects on overall engine performance. Radiation losses yielded a 1 percent decrease in engine I(sub sp). Results also indicate that secondary atomization may have less effect on two-phase losses than it does on propellant burnout distance and no effect if oxide particle coagulation and shear induced droplet breakup govern oxide particle size. Engine I(sub sp) was found to decrease from 337.4 to 293.7 seconds as gel aluminum mass loading was varied from 0-70 wt percent. Engine I(sub sp) efficiencies, accounting for radiation and two-phase flow effects, on the order of 0.946 were calculated for a 60 wt percent gel, assuming a fragmentation ratio of 5.

Early Rockets

NASA Image and Video Library

1953-08-30

U.S. Army Redstone Rocket: The Redstone ballistic missile was a high-accuracy, liquid-propelled, surface-to-surface missile developed by the Army Ballistic Missile Agency, Redstone Arsenal, in Huntsville, Alabama, under the direction of Dr. von Braun. The Redstone rocket was also known as "Old Reliable" because of its many diverse missions. The first Redstone Missile was launched from Cape Canaveral, Florida on August 30, 1953.

Gas-dynamic modeling of gas flow in semi-closed space including channel surface fluctuation

NASA Astrophysics Data System (ADS)

Petrova, E. N.; Salnikov, A. F.

2016-10-01

In this article frequency interaction conditions, that affect on acoustic stability of solid-propellant rocket engine (SPRE) action, and its influence on level change of pressure fluctuations with longitudinal gas oscillations in the combustion chamber (CC) are considered. Studies of CC in the assessment of the operating rocket engine stability are reported.

Cryogenic Gellant and Fuel Formulation for Metallized Gelled Propellants: Hydrocarbons and Hydrogen with Aluminum

NASA Technical Reports Server (NTRS)

Wong, Wing; Starkovich, John; Adams, Scott; Palaszewski, Bryan; Davison, William; Burt, William; Thridandam, Hareesh; Hu-Peng, Hsiao; Santy, Myrrl J.

1994-01-01

An experimental program to determine the viability of nanoparticulate gellant materials for gelled hydrocarbons and gelled liquid hydrogen was conducted. The gellants included alkoxides (BTMSE and BTMSH) and silica-based materials. Hexane, ethane, propane and hydrogen were gelled with the newly-formulated materials and their rheological properties were determined: shear stress versus shear rate and their attendant viscosities. Metallized hexane with aluminum particles was also rheologically characterized. The propellant and gellant formulations were selected for the very high surface area and relatively-high energy content of the gellants. These new gellants can therefore improve rocket engine specific impulse over that obtained with traditional cryogenic-fuel gellant materials silicon dioxide, frozen methane, or frozen ethane particles. Significant reductions in the total mass of the gellant were enabled in the fuels. In gelled liquid hydrogen, the total mass of gellant was reduced from 10-40 wt percent of frozen hydrocarbon particles to less that 8 wt percent with the alkoxide.

Romanian MRE Rocket Engines Program - An Early Endeavor

NASA Astrophysics Data System (ADS)

Rugescu, R. E.

2002-01-01

(MRE) was initiated in the years '60 of the past century at the Chair of Aerospace Sciences "Elie Carafoli" from the "Politehnica" University in Bucharest (PUB). Consisting of theoretical and experimental investigations in the form of computational methods and technological solutions for small size MRE-s and the concept of the test stand for these engines, the program ended in the construction of the first Romanian liquid rocket motors. Hermann Oberth and Dorin Pavel, were known from 1923, no experimental practice was yet tempted, at the time level of 1960. It was the intention of the developers at PUB to cover this gap and initiate a feasible, low-cost, demonstrative program of designing and testing experimental models of MRE. The research program was oriented towards future development of small size space carrier vehicles for scientific applications only, as an independent program with no connection to other defense programs imagined by the authorities in Bucharest, at that time. Consequently the entire financial support was assured by "Politehnica" university. computerized methods in the thermochemistry of heterogeneous combustion, for both steady and unsteady flows with chemical reactions and two phase flows. The research was gradually extended to the production of a professional CAD program for steady-state heat transfer simulations and the loading capacity analyses of the double wall, cooled thrust chamber. The resulting computer codes were run on a 360-30 IMB machine, beginning in 1968. Some of the computational methods were first exposed at the 9th International Conference on Applied Mechanics, held in Bucharest between June 23-27, 1969. hot testing of a series of storable propellant, variable thrust, variable geometry, liquid rocket motors, with a maximal thrust of 200N. A remotely controlled, portable test bad, actuated either automatically or manually and consisting of a 6-modules construction was built for this motor series, with a simple 8 analog-channel and 5 digital-channel data measuring and recording system. The first hot test firing of the MRE-1B motor took place successfully on April 9th, 1969 in Bucharest, at the "Elie Carafoli" Chair of UPB. The research program continued with the development of a series of solid, double base propellant rocket and ram-rocket motors, with emphasize on the optimization of the gasdynamic contour of the engine, in order to increase the flight performances. Increments of up to 8% in specific thrust were measured on the test stand, with mass savings and no extra costs. The test firing of the first Romanian, air-breathing ram-rocket engine took place successfully in august 1987 at the Chemical Works in Fagaras, Romania. Astronautics", founded in Bucharest. The principles and history of the "MRE" research program are presented in the proposed paper.

Tail Service Mast Umbilical Arrival

NASA Image and Video Library

2016-08-02

A heavy-lift transport truck arrives at the Launch Equipment Test Facility (LETF) at NASA’s Kennedy Space Center in Florida, with the first of two Tail Service Mast Umbilicals (TSMU) for NASA’s Space Launch System (SLS). Two TSMUs will provide liquid propellants and power to the Space Launch System (SLS) rocket’s core stage engine. Both TSMUs will connect to the zero-level deck on the mobile launcher, providing fuel and electricity to the SLS rocket before it launches on Exploration Mission 1. The TSMU will undergo testing and validation at the LETF to verify it is functioning properly. The center’s Engineering Directorate and the Ground Systems Development and Operations Program are overseeing processing and testing of the umbilicals.

Tail Service Mast Umbilical Arrival

NASA Image and Video Library

2016-08-02

Technicians assist as a crane is used to lift the first Tail Service Mast Umbilical (TSMU) into the vertical position at the Launch Equipment Test Facility (LETF) at NASA’s Kennedy Space Center in Florida. Two TSMUs will provide liquid propellants and power to the Space Launch System (SLS) rocket’s core stage engine. Both TSMUs will connect to the zero-level deck on the mobile launcher, providing fuel and electricity to the SLS rocket before it launches on Exploration Mission 1. The TSMU will undergo testing and validation at the LETF to verify it is functioning properly. The center’s Engineering Directorate and the Ground Systems Development and Operations Program are overseeing processing and testing of the umbilicals.

Tail Service Mast Umbilical Arrival

NASA Image and Video Library

2016-08-02

A crane is prepared to help lift the first Tail Service Mast Umbilical (TSMU) for NASA’s Space Launch System (SLS) at the Launch Equipment Test Facility (LETF) at NASA’s Kennedy Space Center in Florida. Two TSMUs will provide liquid propellants and power to the Space Launch System (SLS) rocket’s core stage engine. Both TSMUs will connect to the zero-level deck on the mobile launcher, providing fuel and electricity to the SLS rocket before it launches on Exploration Mission 1. The TSMU will undergo testing and validation at the LETF to verify it is functioning properly. The center’s Engineering Directorate and the Ground Systems Development and Operations Program are overseeing processing and testing of the umbilicals.

Tail Service Mast Umbilical Arrival

NASA Image and Video Library

2016-08-02

A crane is attached to the first Tail Service Mast Umbilical (TSMU) for NASA’s Space Launch System (SLS) at the Launch Equipment Test Facility at NASA’s Kennedy Space Center in Florida. Two TSMUs will provide liquid propellants and power to the Space Launch System (SLS) rocket’s core stage engine. Both TSMUs will connect to the zero-level deck on the mobile launcher, providing fuel and electricity to the SLS rocket before it launches on Exploration Mission 1. The TSMU will undergo testing and validation at the LETF to verify it is functioning properly. The center’s Engineering Directorate and the Ground Systems Development and Operations Program are overseeing processing and testing of the umbilicals.

Tail Service Mast Umbilical Arrival

NASA Image and Video Library

2016-08-02

A crane lifts the first Tail Service Mast Umbilical (TSMU) up for placement on a test stand at the Launch Equipment Test Facility at NASA’s Kennedy Space Center in Florida. Two TSMUs will provide liquid propellants and power to the Space Launch System (SLS) rocket’s core stage engine. Both TSMUs will connect to the zero-level deck on the mobile launcher, providing fuel and electricity to the SLS rocket before it launches on Exploration Mission 1. The TSMU will undergo testing and validation at the LETF to verify it is functioning properly. The center’s Engineering Directorate and the Ground Systems Development and Operations Program are overseeing processing and testing of the umbilicals.

Tail Service Mast Umbilical Arrival

NASA Image and Video Library

2016-08-02

Technician monitors the progress as a crane lowers the first Tail Service Mast Umbilical (TSMU) onto a test stand at the Launch Equipment Test Facility at NASA’s Kennedy Space Center in Florida. Two TSMUs will provide liquid propellants and power to the Space Launch System (SLS) rocket’s core stage engine. Both TSMUs will connect to the zero-level deck on the mobile launcher, providing fuel and electricity to the SLS rocket before it launches on Exploration Mission 1. The TSMU will undergo testing and validation at the LETF to verify it is functioning properly. The center’s Engineering Directorate and the Ground Systems Development and Operations Program are overseeing processing and testing of the umbilicals.

Tail Service Mast Umbilical Arrival

NASA Image and Video Library

2016-08-02

A technician monitors the progress as a crane lifts the first Tail Service Mast Umbilical (TSMU) for transfer to a test stand at the Launch Equipment Test Facility at NASA’s Kennedy Space Center in Florida. Two TSMUs will provide liquid propellants and power to the Space Launch System (SLS) rocket’s core stage engine. Both TSMUs will connect to the zero-level deck on the mobile launcher, providing fuel and electricity to the SLS rocket before it launches on Exploration Mission 1. The TSMU will undergo testing and validation at the LETF to verify it is functioning properly. The center’s Engineering Directorate and the Ground Systems Development and Operations Program are overseeing processing and testing of the umbilicals.

Analysis of a Nuclear Enhanced Airbreathing Rocket for Earth to Orbit Applications

NASA Technical Reports Server (NTRS)

Adams, Robert B.; Landrum, D. Brian; Brown, Norman (Technical Monitor)

2001-01-01

The proposed engine concept is the Nuclear Enhanced Airbreathing Rocket (NEAR). The NEAR concept uses a fission reactor to thermally heat a propellant in a rocket plenum. The rocket is shrouded, thus the exhaust mixes with ingested air to provide additional thermal energy through combustion. The combusted flow is then expanded through a nozzle to provide thrust.

Early Rockets

NASA Image and Video Library

1950-01-01

Test firing of a Redstone Missile at Redstone Test Stand in the early 1950's. The Redstone was a high-accuracy, liquid-propelled, surface-to-surface missile developed by the von Braun Team under the management of the U.S. Army. The Redstone was the first major rocket development program in the United States.

Low heat transfer oxidizer heat exchanger design and analysis

NASA Technical Reports Server (NTRS)

Kanic, P. G.; Kmiec, T. D.; Peckham, R. J.

1987-01-01

The RL10-IIB engine, a derivative of the RLIO, is capable of multi-mode thrust operation. This engine operates at two low thrust levels: tank head idle (THI), which is approximately 1 to 2 percent of full thrust, and pumped idle (PI), which is 10 percent of full thrust. Operation at THI provides vehicle propellant settling thrust and efficient engine thermal conditioning; PI operation provides vehicle tank pre-pressurization and maneuver thrust for log-g deployment. Stable combustion of the RL10-IIB engine at THI and PI thrust levels can be accomplished by providing gaseous oxygen at the propellant injector. Using gaseous hydrogen from the thrust chamber jacket as an energy source, a heat exchanger can be used to vaporize liquid oxygen without creating flow instability. This report summarizes the design and analysis of a United Aircraft Products (UAP) low-rate heat transfer heat exchanger concept for the RL10-IIB rocket engine. The design represents a second iteration of the RL10-IIB heat exchanger investigation program. The design and analysis of the first heat exchanger effort is presented in more detail in NASA CR-174857. Testing of the previous design is detailed in NASA CR-179487.

Design and evaluation of high performance rocket engine injectors for use with hydrocarbon fuels

NASA Technical Reports Server (NTRS)

Pavli, A. J.

1979-01-01

The feasibility of using a heavy hydrocarbon fuel as a rocket propellant is examined. A method of predicting performance of a heavy hydrocarbon in terms of vaporization effectiveness is described and compared to other fuels and to experimental test results. Experiments were done at a chamber pressure of 4137 KN/sq M (600 psia) with RP-1, JP-10, and liquefied natural gas as fuels, and liquid oxygen as the oxidizer. Combustion length effects were explored over a range of 21.6 cm (8 1/2 in) to 55.9 cm (22 in). Four injector types were tested, each over a range of mixture ratios. Further configuration modifications were obtained by reaming each injector several times to provide test data over a range of injector pressure drop.

Grooved Fuel Rings for Nuclear Thermal Rocket Engines

NASA Technical Reports Server (NTRS)

Emrich, William

2009-01-01

An alternative design concept for nuclear thermal rocket engines for interplanetary spacecraft calls for the use of grooved-ring fuel elements. Beyond spacecraft rocket engines, this concept also has potential for the design of terrestrial and spacecraft nuclear electric-power plants. The grooved ring fuel design attempts to retain the best features of the particle bed fuel element while eliminating most of its design deficiencies. In the grooved ring design, the hydrogen propellant enters the fuel element in a manner similar to that of the Particle Bed Reactor (PBR) fuel element.

Advanced Transportation System Studies. Technical Area 3: Alternate Propulsion Subsystem Concepts. Volume 1; Executive Summary

NASA Technical Reports Server (NTRS)

Levack, Daniel J. H.

2000-01-01

The Alternate Propulsion Subsystem Concepts contract had seven tasks defined that are reported under this contract deliverable. The tasks were: FAA Restart Study, J-2S Restart Study, Propulsion Database Development. SSME Upper Stage Use. CERs for Liquid Propellant Rocket Engines. Advanced Low Cost Engines, and Tripropellant Comparison Study. The two restart studies, F-1A and J-2S, generated program plans for restarting production of each engine. Special emphasis was placed on determining changes to individual parts due to obsolete materials, changes in OSHA and environmental concerns, new processes available, and any configuration changes to the engines. The Propulsion Database Development task developed a database structure and format which is easy to use and modify while also being comprehensive in the level of detail available. The database structure included extensive engine information and allows for parametric data generation for conceptual engine concepts. The SSME Upper Stage Use task examined the changes needed or desirable to use the SSME as an upper stage engine both in a second stage and in a translunar injection stage. The CERs for Liquid Engines task developed qualitative parametric cost estimating relationships at the engine and major subassembly level for estimating development and production costs of chemical propulsion liquid rocket engines. The Advanced Low Cost Engines task examined propulsion systems for SSTO applications including engine concept definition, mission analysis. trade studies. operating point selection, turbomachinery alternatives, life cycle cost, weight definition. and point design conceptual drawings and component design. The task concentrated on bipropellant engines, but also examined tripropellant engines. The Tripropellant Comparison Study task provided an unambiguous comparison among various tripropellant implementation approaches and cycle choices, and then compared them to similarly designed bipropellant engines in the SSTO mission This volume overviews each of the tasks giving its objectives, main results. and conclusions. More detailed Final Task Reports are available on each individual task.

Theoretical performance of some rocket propellants containing hydrogen, nitrogen, and oxygen

NASA Technical Reports Server (NTRS)

Miller, Riley O; Ordin, Paul M

1948-01-01

Theoretical performance data including nozzle-exit temperature, specific impulse, volume specific impulse and composition, temperature, and mean molecular weight of reaction products based on frozen equilibrium and isentropic expansion are presented for 13 propellant combinations at reaction pressure of 300 pounds per square inch absolute and expansion ratio of 20.4. On basis of maximum specific impulse alone, five fuels had the following order for any given oxidant: liquid hydrogen, hydrazine, liquid ammonia, and either hydrazine hydrate or hydroxylamine. Three oxidants with a given fuel had the following order: liquid ozone, liquid oxygen, and 100-percent hydrogen peroxide.

Laser Ignition Technology for Bi-Propellant Rocket Engine Applications

NASA Technical Reports Server (NTRS)

Thomas, Matt; Bossard, John; Early, Jim; Trinh, Huu; Dennis, Jay; Turner, James (Technical Monitor)

2001-01-01

This viewgraph presentation gives an overview of laser ignition technology for bipropellant rocket engines applications. The objectives of this project include: (1) the selection test chambers and flows; (2) definition of the laser ignition setup; (3) pulse format optimization; (4) fiber optic coupled laser ignition system analysis; and (5) chamber integration issues definition. The testing concludes that rocket combustion chamber laser ignition is imminent. Support technologies (multiplexing, window durability/cleaning, and fiber optic durability) are feasible.

A methodology to study the possible occurrence of chugging in liquid rocket engines during transient start-up

NASA Astrophysics Data System (ADS)

Leonardi, Marco; Nasuti, Francesco; Di Matteo, Francesco; Steelant, Johan

2017-10-01

An investigation on the low frequency combustion instabilities due to the interaction of combustion chamber and feed line dynamics in a liquid rocket engine is carried out implementing a specific module in the system analysis software EcosimPro. The properties of the selected double time lag model are identified according to the two classical assumptions of constant and variable time lag. Module capabilities are evaluated on a literature experimental set up consisting of a combustion chamber decoupled from the upstream feed lines. The computed stability map results to be in good agreement with both experimental data and analytical models. Moreover, the first characteristic frequency of the engine is correctly predicted, giving confidence on the use of the module for the analysis of chugging instabilities. As an example of application, a study is carried out on the influence of the feed lines on the system stability, correctly capturing that the lines extend the stable regime of the combustion chamber and that the propellant domes play a key role in coupling the dynamics of combustion chamber and feed lines. A further example is presented to discuss on the role of pressure growth rate and of the combustion chamber properties on the possible occurrence of chug instability during engine start-up and on the conditions that lead to its damping or growth.

A detailed numerical simulation of a liquid-propellant rocket engine ground test experiment

NASA Astrophysics Data System (ADS)

Lankford, D. W.; Simmons, M. A.; Heikkinen, B. D.

1992-07-01

A computational simulation of a Liquid Rocket Engine (LRE) ground test experiment was performed using two modeling approaches. The results of the models were compared with selected data to assess the validity of state-of-the-art computational tools for predicting the flowfield and radiative transfer in complex flow environments. The data used for comparison consisted of in-band station radiation measurements obtained in the near-field portion of the plume exhaust. The test article was a subscale LRE with an afterbody, resulting in a large base region. The flight conditions were such that afterburning regions were observed in the plume flowfield. A conventional standard modeling approach underpredicted the extent of afterburning and the associated radiation levels. These results were attributed to the absence of the base flow region which is not accounted for in this model. To assess the effects of the base region a Navier-Stokes model was applied. The results of this calculation indicate that the base recirculation effects are dominant features in the immediate expansion region and resulted in a much improved comparison. However, the downstream in-band station radiation data remained underpredicted by this model.

REIMR - A Process for Utilizing Liquid Rocket Propulsion-Oriented 'Lessons Learned' to Mitigate Development Risk in Nuclear Thermal Propulsion

NASA Technical Reports Server (NTRS)

Ballard, RIchard O.

2006-01-01

This paper is a summary overview of a study conducted at the NASA Marshall Space Flight Center (NASA 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). One of the primary 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 propulsion system development efforts, such as that which is currently envisioned in the field of nuclear thermal propulsion (NTF). This paper will discuss 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 in the development of an NTP system

Determination of the Flow Field in the Propellant Tank of a Rocket Engine on Completion of the Mission

NASA Astrophysics Data System (ADS)

Fedorov, A. V.; Bedarev, I. A.; Lavruk, S. A.; Trushlyakov, V. I.; Kudentsov, V. Yu.

2018-03-01

In the present work, a method of mathematical simulation is employed to describe processes occurring in the specimens of new equipment and using the remaining propellant in rocket-engine tanks. Within the framework of certain turbulence models, the authors perform a calculation of the flow field in the volume of the tank of the launch-vehicle stage when a hot gas jet is injected into it. A vortex flow structure is revealed; the characteristics of heat transfer for different angles of injection of the jet are determined. The obtained correlation Nu = Nu(Re) satisfactorily describes experimental data.

Determination of the Flow Field in the Propellant Tank of a Rocket Engine on Completion of the Mission

NASA Astrophysics Data System (ADS)

Fedorov, A. V.; Bedarev, I. A.; Lavruk, S. A.; Trushlyakov, V. I.; Kudentsov, V. Yu.

2018-05-01

In the present work, a method of mathematical simulation is employed to describe processes occurring in the specimens of new equipment and using the remaining propellant in rocket-engine tanks. Within the framework of certain turbulence models, the authors perform a calculation of the flow field in the volume of the tank of the launch-vehicle stage when a hot gas jet is injected into it. A vortex flow structure is revealed; the characteristics of heat transfer for different angles of injection of the jet are determined. The obtained correlation Nu = Nu(Re) satisfactorily describes experimental data.

Wing-Nacelle-Propeller Tests - Comparative Tests of Liquid-Cooled and Air-Cooled Engine Nacelles

NASA Technical Reports Server (NTRS)

Wood, Donald H.

1934-01-01

This report gives the results of measurements of the lift, drag, and propeller characteristics of several wing and nacelle combinations with a tractor propeller. The nacelles were so located that the propeller was about 31% of the wing chord directly ahead of the leading edge of the wing, a position which earlier tests (NASA Report No. 415) had shown to be efficient. The nacelles were scale models of an NACA cowled nacelle for a radial air-cooled engine, a circular nacelle with the V-type engine located inside and the radiator for the cooling liquid located inside and the radiator for the type, and a nacelle shape simulating the housing which would be used for an extension shaft if the engine were located entirely within the wing. The propeller used in all cases was a 4-foot model of Navy No. 4412 adjustable metal propeller. The results of the tests indicate that, at the angles of attack corresponding to high speeds of flight, there is no marked advantage of one type of nacelle over the others as far as low drag is concerned, since the drag added by any of the nacelles in the particular location ahead of the wing is very small. The completely cowled nacelle for a radial air-cooled engine appears to have the highest drag, the liquid-cooled engine appears to have the highest drag, the liquid-cooled engine nacelle with external radiator slightly less drag. The liquid-cooled engine nacelle with radiator in the cowling hood has about half the drag of the cowled radial air-cooled engine nacelle. The extension-shaft housing shows practically no increase in drag over that of the wing alone. A large part of the drag of the liquid-cooled engine nacelle appears to be due to the external radiator. The maximum propulsive efficiency for a given propeller pitch setting is about 2% higher for the liquid-cooled engine nacelle with the radiator in the cowling hood than that for the other cowling arrangements.

Application of Chaboche Model in Rocket Thrust Chamber Analysis

NASA Astrophysics Data System (ADS)

Asraff, Ahmedul Kabir; Suresh Babu, Sheela; Babu, Aneena; Eapen, Reeba

2017-06-01

Liquid Propellant Rocket Engines are commonly used in space technology. Thrust chamber is one of the most important subsystems of a rocket engine. The thrust chamber generates propulsive thrust force for flight of the rocket by ejection of combustion products at supersonic speeds. Often double walled construction is employed for these chambers. The thrust chamber investigated here has its hot inner wall fabricated out of a high thermal conductive material like copper alloy and outer wall made of stainless steel. Inner wall is subjected to high thermal and pressure loads during operation of engine due to which it will be in the plastic regime. Main reasons for the failure of such chambers are fatigue in the plastic range (called as low cycle fatigue since the number of cycles to failure will be low in plastic range), creep and thermal ratcheting. Elasto plastic material models are required to simulate the above effects through a cyclic stress analysis. This paper gives the details of cyclic stress analysis carried out for the thrust chamber using different plasticity model combinations available in ANSYS (Version 15) FE code. The best model among the above is applied in the cyclic stress analysis of two dimensional (plane strain and axisymmetric) and three dimensional finite element models of thrust chamber. Cyclic life of the chamber is calculated from stress-strain graph obtained from above analyses.

Taming Liquid Hydrogen: The Centaur Upper Stage Rocket, 1958-2002

NASA Technical Reports Server (NTRS)

Dawson, Virginia P.; Bowles, Mark D.

2004-01-01

During its maiden voyage in May 1962, a Centaur upper stage rocket, mated to an Atlas booster, exploded 54 seconds after launch, engulfing the rocket in a huge fireball. Investigation revealed that Centaur's light, stainless-steel tank had split open, spilling its liquid-hydrogen fuel down its sides, where the flame of the rocket exhaust immediately ignited it. Coming less than a year after President Kennedy had made landing human beings on the Moon a national priority, the loss of Centaur was regarded as a serious setback for the National Aeronautics and Space Administration (NASA). During the failure investigation, Homer Newell, Director of Space Sciences, ruefully declared: "Taming liquid hydrogen to the point where expensive operational space missions can be committed to it has turned out to be more difficult than anyone supposed at the outset." After this failure, Centaur critics, led by Wernher von Braun, mounted a campaign to cancel the program. In addition to the unknowns associated with liquid hydrogen, he objected to the unusual design of Centaur. Like the Atlas rocket, Centaur depended on pressure to keep its paper-thin, stainless-steel shell from collapsing. It was literally inflated with its propellants like a football or balloon and needed no internal structure to give it added strength and stability. The so-called "pressure-stabilized structure" of Centaur, coupled with the light weight of its high- energy cryogenic propellants, made Centaur lighter and more powerful than upper stages that used conventional fuel. But, the critics argued, it would never become the reliable rocket that the United States needed.

Raman Gas Species Measurements in Hydrocarbon-Fueled Rocket Engine Injector Flows

NASA Technical Reports Server (NTRS)

Wehrmeyer, Joseph A.; Trinh, Huu Phuoc; Hartfield, Roy J.; Dobson, Christopher C.; Eskridge, Richard H.

2000-01-01

Propellent injector development at MSFC (Marshall Space Flight Center) includes experimental analysis using optical techniques, such as Raman, fluorescence, or Mie scattering. For the application of spontaneous Raman scattering to hydrocarbon-fueled flows a technique needs to be developed to remove the interfering polycyclic aromatic hydrocarbon fluorescence from the relatively weak Raman signals. A current application of such a technique is to the analysis of the mixing and combustion performance of multijet, impinging-jet candidate fuel injectors for the baseline Mars ascent engine, which will burn methane and liquid oxygen produced in-situ on Mars to reduce the propellent mass transported to Mars for future manned Mars missions. The present technique takes advantage of the strongly polarized nature of Raman scattering. It is shown to be discernable from unpolarized fluorescence interference by subtracting one polarized image from another. Both of these polarized images are obtained from a single laser pulse by using a polarization-separating calcite rhomb mounted in the imaging spectrograph. A demonstration in a propane-air flame is presented.

Ion propulsion engine installed on Deep Space 1 at CCAS

NASA Technical Reports Server (NTRS)

1998-01-01

Workers at the Defense Satellite Communications System Processing Facility (DPF), Cape Canaveral Air Station (CCAS), attach a strap during installation of the ion propulsion engine on Deep Space 1. The first flight in NASA's New Millennium Program, Deep Space 1 is designed to validate 12 new technologies for scientific space missions of the next century, including the engine. Propelled by the gas xenon, the engine is being flight-tested for future deep space and Earth-orbiting missions. Deceptively powerful, the ion drive emits only an eerie blue glow as ionized atoms of xenon are pushed out of the engine. While slow to pick up speed, over the long haul it can deliver 10 times as much thrust per pound of fuel as liquid or solid fuel rockets. Other onboard experiments include software that tracks celestial bodies so the spacecraft can make its own navigation decisions without the intervention of ground controllers. Deep Space 1 will complete most of its mission objectives within the first two months, but will also do a flyby of a near-Earth asteroid, 1992 KD, in July 1999. Deep Space 1 will be launched aboard a Boeing Delta 7326 rocket from Launch Pad 17A, CCAS, in October.

Ion propulsion engine installed on Deep Space 1 at CCAS

NASA Technical Reports Server (NTRS)

1998-01-01

Workers in the Defense Satellite Communications Systems Processing Facility (DPF) at Cape Canaveral Air Station (CCAS) finish installing the ion propulsion engine on Deep Space 1. The first flight in NASA's New Millennium Program, Deep Space 1 is designed to validate 12 new technologies for scientific space missions of the next century, including the engine. Propelled by the gas xenon, the engine is being flight-tested for future deep space and Earth-orbiting missions. Deceptively powerful, the ion drive emits only an eerie blue glow as ionized atoms of xenon are pushed out of the engine. While slow to pick up speed, over the long haul it can deliver 10 times as much thrust per pound of fuel as liquid or solid fuel rockets. Other onboard experiments include software that tracks celestial bodies so the spacecraft can make its own navigation decisions without the intervention of ground controllers. Deep Space 1 will complete most of its mission objectives within the first two months, but will also do a flyby of a near-Earth asteroid, 1992 KD, in July 1999. Deep Space 1 will be launched Oct. 25 aboard a Boeing Delta 7326 rocket from Launch Pad 17A, CCAS.

Ion propulsion engine installed on Deep Space 1 at CCAS

NASA Technical Reports Server (NTRS)

1998-01-01

Workers at the Defense Satellite Communications System Processing Facility (DPF), Cape Canaveral Air Station (CCAS), maneuver the ion propulsion engine into place before installation on Deep Space 1. The first flight in NASA's New Millennium Program, Deep Space 1 is designed to validate 12 new technologies for scientific space missions of the next century, including the engine. Propelled by the gas xenon, the engine is being flight- tested for future deep space and Earth-orbiting missions. Deceptively powerful, the ion drive emits only an eerie blue glow as ionized atoms of xenon are pushed out of the engine. While slow to pick up speed, over the long haul it can deliver 10 times as much thrust per pound of fuel as liquid or solid fuel rockets. Other onboard experiments include software that tracks celestial bodies so the spacecraft can make its own navigation decisions without the intervention of ground controllers. Deep Space 1 will complete most of its mission objectives within the first two months, but will also do a flyby of a near-Earth asteroid, 1992 KD, in July 1999. Deep Space 1 will be launched aboard a Boeing Delta 7326 rocket from Launch Pad 17A, CCAS, in October.

Ion propulsion engine installed on Deep Space 1 at CCAS

NASA Technical Reports Server (NTRS)

1998-01-01

Workers at the Defense Satellite Communications System Processing Facility (DPF), Cape Canaveral Air Station (CCAS), install an ion propulsion engine on Deep Space 1. The first flight in NASA's New Millennium Program, Deep Space 1 is designed to validate 12 new technologies for scientific space missions of the next century, including the engine. Propelled by the gas xenon, the engine is being flight-tested for future deep space and Earth-orbiting missions. Deceptively powerful, the ion drive emits only an eerie blue glow as ionized atoms of xenon are pushed out of the engine. While slow to pick up speed, over the long haul it can deliver 10 times as much thrust per pound of fuel as liquid or solid fuel rockets. Other onboard experiments include software that tracks celestial bodies so the spacecraft can make its own navigation decisions without the intervention of ground controllers. Deep Space 1 will complete most of its mission objectives within the first two months, but will also do a flyby of a near-Earth asteroid, 1992 KD, in July 1999. Deep Space 1 will be launched aboard a Boeing Delta 7326 rocket from Launch Pad 17A, CCAS, in October.

Ion propulsion engine installed on Deep Space 1 at CCAS

NASA Technical Reports Server (NTRS)

1998-01-01

Workers in the Defense Satellite Communications Systems Processing Facility (DPF) at Cape Canaveral Air Station (CCAS) make adjustments while installing the ion propulsion engine on Deep Space 1. The first flight in NASA's New Millennium Program, Deep Space 1 is designed to validate 12 new technologies for scientific space missions of the next century, including the engine. Propelled by the gas xenon, the engine is being flight- tested for future deep space and Earth-orbiting missions. Deceptively powerful, the ion drive emits only an eerie blue glow as ionized atoms of xenon are pushed out of the engine. While slow to pick up speed, over the long haul it can deliver 10 times as much thrust per pound of fuel as liquid or solid fuel rockets. Other onboard experiments include software that tracks celestial bodies so the spacecraft can make its own navigation decisions without the intervention of ground controllers. Deep Space 1 will complete most of its mission objectives within the first two months, but will also do a flyby of a near-Earth asteroid, 1992 KD, in July 1999. Deep Space 1 will be launched Oct. 25 aboard a Boeing Delta 7326 rocket from Launch Pad 17A, CCAS.

Ion propulsion engine installed on Deep Space 1 at CCAS

NASA Technical Reports Server (NTRS)

1998-01-01

Workers at the Defense Satellite Communications System Processing Facility (DPF), Cape Canaveral Air Station (CCAS), make adjustments while installing the ion propulsion engine on Deep Space 1. The first flight in NASA's New Millennium Program, Deep Space 1 is designed to validate 12 new technologies for scientific space missions of the next century, including the engine. Propelled by the gas xenon, the engine is being flight- tested for future deep space and Earth-orbiting missions. Deceptively powerful, the ion drive emits only an eerie blue glow as ionized atoms of xenon are pushed out of the engine. While slow to pick up speed, over the long haul it can deliver 10 times as much thrust per pound of fuel as liquid or solid fuel rockets. Other onboard experiments include software that tracks celestial bodies so the spacecraft can make its own navigation decisions without the intervention of ground controllers. Deep Space 1 will complete most of its mission objectives within the first two months, but will also do a flyby of a near-Earth asteroid, 1992 KD, in July 1999. Deep Space 1 will be launched aboard a Boeing Delta 7326 rocket from Launch Pad 17A, CCAS, in October.

The Benefits of Nuclear Thermal Propulsion (NTP) in an Evolvable Mars Campaign

NASA Technical Reports Server (NTRS)

Borowski, Stanley K.; Mccurdy, David R.

2014-01-01

NTR: High thrust high specific impulse (2 x LOXLH2chemical) engine uses high power density fission reactor with enriched uranium fuel as thermal power source. Reactor heat is removed using H2propellant which is then exhausted to produce thrust. Conventional chemical engine LH2tanks, turbopumps, regenerative nozzles and radiation-cooled shirt extensions used --NTR is next evolutionary step in high performance liquid rocket engines During the Rover program, a common fuel element tie tube design was developed and used in the design of the 50 klbf Kiwi-B4E (1964), 75 klbf Phoebus-1B (1967), 250 klbf Phoebus-2A (June 1968), then back down to the 25 klbf Pewee engine (Nov-Dec 1968) NASA and DOE are using this same approach: design, build, ground then flight test a small engine using a common fuel element that is scalable to a larger 25 klbf thrust engine needed for human missions

Acoustically Forced Coaxial Hydrogen / Liquid Oxygen Jet Flames

DTIC Science & Technology

2016-05-15

serious problems in the development of liquid rocket engines. In order to understand and predict them, it is necessary to understand how representative...liquid rocket injector flames react to acoustic waves. In this study, a representative coaxial gaseous hydrogen / liquid oxygen (LOX) jet flame is...Combustion instabilities can pose serious problems in the development of liquid rocket engines. In order to under- stand and predict them, it is

Propulsion Risk Reduction Activities for Non-Toxic Cryogenic Propulsion

NASA Technical Reports Server (NTRS)

Smith, Timothy D.; Klem, Mark D.; Fisher, Kenneth

2010-01-01

The Propulsion and Cryogenics Advanced Development (PCAD) Project s primary objective is to develop propulsion system technologies for non-toxic or "green" propellants. The PCAD project focuses on the development of non-toxic propulsion technologies needed to provide necessary data and relevant experience to support informed decisions on implementation of non-toxic propellants for space missions. Implementation of non-toxic propellants in high performance propulsion systems offers NASA an opportunity to consider other options than current hypergolic propellants. The PCAD Project is emphasizing technology efforts in reaction control system (RCS) thruster designs, ascent main engines (AME), and descent main engines (DME). PCAD has a series of tasks and contracts to conduct risk reduction and/or retirement activities to demonstrate that non-toxic cryogenic propellants can be a feasible option for space missions. Work has focused on 1) reducing the risk of liquid oxygen/liquid methane ignition, demonstrating the key enabling technologies, and validating performance levels for reaction control engines for use on descent and ascent stages; 2) demonstrating the key enabling technologies and validating performance levels for liquid oxygen/liquid methane ascent engines; and 3) demonstrating the key enabling technologies and validating performance levels for deep throttling liquid oxygen/liquid hydrogen descent engines. The progress of these risk reduction and/or retirement activities will be presented.

Propulsion Risk Reduction Activities for Nontoxic Cryogenic Propulsion

NASA Technical Reports Server (NTRS)

Smith, Timothy D.; Klem, Mark D.; Fisher, Kenneth L.

2010-01-01

The Propulsion and Cryogenics Advanced Development (PCAD) Project s primary objective is to develop propulsion system technologies for nontoxic or "green" propellants. The PCAD project focuses on the development of nontoxic propulsion technologies needed to provide necessary data and relevant experience to support informed decisions on implementation of nontoxic propellants for space missions. Implementation of nontoxic propellants in high performance propulsion systems offers NASA an opportunity to consider other options than current hypergolic propellants. The PCAD Project is emphasizing technology efforts in reaction control system (RCS) thruster designs, ascent main engines (AME), and descent main engines (DME). PCAD has a series of tasks and contracts to conduct risk reduction and/or retirement activities to demonstrate that nontoxic cryogenic propellants can be a feasible option for space missions. Work has focused on 1) reducing the risk of liquid oxygen/liquid methane ignition, demonstrating the key enabling technologies, and validating performance levels for reaction control engines for use on descent and ascent stages; 2) demonstrating the key enabling technologies and validating performance levels for liquid oxygen/liquid methane ascent engines; and 3) demonstrating the key enabling technologies and validating performance levels for deep throttling liquid oxygen/liquid hydrogen descent engines. The progress of these risk reduction and/or retirement activities will be presented.

Enhancement and Extension of Porosity Model in the FDNS-500 Code to Provide Enhanced Simulations of Rocket Engine Components

NASA Technical Reports Server (NTRS)

Cheng, Gary

2003-01-01

In the past, the design of rocket engines has primarily relied on the cold flow/hot fire test, and the empirical correlations developed based on the database from previous designs. However, it is very costly to fabricate and test various hardware designs during the design cycle, whereas the empirical model becomes unreliable in designing the advanced rocket engine where its operating conditions exceed the range of the database. The main goal of the 2nd Generation Reusable Launching Vehicle (GEN-II RLV) is to reduce the cost per payload and to extend the life of the hardware, which poses a great challenge to the rocket engine design. Hence, understanding the flow characteristics in each engine components is thus critical to the engine design. In the last few decades, the methodology of computational fluid dynamics (CFD) has been advanced to be a mature tool of analyzing various engine components. Therefore, it is important for the CFD design tool to be able to properly simulate the hot flow environment near the liquid injector, and thus to accurately predict the heat load to the injector faceplate. However, to date it is still not feasible to conduct CFD simulations of the detailed flowfield with very complicated geometries such as fluid flow and heat transfer in an injector assembly and through a porous plate, which requires gigantic computer memories and power to resolve the detailed geometry. The rigimesh (a sintered metal material), utilized to reduce the heat load to the faceplate, is one of the design concepts for the injector faceplate of the GEN-II RLV. In addition, the injector assembly is designed to distribute propellants into the combustion chamber of the liquid rocket engine. A porosity mode thus becomes a necessity for the CFD code in order to efficiently simulate the flow and heat transfer in these porous media, and maintain good accuracy in describing the flow fields. Currently, the FDNS (Finite Difference Navier-Stakes) code is one of the CFD codes which are most widely used by research engineers at NASA Marshall Space Flight Center (MSFC) to simulate various flow problems related to rocket engines. The objective of this research work during the 10-week summer faculty fellowship program was to 1) debug the framework of the porosity model in the current FDNS code, and 2) validate the porosity model by simulating flows through various porous media such as tube banks and porous plate.

Analytic Modeling of Pressurization and Cryogenic Propellant

NASA Technical Reports Server (NTRS)

Corpening, Jeremy H.

2010-01-01

An analytic model for pressurization and cryogenic propellant conditions during all mission phases of any liquid rocket based vehicle has been developed and validated. The model assumes the propellant tanks to be divided into five nodes and also implements an empirical correlation for liquid stratification if desired. The five nodes include a tank wall node exposed to ullage gas, an ullage gas node, a saturated propellant vapor node at the liquid-vapor interface, a liquid node, and a tank wall node exposed to liquid. The conservation equations of mass and energy are then applied across all the node boundaries and, with the use of perfect gas assumptions, explicit solutions for ullage and liquid conditions are derived. All fluid properties are updated real time using NIST Refprop.1 Further, mass transfer at the liquid-vapor interface is included in the form of evaporation, bulk boiling of liquid propellant, and condensation given the appropriate conditions for each. Model validation has proven highly successful against previous analytic models and various Saturn era test data and reasonably successful against more recent LH2 tank self pressurization ground test data. Finally, this model has been applied to numerous design iterations for the Altair Lunar Lander, Ares V Core Stage, and Ares V Earth Departure Stage in order to characterize Helium and autogenous pressurant requirements, propellant lost to evaporation and thermodynamic venting to maintain propellant conditions, and non-uniform tank draining in configurations utilizing multiple LH2 or LO2 propellant tanks. In conclusion, this model provides an accurate and efficient means of analyzing multiple design configurations for any cryogenic propellant tank in launch, low-acceleration coast, or in-space maneuvering and supplies the user with pressurization requirements, unusable propellants from evaporation and liquid stratification, and general ullage gas, liquid, and tank wall conditions as functions of time.

Fiber reinforced superalloys for rocket engines

NASA Technical Reports Server (NTRS)

Petrasek, Donald W.; Stephens, Joseph R.

1989-01-01

High pressure turbopumps for advanced reusable liquid propellant rocket engines such as that for the Space Shuttle Main Engine (SSME) require turbine blade materials that operate under extreme conditions of temperature, hydrogen environment, high-cycle fatigue loading, thermal fatigue and thermal shock. Such requirements tax the capabilities of current blade materials. Based on projections of properties for tungsten fiber reinforced superalloy (FRS) composites, it was concluded that FRS turbine blades offer the potential of a several fold increase in life and over a 200 C increase in temperature capability over the current SSME blade material. FRS composites were evaluated with respect to mechanical property requirements for SSME blade applications. Compared to the current blade material, the thermal shock resistance of FRS materials is excellent, two to nine times better, and their thermal fatigue resistance is equal to or higher than the current blade material. FRS materials had excellent low and high-cycle fatigue strengths, and thermal shock-induced surface microcracks had no influence on their fatigue strength. The material also exhibited negligible embrittlement when exposed to a hydrogen environment.

KSC-98pc1334

NASA Image and Video Library

1998-10-12

KENNEDY SPACE CENTER, FLA. -- On Launch Pad 17A at Cape Canaveral Air Station, Deep Space 1 is viewed from above after installation on a Boeing Delta 7326 rocket . Targeted for launch on Oct. 25, Deep Space 1 is the first flight in NASA's New Millennium Program, and is designed to validate 12 new technologies for scientific space missions of the next century, including the engine. Propelled by the gas xenon, the engine is being flight-tested for future deep space and Earth-orbiting missions. Deceptively powerful, the ion drive emits only an eerie blue glow as ionized atoms of xenon are pushed out of the engine. While slow to pick up speed, over the long haul it can deliver 10 times as much thrust per pound of fuel as liquid or solid fuel rockets. Other onboard experiments include software that tracks celestial bodies so the spacecraft can make its own navigation decisions without the intervention of ground controllers. Deep Space 1 will complete most of its mission objectives within the first two months, but will also do a flyby of a near-Earth asteroid, 1992 KD, in July 1999

KSC-98pc1354

NASA Image and Video Library

1998-10-16

KENNEDY SPACE CENTER, FLA. -- On Launch Pad 17A at Cape Canaveral Air Station, workers maneuver the second half of the fairing to encapsulate Deep Space 1, targeted for launch aboard a Boeing Delta II rocket on Oct. 24. The first flight in NASA's New Millennium Program, Deep Space 1 is designed to validate 12 new technologies for scientific space missions of the next century, including the engine. Propelled by the gas xenon, the engine is being flight-tested for future deep space and Earth-orbiting missions. Deceptively powerful, the ion drive emits only an eerie blue glow as ionized atoms of xenon are pushed out of the engine. While slow to pick up speed, over the long haul it can deliver 10 times as much thrust per pound of fuel as liquid or solid fuel rockets. Other onboard experiments include software that tracks celestial bodies so the spacecraft can make its own navigation decisions without the intervention of ground controllers. Deep Space 1 will complete most of its mission objectives within the first two months, but will also do a flyby of a near-Earth asteroid, 1992 KD, in July 1999

Fiber reinforced superalloys for rocket engines

NASA Technical Reports Server (NTRS)

Petrasek, Donald W.; Stephens, Joseph R.

1988-01-01

High-pressure turbopumps for advanced reusable liquid-propellant rocket engines such as that for the Space Shuttle Main Engine (SSME) require turbine blade materials that operate under extreme conditions of temperature, hydrogen environment, high-cycle fatigue loading, thermal fatigue and thermal shock. Such requirements tax the capabilities of current blade materials. Based on projections of properties for tungsten fiber reinforced superalloy (FRS) composites, it was concluded that FRS turbine blades offer the potential of a several-fold increase in life and over a 200C increase in temperature capability over current SSME blade material. FRS composites were evaluated with respect to mechanical property requirements for SSME blade applications. Compared to the current blade material, the thermal shock resistance of FRS materials is excellent, two to nine times better, and their thermal fatigue resistance is equal to or higher than the current blade material. FRS materials had excellent low and high-cycle fatigue strengths, and thermal shock-induced surface microcracks had no influence on their fatigue strength. The material also exhibited negligible embrittlement when exposed to a hydrogen environment.

KSC-98pc1335

NASA Image and Video Library

1998-10-12

KENNEDY SPACE CENTER, FLA. -- On Launch Pad 17A at Cape Canaveral Air Station, Deep Space 1 is uncovered after installation on a Boeing Delta 7326 rocket. Targeted for launch on Oct. 25, Deep Space 1 is the first flight in NASA's New Millennium Program, and is designed to validate 12 new technologies for scientific space missions of the next century, including the engine. Propelled by the gas xenon, the engine is being flight-tested for future deep space and Earth-orbiting missions. Deceptively powerful, the ion drive emits only an eerie blue glow as ionized atoms of xenon are pushed out of the engine. While slow to pick up speed, over the long haul it can deliver 10 times as much thrust per pound of fuel as liquid or solid fuel rockets. Other onboard experiments include software that tracks celestial bodies so the spacecraft can make its own navigation decisions without the intervention of ground controllers. Deep Space 1 will complete most of its mission objectives within the first two months, but will also do a flyby of a near-Earth asteroid, 1992 KD, in July 1999

KSC-98pc1355

NASA Image and Video Library

1998-10-16

KENNEDY SPACE CENTER, FLA. -- On Launch Pad 17A at Cape Canaveral Air Station, workers check make a final check of the fairing encapsulating Deep Space 1, which is targeted for launch aboard a Boeing Delta II rocket on Oct. 24. The first flight in NASA's New Millennium Program, Deep Space 1 is designed to validate 12 new technologies for scientific space missions of the next century, including the engine. Propelled by the gas xenon, the engine is being flight-tested for future deep space and Earth-orbiting missions. Deceptively powerful, the ion drive emits only an eerie blue glow as ionized atoms of xenon are pushed out of the engine. While slow to pick up speed, over the long haul it can deliver 10 times as much thrust per pound of fuel as liquid or solid fuel rockets. Other onboard experiments include software that tracks celestial bodies so the spacecraft can make its own navigation decisions without the intervention of ground controllers. Deep Space 1 will complete most of its mission objectives within the first two months, but will also do a flyby of a near-Earth asteroid, 1992 KD, in July 1999

KSC-98pc1331

NASA Image and Video Library

1998-10-12

KENNEDY SPACE CENTER, FLA. -- On Launch Pad 17A at Cape Canaveral Air Station, Deep Space 1 is lowered in the white room for installation on a Boeing Delta 7326 rocket . The spacecraft is targeted for launch on Oct. 25. Deep Space 1 is the first flight in NASA's New Millennium Program, and is designed to validate 12 new technologies for scientific space missions of the next century, including the engine. Propelled by the gas xenon, the engine is being flight-tested for future deep space and Earth-orbiting missions. Deceptively powerful, the ion drive emits only an eerie blue glow as ionized atoms of xenon are pushed out of the engine. While slow to pick up speed, over the long haul it can deliver 10 times as much thrust per pound of fuel as liquid or solid fuel rockets. Other onboard experiments include software that tracks celestial bodies so the spacecraft can make its own navigation decisions without the intervention of ground controllers. Deep Space 1 will complete most of its mission objectives within the first two months, but will also do a flyby of a near-Earth asteroid, 1992 KD, in July 1999

KSC-98pc1333

NASA Image and Video Library

1998-10-12

KENNEDY SPACE CENTER, FLA. -- On Launch Pad 17A at Cape Canaveral Air Station, workers remove the transportation canister around Deep Space 1 after installation on a Boeing Delta 7326 rocket . Targeted for launch on Oct. 25, Deep Space 1 is the first flight in NASA's New Millennium Program, and is designed to validate 12 new technologies for scientific space missions of the next century, including the engine. Propelled by the gas xenon, the engine is being flight-tested for future deep space and Earth-orbiting missions. Deceptively powerful, the ion drive emits only an eerie blue glow as ionized atoms of xenon are pushed out of the engine. While slow to pick up speed, over the long haul it can deliver 10 times as much thrust per pound of fuel as liquid or solid fuel rockets. Other onboard experiments include software that tracks celestial bodies so the spacecraft can make its own navigation decisions without the intervention of ground controllers. Deep Space 1 will complete most of its mission objectives within the first two months, but will also do a flyby of a near-Earth asteroid, 1992 KD, in July 1999

KSC-98pc1346

NASA Image and Video Library

1998-10-16

KENNEDY SPACE CENTER, FLA. -- On Launch Pad 17A at Cape Canaveral Air Station, workers begin encapsulating Deep Space 1 with the fairing (right side). Targeted for launch aboard a Boeing Delta 7326 rocket on Oct. 25, Deep Space 1 is the first flight in NASA's New Millennium Program, and is designed to validate 12 new technologies for scientific space missions of the next century, including the engine. Propelled by the gas xenon, the engine is being flight-tested for future deep space and Earth-orbiting missions. Deceptively powerful, the ion drive emits only an eerie blue glow as ionized atoms of xenon are pushed out of the engine. While slow to pick up speed, over the long haul it can deliver 10 times as much thrust per pound of fuel as liquid or solid fuel rockets. Other onboard experiments include software that tracks celestial bodies so the spacecraft can make its own navigation decisions without the intervention of ground controllers. Deep Space 1 will complete most of its mission objectives within the first two months, but will also do a flyby of a near-Earth asteroid, 1992 KD, in July 1999

Main Chamber and Preburner Injector Technology

NASA Technical Reports Server (NTRS)

Santoro, Robert J.; Merkle, Charles L.

1999-01-01

This document reports the experimental and analytical research carried out at the Penn State Propulsion Engineering Research Center in support of NASA's plan to develop advanced technologies for future single stage to orbit (SSTO) propulsion systems. The focus of the work is on understanding specific technical issues related to bi-propellant and tri-propellant thrusters. The experiments concentrate on both cold flow demonstrations and hot-fire uni-element tests to demonstrate concepts that can be incorporated into hardware design and development. The analysis is CFD-based and is intended to support the design and interpretation of the experiments and to extrapolate findings to full-scale designs. The research is divided into five main categories that impact various SSTO development scenarios. The first category focuses on RP-1/gaseous hydrogen (GH2)/gaseous oxygen (GO2) tri-propellant combustion with specific emphasis on understanding the benefits of hydrogen addition to RP-1/oxygen combustion and in developing innovative injector technology. The second category investigates liquid oxygen (LOX)/GH2 combustion at main chamber near stoichiometric conditions to improve understanding of existing LOX/GH2 rocket systems. The third and fourth categories investigate the technical issues related with oxidizer-rich and fuel-rich propulsive concepts, issues that are necessary for developing the full-flow engine cycle. Here, injector technology issues for both LOX/GH2 and LOX/RP-1 propellants are examined. The last category, also related to the full-flow engine cycle, examines injector technology needs for GO2/GH2 propellant combustion at near-stoichiometric conditions for main chamber application.

Theoretical and Experimental Investigations of Ignition, Combustion and Expansion Processes of Hypergolic Liquid Fuel Combinations at Gas Temperatures up to 3000 K. Thesis - Rhein-Westfalia Technical Coll., 1967

NASA Technical Reports Server (NTRS)

Schulz, Harry

1987-01-01

The ignition, combustion, and expansion characteristics of hypergolic liquid propellant mixtures in small rocket engines are studied theoretically and experimentally. It is shown by using the Bray approximation procedure that the reaction H + OH + M = H2O + M (where M is the molecular mass of the gas mixture) has a strong effect on the combustion efficiency. Increases in recombination energies ranging from 30 to 65% were obtained when the rate of this reaction was increased by a factor of 10 in gas mixtures containing 90% oxygen. The effect of aluminum additions and various injection techniques on the combustion process is investigated.

62. Historic propellant piping diagram of oxidant pit at Building ...

Library of Congress Historic Buildings Survey, Historic Engineering Record, Historic Landscapes Survey

62. Historic propellant piping diagram of oxidant pit at Building 202, January 6, 1956. NASA GRC drawing no. CF-101644. (On file at NASA Glenn Research Center). - Rocket Engine Testing Facility, GRC Building No. 202, NASA Glenn Research Center, Cleveland, Cuyahoga County, OH

Development of Life Prediction Capabilities for Liquid Propellant Rocket Engines. Task 3. Sensor Data Validation and Reconstruction. Phase 1. System Architecture Study

DTIC Science & Technology

1991-06-01

Validation And Reconstruction -~ Phase 1: System Architecture Study i ".- Contract NAS 3 -25883 I - _ CR-187124 -4 Phase I Final Report,, " , I Prepared for...131 NAS 3 -25883 1.0 INTRODUCTION 1 2.0 EXECUTIVE SUMMARY 2 3.0 TECHNICAL DISCUSSION 8 3.1 Review of SSME Test Data and Validation Procedure 8 3.1.1...NAS 3 -25883 FIGURES FigureNo. e 1 Elements The Sensor Data Validation and Signal Reconstuction System 7 3 Current NASA MSFC Data Review Process 12 4

Tests of Five Full-Scale Propellers in the Presence of a Radial and a Liquid-Cooled Engine Nacelle, Including Tests of Two Spinners

NASA Technical Reports Server (NTRS)

Biermann, David; Hartman, Edwin P

1938-01-01

Wind-tunnel tests are reported of five 3-blade 10-foot propellers operating in front of a radial and a liquid-cooled engine nacelle. The range of blade angles investigated extended from 15 degrees to 45 degrees. Two spinners were tested in conjunction with the liquid-cooled engine nacelle. Comparisons are made between propellers having different blade-shank shapes, blades of different thickness, and different airfoil sections. The results show that propellers operating in front of the liquid-cooled engine nacelle had higher take-off efficiencies than when operating in front of the radial engine nacelle; the peak efficiency was higher only when spinners were employed. One spinner increased the propulsive efficiency of the liquid-cooled unit 6 percent for the highest blade-angle setting investigated and less for lower blade angles. The propeller having airfoil sections extending into the hub was superior to one having round blade shanks. The thick propeller having a Clark y section had a higher take-off efficiency than the thinner one, but its maximum efficiency was possibly lower. Of the three blade sections tested, Clark y, R.A.F. 6, and NACA 2400-34, the Clark y was superior for the high-speed condition, but the R.A.F. 6 excelled for the take-off condition.

Air breathing engine/rocket trajectory optimization

NASA Technical Reports Server (NTRS)

Smith, V. K., III

1979-01-01

This research has focused on improving the mathematical models of the air-breathing propulsion systems, which can be mated with the rocket engine model and incorporated in trajectory optimization codes. Improved engine simulations provided accurate representation of the complex cycles proposed for advanced launch vehicles, thereby increasing the confidence in propellant use and payload calculations. The versatile QNEP (Quick Navy Engine Program) was modified to allow treatment of advanced turboaccelerator cycles using hydrogen or hydrocarbon fuels and operating in the vehicle flow field.

Interactive Schematic Integration Within the Propellant System Modeling Environment

NASA Technical Reports Server (NTRS)

Coote, David; Ryan, Harry; Burton, Kenneth; McKinney, Lee; Woodman, Don

2012-01-01

Task requirements for rocket propulsion test preparations of the test stand facilities drive the need to model the test facility propellant systems prior to constructing physical modifications. The Propellant System Modeling Environment (PSME) is an initiative designed to enable increased efficiency and expanded capabilities to a broader base of NASA engineers in the use of modeling and simulation (M&S) technologies for rocket propulsion test and launch mission requirements. PSME will enable a wider scope of users to utilize M&S of propulsion test and launch facilities for predictive and post-analysis functionality by offering a clean, easy-to-use, high-performance application environment.

The engineering of a nuclear thermal landing and ascent vehicle utilizing indigenous Martian propellant

NASA Technical Reports Server (NTRS)

Zubrin, Robert M.

1991-01-01

The following paper reports on a design study of a novel space transportation concept known as a 'NIMF' (Nuclear rocket using Indigenous Martian Fuel). The NIMF is a ballistic vehicle which obtains its propellant out of the Martian air by compression and liquefaction of atmospheric CO2. This propellant is subsequently used to generate rocket thrust at a specific impulse of 264 s by being heated to high temperature (2800 K) gas in the NIMFs' nuclear thermal rocket engines. The vehicle is designed to provide surface to orbit and surface to surface transportation, as well as housing, for a crew of three astronauts. It is capable of refueling itself for a flight to its maximum orbit in less than 50 days. The ballistic NIMF has a mass of 44.7 tonnes and, with the assumed 2800 K propellant temperature, is capable of attaining highly energetic (250 km by 34,000 km elliptical) orbits. This allows it to rendezvous with interplanetary transfer vehicles which are only very loosely bound into orbit around Mars. If a propellant temperature of 2000 K is assumed, then low Mars orbit can be attained; while if 3100 K is assumed, then the ballistic NIMF is capable of injecting itself onto a minimum energy transfer orbit to Earth in a direct ascent from the Martian surface.

Powdered aluminum and oxygen rocket propellants: Subscale combustion experiments

NASA Technical Reports Server (NTRS)

Meyer, Mike L.

1993-01-01

Aluminum combined with oxygen has been proposed as a potential lunar in situ propellant for ascent/descent and return missions for future lunar exploration. Engine concepts proposed to use this propellant have not previously been demonstrated, and the impact on performance from combustion and two-phase flow losses could only be estimated. Therefore, combustion tests were performed for aluminum and aluminum/magnesium alloy powders with oxygen in subscale heat-sink rocket engine hardware. The metal powder was pneumatically injected, with a small amount of nitrogen, through the center orifice of a single element O-F-O triplet injector. Gaseous oxygen impinged on the fuel stream. Hot-fire tests of aluminum/oxygen were performed over a mixture ratio range of 0.5 to 3.0, and at a chamber pressure of approximately 480 kPa (70 psia). The theoretical performance of the propellants was analyzed over a mixture ratio range of 0.5 to 5.0. In the theoretical predictions the ideal one-dimensional equilibrium rocket performance was reduced by loss mechanisms including finite rate kinetics, two-dimensional divergence losses, and boundary layer losses. Lower than predicted characteristic velocity and specific impulse performance efficiencies were achieved in the hot-fire tests, and this was attributed to poor mixing of the propellants and two-phase flow effects. Several tests with aluminum/9.8 percent magnesium alloy powder did not indicate any advantage over the pure aluminum fuel.

Comparison of Laminar and Linear Eddy Model Closures for Combustion Instability Simulations

DTIC Science & Technology

2015-07-01

14. ABSTRACT Unstable liquid rocket engines can produce highly complex dynamic flowfields with features such as rapid changes in temperature and...applicability. In the present study, the linear eddy model (LEM) is applied to an unstable single element liquid rocket engine to assess its performance and to...Sankaran‡ Air Force Research Laboratory, Edwards AFB, CA, 93524 Unstable liquid rocket engines can produce highly complex dynamic flowfields with features

Coaxial atomization of a round liquid jet in a high speed gas stream: A phenomenological study

NASA Astrophysics Data System (ADS)

Mayer, W. O. H.

1994-05-01

Coaxial injectors have proven to be advantageous for the injection, atomization and mixing of propellants in cryogenic H2/O2 rocket engines. Thereby, a round liquid oxygen jet is atomized by a fast, coaxial gaseous hydrogen jet. This article summarizes phenomenological studies of coaxial spray generation under a broad variation of influencing parameters including injector design, inflow, and fluid conditions. The experimental investigations, performed using spark light photography and high speed cinematography in a shadow graph setup as main diagnostic means, illuminate the most important processes leading to atomization. These are identified as turbulence in the liquid jet, surface instability, surface wave growth and droplet detachment. Numerical simulations including free surface flow phenomena are a further diagnostic tool to elucidate some atomization particulars. The results of the study are of general importance in the field of liquid atomization.

Hot-Fire Testing of 100 LB(sub F) LOX/LCH4 Reaction Control Engine at Altitude Conditions

NASA Technical Reports Server (NTRS)

Marshall, William M.; Kleinhenz, Julie E.

2010-01-01

Liquid oxygen/liquid methane (LO2/LCH4 ) has recently been viewed as a potential green propulsion system for both the Altair ascent main engine (AME) and reaction control system (RCS). The Propulsion and Cryogenic Advanced Development Project (PCAD) has been tasked by NASA to develop these green propellant systems to enable safe and cost effective exploration missions. However, experience with LO2/LCH4 as a propellant combination is limited, so testing of these systems is critical to demonstrating reliable ignition and performance. A test program of a 100 lb f reaction control engine (RCE) is underway at the Altitude Combustion Stand (ACS) of the NASA Glenn Research Center, with a focus on conducting tests at altitude conditions. These tests include a unique propellant conditioning feed system (PCFS) which allows for the inlet conditions of the propellant to be varied to test warm to subcooled liquid propellant temperatures. Engine performance, including thrust, c* and vacuum specific impulse (I(sub sp,vac)) will be presented as a function of propellant temperature conditions. In general, the engine performed as expected, with higher performance at warmer propellant temperatures but better efficiency at lower propellant temperatures. Mixture ratio effects were inconclusive within the uncertainty bands of data, but qualitatively showed higher performance at lower ratios.

Determination of the availability of appropriate aged flight rocket motors. [captive tests to determine case bond separation and grain bore cracking

NASA Technical Reports Server (NTRS)

Martin, P. J.

1974-01-01

A program to identify surplus solid rocket propellant engines which would be available for a program of functional integrity testing was conducted. The engines are classified as: (1) upper stage and apogee engines, (2) sounding rocket and launch vehicle engines, and (3) jato, sled, and tactical engines. Nearly all the engines were available because their age exceeds the warranted shelf life. The preference for testing included tests at nominal flight conditions, at design limits, and to establish margin limits. The principal failure modes of interest were case bond separation and grain bore cracking. Data concerning the identification and characteristics of each engine are tabulated. Methods for conducting the tests are described.

Combustion Stability of the Gas Generator Assembly from J-2X Engine E10001 and Powerpack Tests

NASA Technical Reports Server (NTRS)

Hulka, J. R.; Kenny, R. L.; Casiano, M. J.

2013-01-01

Testing of a powerpack configuration (turbomachinery and gas generator assembly) and the first complete engine system of the liquid oxygen/liquid hydrogen propellant J-2X rocket engine have been completed at the NASA Stennis Space Center. The combustion stability characteristics of the gas generator assemblies on these two systems are of interest for reporting since considerable effort was expended to eliminate combustion instability during early development of the gas generator assembly with workhorse hardware. Comparing the final workhorse gas generator assembly development test data to the powerpack and engine system test data provides an opportunity to investigate how the nearly identical configurations of gas generator assemblies operate with two very different propellant supply systems one the autonomous pressure-fed test configuration on the workhorse development test stand, the other the pump-fed configurations on the powerpack and engine systems. The development of the gas generator assembly and the elimination of the combustion instability on the pressure-fed workhorse test stand have been reported extensively in the two previous Liquid Propulsion Subcommittee meetings 1-7. The powerpack and engine system testing have been conducted from mid-2011 through 2012. All tests of the powerpack and engine system gas generator systems to date have been stable. However, measureable dynamic behavior, similar to that observed on the pressure-fed test stand and reported in Ref. [6] and attributed to an injection-coupled response, has appeared in both powerpack and engine system tests. As discussed in Ref. [6], these injection-coupled responses are influenced by the interaction of the combustion chamber with a branch pipe in the hot gas duct that supplies gaseous helium to pre-spin the turbine during the start transient. This paper presents the powerpack and engine system gas generator test data, compares these data to the development test data, and provides additional combustion stability analyses of the configurations.

The Peak of Rocket Production: The Designer of Ballistic Missiles V.F. Utkin (1923-2000)

NASA Astrophysics Data System (ADS)

Prisniakov, V.; Sitnikova, N.

2002-01-01

The main landmarks of the biography of the general designer of the most power missiles Vladimira Fe- dorovicha Utkina are stated. Formation of character of outstanding scientist of 20 century as the son of the Soviet epoch is shown. He belongs to that generation which had many difficulties and afflictions - hungry time, the heavy years of the second world war, post-war disruption, but also many happy days - the Victory above fascism, restoration of the country, pride of successes in conquest of Space. In June, 1941 Vladimir has finished school with honours certificate and since October, 1941 up to the end of the second world war was on various fronts. After the ending of Leningrad military-mechanical insti- tute the young engineer came in Southern engineering works in Dnipropetrovsk (Yugmach). Here for 40 years there was a dizzy ascent of the beginner-designer over a ladder of space-rocket's Olympus up to the chief designer and the general director of the biggest in the world of rocket concern (9000 high quality engineers of Design office Yugnoe (DOYu) and 60 thousand workers Yugmach). After death in 1971 to year of main designer M. Jangele V. F. Utkin has headed Design office Yugnoe. Under ma- nual of V. Utkin four strategic rocket complexes of new generation SS-17, SS-18 (three updatings with divided head parts with weight of 8 tons), SS-24 (railway and shaft basing) were developed and han- ded over on arms. Among development of academician V. Utkin there is a rocket-carrier "Zenit" which delivers to an orbit over 12 tons of a payload. This rocket is also a basis of the first stage of reusable transport space system "Energia-Buran". Under manual of V. Utkin were developed and used the con- version carrier-rockets "Ziclon" and "Kosmos", as well the effective satellites of a defensive, scientific and economic direction, among which family of satellites "Kosmos", the satellite "Ocean", equipped by a locator of the side observation too. Largest scientific achievements V. Utkin and his pupils are crea- tion unique "mortar" launching of a heavy liquid rocket from shaft, the decision of a complex of prob- lems on maintenance ready for military action (continuous attendance) of liquid rockets in the filled condi-tion for many years, maintenance of stability of rockets at action on them of striking factors of nuclear explosion. With personal participation of academician IAA V. Utkin the following large scien- tific and technical results were received: (a) a military railway rocket complex with intercontinental solid-propellant rocket with starting weight of 105 tons and with 10 warheads; (b) a method of war manage-ment with the help of command rockets; (c) a method of definition of characteristics of means of overcoming of antimissile defense; (d) war intercontinental rockets with the increased accuracy, with the survivability, with the availability for action; (e) a commanding rocket. Design' decisions not ha- ving the analogues in world: (a) managements of flight solid-propellant an intercontinental ballistic missiles by means of a deviating head part; (b) managements solid-propellant rocket by method of inje- ction of gas in supercritical part of nozzle; (c) industrial introduction of the newest materials etc.V. Ut- kin is the active participant of works in the field of the international cooperation in research and deve- lopment of a space. In 1990 V. Utkin hold a high post of the director of ZSNIIMACH which is leading organization of a space-rocket industry of Russia. Under manual V. Utkin the Federal space program of Russia was developed. V. Utkin had huge authority as the chairman of Advice of the Main designers of the USSR. He was the co-chairman combined commission of experts V. Utkin - T. Stafford" on problems of maintenance joint manned flights. He was the chairman of Coordination advice under the program of researches on manned space complexes. V. Utkin dreamed to be the active participant of a new stage of the outer space exploration, connected with commissioning ISS. He believed, that Human mind and integration of efforts of all world community will be prevailed over the world.

Tail Service Mast Umbilical Arrival

NASA Image and Video Library

2016-08-02

Technicians assist as a crane is used to lift the first Tail Service Mast Umbilical (TSMU) up from the flatbed of the transport truck at the Launch Equipment Test Facility (LETF) at NASA’s Kennedy Space Center in Florida. Two TSMUs will provide liquid propellants and power to the Space Launch System (SLS) rocket’s core stage engine. Both TSMUs will connect to the zero-level deck on the mobile launcher, providing fuel and electricity to the SLS rocket before it launches on Exploration Mission 1. The TSMU will undergo testing and validation at the LETF to verify it is functioning properly. The center’s Engineering Directorate and the Ground Systems Development and Operations Program are overseeing processing and testing of the umbilicals.

Tail Service Mast Umbilical Arrival

NASA Image and Video Library

2016-08-02

Technicians assist as a crane is used to lift the first Tail Service Mast Umbilical (TSMU) away from the flatbed of the transport truck at the Launch Equipment Test Facility (LETF) at NASA’s Kennedy Space Center in Florida. Two TSMUs will provide liquid propellants and power to the Space Launch System (SLS) rocket’s core stage engine. Both TSMUs will connect to the zero-level deck on the mobile launcher, providing fuel and electricity to the SLS rocket before it launches on Exploration Mission 1. The TSMU will undergo testing and validation at the LETF to verify it is functioning properly. The center’s Engineering Directorate and the Ground Systems Development and Operations Program are overseeing processing and testing of the umbilicals.

Aluminum/hydrocarbon gel propellants: An experimental and theoretical investigation of secondary atomization and predicted rocket engine performance

NASA Astrophysics Data System (ADS)

Mueller, Donn Christopher

1997-12-01

Experimental and theoretical investigations of aluminum/hydrocarbon gel propellant secondary atomization and its potential effects on rocket engine performance were conducted. In the experimental efforts, a dilute, polydisperse, gel droplet spray was injected into the postflame region of a burner and droplet size distributions was measured as a function of position above the burner using a laser-based sizing/velocimetry technique. The sizing/velocimetry technique was developed to measure droplets in the 10-125 mum size range and avoids size-biased detection through the use of a uniformly illuminated probe volume. The technique was used to determine particle size distributions and velocities at various axial locations above the burner for JP-10, and 50 and 60 wt% aluminum gels. Droplet shell formation models were applied to aluminum/hydrocarbon gels to examine particle size and mass loading effects on the minimum droplet diameter that will permit secondary atomization. This diameter was predicted to be 38.1 and 34.7 mum for the 50 and 60 wt% gels, which is somewhat greater than the experimentally measured 30 and 25 mum diameters. In the theoretical efforts, three models were developed and an existing rocket code was exercised to gain insights into secondary atomization. The first model was designed to predict gel droplet properties and shell stresses after rigid shell formation, while the second, a one-dimensional gel spray combustion model was created to quantify the secondary atomization process. Experimental and numerical comparisons verify that secondary atomization occurs in 10-125 mum diameter particles although an exact model could not be derived. The third model, a one-dimensional gel-fueled rocket combustion chamber, was developed to evaluate secondary atomization effects on various engine performance parameters. Results show that only modest secondary atomization may be required to reduce propellant burnout distance and radiation losses. A solid propellant engine code was employed to estimate nozzle two-phase flow losses and engine performance for upper-stage and booster missions (3-6% and 2-3%, respectively). Given these losses and other difficulties, metallized gel propellants may be impractical in high-expansion ratio engines. Although uncertainties remain, it appears that performance gains will be minimal in gross-weight limited missions, but that significant gains may arise in volume-limited missions.

Design and evaluation of high performance rocket engine injectors for use with hydrocarbon fuels

NASA Technical Reports Server (NTRS)

Pavli, A. J.

1979-01-01

An experimental program to determine the feasibility of using a heavy hydrocarbon fuel as a rocket propellant is reported herein. A method of predicting performance of a heavy hydrocarbon in terms of vaporization effectiveness is described and compared to other fuels and to experimental test results. The work was done at a chamber pressure of 4137 KN/sq M (600 psia) with RP-1, JP-10, and liquefied natural gas as fuels, and liquid oxygen as the oxidizer. Combustion length effects were explored over a range of 21.6 cm (8 1/2 in.) to 55.9 cm (22 in.). Four injector types were tested, each over a range of mixture ratios. Further configuration modifications were obtained by 'reaming' each injector several times to provide test data over a range of injector pressure drop.

A concept of highly maneuverable experimental space (HIMES) vehicle

NASA Astrophysics Data System (ADS)

Nagatomo, M.; Naruo, Y.; Inatani, Y.

1985-10-01

The development of a highly maneuverable experimental space (HIMES) vehicle is proposed. This reusable sounding rocket is to be propelled by a liquid hydrogen/LOX engine, and have a maximum payload mass of 500 kg at an altitude of 300 km. The main subsystems of HIMES, the fuselage and wing structure, propulsion, and navigation, guidance, and control system, are described and a diagram is provided. The operational features of HIMES are defined by three mission models. In the first model the vehicle is used as a stable platform of low velocity relative to the environment; model two represents the suborbital flight of sounding rockets, and model three is used for orbital reentry experiments and the testing of a new system of winged space vehicles. Typical mission profiles for the three models are presented. A cost estimation of the HIMES vehicle is given.

The Guggenheim Aeronautics Laboratory at Caltech and the creation of the modern rocket motor (1936-1946): How the dynamics of rocket theory became reality

NASA Astrophysics Data System (ADS)

Zibit, Benjamin Seth

This thesis explores and unfolds the story of discovery in rocketry at The California Institute of Technology---specifically at Caltech's Guggenheim Aeronautics Laboratory---in the 1930s and 1940s. Caltech was home to a small group of engineering students and experimenters who, beginning in the winter of 1935--1936, formed a study and research team destined to change the face of rocket science in the United States. The group, known as the Guggenheim Aeronautics Laboratory (GALCIT, for short) Rocket Research Group, invented a new type of solid-rocket propellant, made distinct and influential discoveries in the theory of rocket combustion and design, founded the Jet Propulsion Laboratory, and incorporated the first American industrial concern devoted entirely to rocket motor production: The Aerojet Corporation. The theoretical work of team members, Frank Malina, Hsueh-shen Tsien, Homer J. Stewart, and Mark Mills, is examined in this thesis in detail. The author scrutinizes Frank Malina's doctoral thesis (both its assumptions and its mathematics), and finds that, although Malina's key assertions, his formulae, hold, his work is shown to make key assumptions about rocket dynamics which only stand the test of validity if certain approximations, rather than exact measurements, are accepted. Malina studied the important connection between motor-nozzle design and thrust; in his Ph.D. thesis, he developed mathematical statements which more precisely defined the design/thrust relation. One of Malina's colleagues on the Rocket Research Team, John Whiteside Parsons, created a new type of solid propellant in the winter of 1941--1942. This propellant, known as a composite propellant (because it simply was a relatively inert amalgam of propellant and oxidizer in non-powder form), became the forerunner of all modern solid propellants, and has become one of the seminal discoveries in the field of Twentieth Century rocketry. The latter chapters of this dissertation discuss the creation of the jet Propulsion Laboratory, the founding of the Aerojet Corporation, and emphasizes the issue of JPL's close relation to military development of the rocket becomes a core subject of this thesis. Cooperation between engineers in an academic setting and the military was not merely inevitable in the 1940s---it was actively fostered and proved quite profitable to all concerned. The deep relationship between the Guggenheim Aeronautics Laboratory and the Army Air Force was one model of the evolution of a permanent institutional edifice, weaving academic research and military end-use together. The dissertation concludes that what began as a modest effort to understand rocket theory in greater depth led within ten years to both research and development tracks which have profoundly altered the technological and military definition of modern history.

KSC-98pc1261

NASA Image and Video Library

1998-10-07

KENNEDY SPACE CENTER, FLA. -- Workers at the Defense Satellite Communications System Processing Facility (DPF), Cape Canaveral Air Station (CCAS), attach a strap during installation of the ion propulsion engine on Deep Space 1. The first flight in NASA's New Millennium Program, Deep Space 1 is designed to validate 12 new technologies for scientific space missions of the next century, including the engine. Propelled by the gas xenon, the engine is being flight-tested for future deep space and Earth-orbiting missions. Deceptively powerful, the ion drive emits only an eerie blue glow as ionized atoms of xenon are pushed out of the engine. While slow to pick up speed, over the long haul it can deliver 10 times as much thrust per pound of fuel as liquid or solid fuel rockets. Other onboard experiments include software that tracks celestial bodies so the spacecraft can make its own navigation decisions without the intervention of ground controllers. Deep Space 1 will complete most of its mission objectives within the first two months, but will also do a flyby of a near-Earth asteroid, 1992 KD, in July 1999. Deep Space 1 will be launched aboard a Boeing Delta 7326 rocket from Launch Pad 17A, CCAS, in October

KSC-98pc1262

NASA Image and Video Library

1998-10-07

KENNEDY SPACE CENTER, FLA. -- Workers at the Defense Satellite Communications System Processing Facility (DPF), Cape Canaveral Air Station (CCAS), make adjustments while installing the ion propulsion engine on Deep Space 1. The first flight in NASA's New Millennium Program, Deep Space 1 is designed to validate 12 new technologies for scientific space missions of the next century, including the engine. Propelled by the gas xenon, the engine is being flight-tested for future deep space and Earth-orbiting missions. Deceptively powerful, the ion drive emits only an eerie blue glow as ionized atoms of xenon are pushed out of the engine. While slow to pick up speed, over the long haul it can deliver 10 times as much thrust per pound of fuel as liquid or solid fuel rockets. Other onboard experiments include software that tracks celestial bodies so the spacecraft can make its own navigation decisions without the intervention of ground controllers. Deep Space 1 will complete most of its mission objectives within the first two months, but will also do a flyby of a near-Earth asteroid, 1992 KD, in July 1999. Deep Space 1 will be launched aboard a Boeing Delta 7326 rocket from Launch Pad 17A, CCAS, in October

KSC-98pc1264

NASA Image and Video Library

1998-10-07

KENNEDY SPACE CENTER, FLA. -- Workers in the Defense Satellite Communications Systems Processing Facility (DPF) at Cape Canaveral Air Station (CCAS) make adjustments while installing the ion propulsion engine on Deep Space 1. The first flight in NASA's New Millennium Program, Deep Space 1 is designed to validate 12 new technologies for scientific space missions of the next century, including the engine. Propelled by the gas xenon, the engine is being flight-tested for future deep space and Earth-orbiting missions. Deceptively powerful, the ion drive emits only an eerie blue glow as ionized atoms of xenon are pushed out of the engine. While slow to pick up speed, over the long haul it can deliver 10 times as much thrust per pound of fuel as liquid or solid fuel rockets. Other onboard experiments include software that tracks celestial bodies so the spacecraft can make its own navigation decisions without the intervention of ground controllers. Deep Space 1 will complete most of its mission objectives within the first two months, but will also do a flyby of a near-Earth asteroid, 1992 KD, in July 1999. Deep Space 1 will be launched Oct. 25 aboard a Boeing Delta 7326 rocket from Launch Pad 17A, CCAS

KSC-98pc1260

NASA Image and Video Library

1998-10-07

KENNEDY SPACE CENTER, FLA. -- Workers at the Defense Satellite Communications System Processing Facility (DPF), Cape Canaveral Air Station (CCAS), install an ion propulsion engine on Deep Space 1. The first flight in NASA's New Millennium Program, Deep Space 1 is designed to validate 12 new technologies for scientific space missions of the next century, including the engine. Propelled by the gas xenon, the engine is being flight-tested for future deep space and Earth-orbiting missions. Deceptively powerful, the ion drive emits only an eerie blue glow as ionized atoms of xenon are pushed out of the engine. While slow to pick up speed, over the long haul it can deliver 10 times as much thrust per pound of fuel as liquid or solid fuel rockets. Other onboard experiments include software that tracks celestial bodies so the spacecraft can make its own navigation decisions without the intervention of ground controllers. Deep Space 1 will complete most of its mission objectives within the first two months, but will also do a flyby of a near-Earth asteroid, 1992 KD, in July 1999. Deep Space 1 will be launched aboard a Boeing Delta 7326 rocket from Launch Pad 17A, CCAS, in October

KSC-98pc1265

NASA Image and Video Library

1998-10-07

KENNEDY SPACE CENTER, FLA. -- Workers in the Defense Satellite Communications Systems Processing Facility (DPF) at Cape Canaveral Air Station (CCAS) finish installing the ion propulsion engine on Deep Space 1. The first flight in NASA's New Millennium Program, Deep Space 1 is designed to validate 12 new technologies for scientific space missions of the next century, including the engine. Propelled by the gas xenon, the engine is being flight-tested for future deep space and Earth-orbiting missions. Deceptively powerful, the ion drive emits only an eerie blue glow as ionized atoms of xenon are pushed out of the engine. While slow to pick up speed, over the long haul it can deliver 10 times as much thrust per pound of fuel as liquid or solid fuel rockets. Other onboard experiments include software that tracks celestial bodies so the spacecraft can make its own navigation decisions without the intervention of ground controllers. Deep Space 1 will complete most of its mission objectives within the first two months, but will also do a flyby of a near-Earth asteroid, 1992 KD, in July 1999. Deep Space 1 will be launched Oct. 25 aboard a Boeing Delta 7326 rocket from Launch Pad 17A, CCAS

KSC-98pc1263

NASA Image and Video Library

1998-10-07

KENNEDY SPACE CENTER, FLA. -- Workers at the Defense Satellite Communications System Processing Facility (DPF), Cape Canaveral Air Station (CCAS), maneuver the ion propulsion engine into place before installation on Deep Space 1. The first flight in NASA's New Millennium Program, Deep Space 1 is designed to validate 12 new technologies for scientific space missions of the next century, including the engine. Propelled by the gas xenon, the engine is being flight-tested for future deep space and Earth-orbiting missions. Deceptively powerful, the ion drive emits only an eerie blue glow as ionized atoms of xenon are pushed out of the engine. While slow to pick up speed, over the long haul it can deliver 10 times as much thrust per pound of fuel as liquid or solid fuel rockets. Other onboard experiments include software that tracks celestial bodies so the spacecraft can make its own navigation decisions without the intervention of ground controllers. Deep Space 1 will complete most of its mission objectives within the first two months, but will also do a flyby of a near-Earth asteroid, 1992 KD, in July 1999. Deep Space 1 will be launched aboard a Boeing Delta 7326 rocket from Launch Pad 17A, CCAS, in October

Register of specialized sources for information on selected fuels and oxidizers. [rocket propellants, bibliographies

NASA Technical Reports Server (NTRS)

Ludtke, P. R.

1975-01-01

Thirty-eight (38) organizations are listed and described that catalog and file information in their data systems on fuel and oxidizers. The fuels include hydrogen, methane and hydrazine-type fuels; the oxidizers include oxygen, fluorine, flox, nitrogen tetroxide and ozone. The type of available information covers thermophysical properties, propellant systems, propellant fires-control-extinguishment, propellant explosions, propellant combustion, propellant safety, and fluorine chemistry. These organizations have assembled and collated their information so that it will be useful in the solution of engineering problems.

Development, Verification and Validation of Parallel, Scalable Volume of Fluid CFD Program for Propulsion Applications

NASA Technical Reports Server (NTRS)

West, Jeff; Yang, H. Q.

2014-01-01

There are many instances involving liquid/gas interfaces and their dynamics in the design of liquid engine powered rockets such as the Space Launch System (SLS). Some examples of these applications are: Propellant tank draining and slosh, subcritical condition injector analysis for gas generators, preburners and thrust chambers, water deluge mitigation for launch induced environments and even solid rocket motor liquid slag dynamics. Commercially available CFD programs simulating gas/liquid interfaces using the Volume of Fluid approach are currently limited in their parallel scalability. In 2010 for instance, an internal NASA/MSFC review of three commercial tools revealed that parallel scalability was seriously compromised at 8 cpus and no additional speedup was possible after 32 cpus. Other non-interface CFD applications at the time were demonstrating useful parallel scalability up to 4,096 processors or more. Based on this review, NASA/MSFC initiated an effort to implement a Volume of Fluid implementation within the unstructured mesh, pressure-based algorithm CFD program, Loci-STREAM. After verification was achieved by comparing results to the commercial CFD program CFD-Ace+, and validation by direct comparison with data, Loci-STREAM-VoF is now the production CFD tool for propellant slosh force and slosh damping rate simulations at NASA/MSFC. On these applications, good parallel scalability has been demonstrated for problems sizes of tens of millions of cells and thousands of cpu cores. Ongoing efforts are focused on the application of Loci-STREAM-VoF to predict the transient flow patterns of water on the SLS Mobile Launch Platform in order to support the phasing of water for launch environment mitigation so that vehicle determinantal effects are not realized.

Dual nozzle design update. [on liquid rocket engines for advanced earth-to-orbit transportation systems

NASA Technical Reports Server (NTRS)

Obrien, C. J.

1982-01-01

Dual-nozzle engines, such as the dual-throat and dual-expander engines, are being evaluated for advanced earth-to-orbit transportation systems. Potential derivatives of the Space Shuttle and completely new vehicles might benefit from these advanced engines. In this paper, progress in the design of single-fuel and dual-fuel dual-nozzle engines is summarized. Dual-nozzle engines include those burning propellants such as LOX/RP-1/LH2, LOX/LC3H8/LH2, LOX/LCH4/LH2, LOX/LH2/LH2, LOX/LCH4/LCH4, LOX/LC3H8/C3H8 and N2O4/MMH/LH2. Engine data are applicable for thrust levels from 200,000 through 670,000 lbF. The results indicate that several versions of these engines utilize state-of-the-art technology and that even advanced versions of these engines do not require a major breakthrough in technology.

The Initial Atmospheric Transport (IAT) Code: Description and Validation

DOE Office of Scientific and Technical Information (OSTI.GOV)

Morrow, Charles W.; Bartel, Timothy James

The Initial Atmospheric Transport (IAT) computer code was developed at Sandia National Laboratories as part of their nuclear launch accident consequences analysis suite of computer codes. The purpose of IAT is to predict the initial puff/plume rise resulting from either a solid rocket propellant or liquid rocket fuel fire. The code generates initial conditions for subsequent atmospheric transport calculations. The Initial Atmospheric Transfer (IAT) code has been compared to two data sets which are appropriate to the design space of space launch accident analyses. The primary model uncertainties are the entrainment coefficients for the extended Taylor model. The Titan 34Dmore » accident (1986) was used to calibrate these entrainment settings for a prototypic liquid propellant accident while the recent Johns Hopkins University Applied Physics Laboratory (JHU/APL, or simply APL) large propellant block tests (2012) were used to calibrate the entrainment settings for prototypic solid propellant accidents. North American Meteorology (NAM )formatted weather data profiles are used by IAT to determine the local buoyancy force balance. The IAT comparisons for the APL solid propellant tests illustrate the sensitivity of the plume elevation to the weather profiles; that is, the weather profile is a dominant factor in determining the plume elevation. The IAT code performed remarkably well and is considered validated for neutral weather conditions.« less

Sea-Level Flight Demonstration and Altitude Characterization of a LO2/LCH4 Based Accent Propulsion Lander

NASA Technical Reports Server (NTRS)

Collins, Jacob; Hurlbert, Eric; Romig, Kris; Melcher, John; Hobson, Aaron; Eaton, Phil

2009-01-01

A 1,500 lbf thrust-class liquid oxygen (LO2)/Liquid Methane (LCH4) rocket engine was developed and tested at both sea-level and simulated altitude conditions. The engine was fabricated by Armadillo Aerospace (AA) in collaboration with NASA Johnson Space Center. Sea level testing was conducted at Armadillo Aerospace facilities at Caddo Mills, TX. Sea-level tests were conducted using both a static horizontal test bed and a vertical take-off and landing (VTOL) test bed capable of lift-off and hover-flight in low atmosphere conditions. The vertical test bed configuration is capable of throttling the engine valves to enable liftoff and hover-flight. Simulated altitude vacuum testing was conducted at NASA Johnson Space Center White Sands Test Facility (WSTF), which is capable of providing altitude simulation greater than 120,000 ft equivalent. The engine tests demonstrated ignition using two different methods, a gas-torch and a pyrotechnic igniter. Both gas torch and pyrotechnic ignition were demonstrated at both sea-level and vacuum conditions. The rocket engine was designed to be configured with three different nozzle configurations, including a dual-bell nozzle geometry. Dual-bell nozzle tests were conducted at WSTF and engine performance data was achieved at both ambient pressure and simulated altitude conditions. Dual-bell nozzle performance data was achieved over a range of altitude conditions from 90,000 ft to 50,000 ft altitude. Thrust and propellant mass flow rates were measured in the tests for specific impulse (Isp) and C* calculations.

Recent Advances and Applications in Cryogenic Propellant Densification Technology

NASA Technical Reports Server (NTRS)

Tomsik, Thomas M.

2000-01-01

This purpose of this paper is to review several historical cryogenic test programs that were conducted at the NASA Glenn Research Center (GRC), Cleveland, Ohio over the past fifty years. More recently these technology programs were intended to study new and improved denser forms of liquid hydrogen (LH2) and liquid oxygen (LO2) cryogenic rocket fuels. Of particular interest are subcooled cryogenic propellants. This is due to the fact that they have a significantly higher density (eg. triple-point hydrogen, slush etc.), a lower vapor pressure and improved cooling capacity over the normal boiling point cryogen. This paper, which is intended to be a historical technology overview, will trace the past and recent development and testing of small and large-scale propellant densification production systems. Densifier units in the current GRC fuels program, were designed and are capable of processing subcooled LH2 and L02 propellant at the X33 Reusable Launch Vehicle (RLV) scale. One final objective of this technical briefing is to discuss some of the potential benefits and application which propellant densification technology may offer the industrial cryogenics production and end-user community. Density enhancements to cryogenic propellants (LH2, LO2, CH4) in rocket propulsion and aerospace application have provided the opportunity to either increase performance of existing launch vehicles or to reduce the overall size, mass and cost of a new vehicle system.

Nontoxic Ionic Liquid Fuels for Exploration Applications

NASA Technical Reports Server (NTRS)

Coil, Millicent

2015-01-01

The toxicity of propellants used in conventional propulsion systems increases not only safety risks to personnel but also costs, due to special handling required during the entire lifetime of the propellants. Orbital Technologies Corporation (ORBITEC) has developed and tested novel nontoxic ionic liquid fuels for propulsion applications. In Phase I of the project, the company demonstrated the feasibility of several ionic liquid formulations that equaled the performance of conventional rocket propellant monomethylhydrazine (MMH) and also provided low volatility and low toxicity. In Phase II, ORBITEC refined the formulations, conducted material property tests, and investigated combustion behavior in droplet and microreactor experiments. The company also explored the effect of injector design on performance and demonstrated the fuels in a small-scale thruster. The ultimate goal is to replace propellants such as MMH with fuels that are simultaneously high-performance and nontoxic. The fuels will have uses in NASA's propulsion applications and also in a range of military and commercial functions.

Liquid rocket combustor computer code development

NASA Technical Reports Server (NTRS)

Liang, P. Y.

1985-01-01

The Advanced Rocket Injector/Combustor Code (ARICC) that has been developed to model the complete chemical/fluid/thermal processes occurring inside rocket combustion chambers are highlighted. The code, derived from the CONCHAS-SPRAY code originally developed at Los Alamos National Laboratory incorporates powerful features such as the ability to model complex injector combustion chamber geometries, Lagrangian tracking of droplets, full chemical equilibrium and kinetic reactions for multiple species, a fractional volume of fluid (VOF) description of liquid jet injection in addition to the gaseous phase fluid dynamics, and turbulent mass, energy, and momentum transport. Atomization and droplet dynamic models from earlier generation codes are transplated into the present code. Currently, ARICC is specialized for liquid oxygen/hydrogen propellants, although other fuel/oxidizer pairs can be easily substituted.

U.S. Strategic Nuclear Forces: Background, Developments, and Issues

DTIC Science & Technology

2016-09-27

meet the terms of the New START Treaty. The Air Force is also modernizing the Minuteman missiles, replacing and upgrading their rocket motors...began in 1998 and has been replacing the propellant, the solid rocket fuel, in the Minuteman motors to extend the life of the rocket motors. A...complete the program. It has not requested additional funding in subsequent years. Propulsion System Rocket Engine Program (PSRE) According to the Air

The 2003 Goddard Rocket Replica Project: A Reconstruction of the World's First Functional Liquid Rocket System

NASA Technical Reports Server (NTRS)

Farr, R. A.; Elam, S. K.; Hicks, G. D.; Sanders, T. M.; London, J. R.; Mayne, A. W.; Christensen, D. L.

2003-01-01

As a part of NASA s 2003 Centennial of Flight celebration, engineers and technicians at Marshall Space Flight Center (MSFC), Huntsville, Alabama, in cooperation with the Alabama-Mississippi AIAA Section, have reconstructed historically accurate, functional replicas of Dr. Robert H. Goddard s 1926 first liquid- fuel rocket. The purposes of this project were to clearly understand, recreate, and document the mechanisms and workings of the 1926 rocket for exhibit and educational use, creating a vital resource for researchers studying the evolution of liquid rocketry for years to come. The MSFC team s reverse engineering activity has created detailed engineering-quality drawings and specifications describing the original rocket and how it was built, tested, and operated. Static hot-fire tests, as well as flight demonstrations, have further defined and quantified the actual performance and engineering actual performance and engineering challenges of this major segment in early aerospace history.

Innovative smart micro sensors for Army weaponry applications

NASA Astrophysics Data System (ADS)

Ruffin, Paul B.; Brantley, Christina; Edwards, Eugene

2008-03-01

Micro sensors offer the potential solution to cost, size, and weight issues associated with smart networked sensor systems designed for environmental/missile health monitoring and rocket out-gassing/fuel leak detection, as well as situational awareness on the battlefield. In collaboration with the University of Arkansas (Fayetteville), University of Alabama (Tuscaloosa and Birmingham), Alabama A&M University (Normal), and Streamline Automation (Huntsville, AL), scientists and engineers at the Army Aviation & Missile Research, Development, and Engineering Center (AMRDEC) are investigating several nano-based technologies to solve the problem of sensing extremely small levels of toxic gases associated with both chemical warfare agents (in air and liquids) and potential rocket motor leaks. Innovative techniques are being devised to adapt voltammetry, which is a well established technique for the detection and quantification of substances dissolved in liquids, to low-cost micro sensors for detecting airborne chemical agents and potential missile propellant leakages. In addition, a surface enhanced Raman scattering (SERS) technique, which enhances Raman scattered light by excitation of surface plasmons on nanoporous metal surfaces (nanospheres), is being investigated to develop novel smart sensors for the detection of chemical agents (including rocket motor out-gassing) and potential detection of home-made explosive devices. In this paper, results are delineated that are associated with experimental studies, which are conducted for the aforementioned cases and for several other nano-based technology approaches. The design challenges of each micro sensor technology approach are discussed. Finally, a comparative analysis of the various innovative micro-sensor techniques is provided.

Redstone Missile

NASA Technical Reports Server (NTRS)

2004-01-01

The image depicts Redstone missile being erected. The Redstone ballistic missile was a high-accuracy, liquid-propelled, surface-to-surface missile developed by Army Ballistic Missile Agency, Redstone Arsenal, in Huntsville, Alabama, under the direction of Dr. von Braun. The Redstone engine was a modified and improved version of the Air Force's Navaho cruise missile engine of the late forties. The A-series, as this would be known, utilized a cylindrical combustion chamber as compared with the bulky, spherical V-2 chamber. By 1951, the Army was moving rapidly toward the design of the Redstone missile, and the production was begun in 1952. Redstone rockets became the 'reliable workhorse' for America's early space program. As an example of the versatility, Redstone was utilized in the booster for Explorer 1, the first American satellite, with no major changes to the engine or missile

Testing and Functions of the J2X Gas Generator

NASA Technical Reports Server (NTRS)

Miller, Nicholas

2009-01-01

The Ares I, NASA s new solid rocket based crew launch vehicle, is a two stage in line rocket that has made its waytothe forefront of NASA s endeavors. The Ares I s Upper Stage (US) will be propelled by a J-2X engine which is fueled by liquid hydrogen and liquid oxygen. The J-2X is a variation based on two of its predecessor s, the J-2 and J-2S engines. ET50 is providing the design support for hardware required to run tests on the J-2X Gas Generator (GG) that increases the delivery pressure of the supplied combustion fuels that the engine burns. The test area will be running a series of tests using different lengths and curved segments of pipe and different sized nozzles to determine the configuration that best satisfies the thrust, heat, and stability requirements for the engine. I have had to research the configurations that are being tested and gain an understanding of the purpose of the tests. I then had to research the parts that would be used in the test configurations. I was taken to see parts similar to the ones used in the test configurations and was allowed to review drawings and dimensions used for those parts. My job over this summer has been to use the knowledge I have gained to design, model, and create drawings for the un-fabricated parts that are necessary for the J-2X Workhorse Gas Generator Phase IIcTest.

Study of solid rocket motors for a space shuttle booster. Volume 4: Mass properties report

NASA Technical Reports Server (NTRS)

Vonderesch, A. H.

1972-01-01

Mass properties data for the 156 inch diameter, parallel burn, solid propellant rocket engine for the space shuttle booster are presented. Design ground rules and assumptions applicable to generation of the mass properties data are described, together with pertinent data sources.

Concepts for the design of an antimatter annihilation rocket

NASA Technical Reports Server (NTRS)

Morgan, D. L., Jr.

1982-01-01

Matter-antimatter annihilation is considered for spacecraft propulsion. Annihilation produces considerably more energy per unit mass of propellant than any other known means of energy production. An antimatter annihilation rocket requires several systems and components that are unique to its nature. Among these are an antimatter storage system, a means to extract the antimatter from storage, a system to transport the antimatter to the rocket engine, and the engine wherein annihilation occurs and thrust is produced. Design concepts of these systems and components are presented and discussed.

Launch Vehicles

NASA Image and Video Library

2007-09-09

Under the goals of the Vision for Space Exploration, Ares I is a chief component of the cost-effective space transportation infrastructure being developed by NASA's Constellation Program. This transportation system will safely and reliably carry human explorers back to the moon, and then onward to Mars and other destinations in the solar system. The Ares I effort includes multiple project element teams at NASA centers and contract organizations around the nation, and is managed by the Exploration Launch Projects Office at NASA's Marshall Space Flight Center (MFSC). ATK Launch Systems near Brigham City, Utah, is the prime contractor for the first stage booster. ATK's subcontractor, United Space Alliance of Houston, is designing, developing and testing the parachutes at its facilities at NASA's Kennedy Space Center in Florida. NASA's Johnson Space Center in Houston hosts the Constellation Program and Orion Crew Capsule Project Office and provides test instrumentation and support personnel. Together, these teams are developing vehicle hardware, evolving proven technologies, and testing components and systems. Their work builds on powerful, reliable space shuttle propulsion elements and nearly a half-century of NASA space flight experience and technological advances. Ares I is an inline, two-stage rocket configuration topped by the Crew Exploration Vehicle, its service module, and a launch abort system. The launch vehicle's first stage is a single, five-segment reusable solid rocket booster derived from the Space Shuttle Program's reusable solid rocket motor that burns a specially formulated and shaped solid propellant called polybutadiene acrylonitrile (PBAN). The second or upper stage will be propelled by a J-2X main engine fueled with liquid oxygen and liquid hydrogen. This HD video image depicts a test firing of a 40k subscale J2X injector at MSFC's test stand 115. (Highest resolution available)

Boundary cooled rocket engines for space storable propellants

NASA Technical Reports Server (NTRS)

Kesselring, R. C.; Mcfarland, B. L.; Knight, R. M.; Gurnitz, R. N.

1972-01-01

An evaluation of an existing analytical heat transfer model was made to develop the technology of boundary film/conduction cooled rocket thrust chambers to the space storable propellant combination oxygen difluoride/diborane. Critical design parameters were identified and their importance determined. Test reduction methods were developed to enable data obtained from short duration hot firings with a thin walled (calorimeter) chamber to be used quantitatively evaluate the heat absorbing capability of the vapor film. The modification of the existing like-doublet injector was based on the results obtained from the calorimeter firings.

Design criteria monograph for pressurized metal cases

NASA Technical Reports Server (NTRS)

1972-01-01

Organiation and presentation of data pertaining to design of solid propellant rocket engine cases are discussed. Design criteria are presented in form of monograph based on accumulated experience and knowledge. Improvements in reliability, cost effectiveness, and engine efficiency are stressed.

On Nonlinear Combustion Instability in Liquid Propellant Rocket Motors

NASA Technical Reports Server (NTRS)

Sims, J. D. (Technical Monitor); Flandro, Gary A.; Majdalani, Joseph; Sims, Joseph D.

2004-01-01

All liquid propellant rocket instability calculations in current use have limited value in the predictive sense and serve mainly as a correlating framework for the available data sets. The well-known n-t model first introduced by Crocco and Cheng in 1956 is still used as the primary analytical tool of this type. A multitude of attempts to establish practical analytical methods have achieved only limited success. These methods usually produce only stability boundary maps that are of little use in making critical design decisions in new motor development programs. Recent progress in understanding the mechanisms of combustion instability in solid propellant rockets"' provides a firm foundation for a new approach to prediction, diagnosis, and correction of the closely related problems in liquid motor instability. For predictive tools to be useful in the motor design process, they must have the capability to accurately determine: 1) time evolution of the pressure oscillations and limit amplitude, 2) critical triggering pulse amplitude, and 3) unsteady heat transfer rates at injector surfaces and chamber walls. The method described in this paper relates these critical motor characteristics directly to system design parameters. Inclusion of mechanisms such as wave steepening, vorticity production and transport, and unsteady detonation wave phenomena greatly enhance the representation of key features of motor chamber oscillatory behavior. The basic theoretical model is described and preliminary computations are compared to experimental data. A plan to develop the new predictive method into a comprehensive analysis tool is also described.

Combustion Instability Analysis and the Effects of Drop Size on Acoustic Driving Rocket Flow

NASA Technical Reports Server (NTRS)

Harper, Brent (Technical Monitor); Ellison, L. Renea; Moser, Marlow D.

2004-01-01

High frequency combustion instability, the most destructive kind, is generally solved on a per engine basis. The instability often is the result of compounding acoustic oscillations, usually from the propellant combustion itself. To counteract the instability the chamber geometry can be changed and/or the method of propellant injection can be altered. This experiment will alter the chamber dimensions slightly; using a cylindrical shape of constant diameter and the length will be varied from six to twelve inches in three-inch increments. The main flowfield will be the products of a high OF hydrogen/oxygen flow. The liquid fuel will be injected into this flowfield using a modulated injector. It will allow for varied droplet size, feed rate, spray pattern, and location for the mixture within the chamber. The response will be deduced from the chamber pressure oscillations.

Additive Manufacturing for Affordable Rocket Engines

NASA Technical Reports Server (NTRS)

West, Brian; Robertson, Elizabeth; Osborne, Robin; Calvert, Marty

2016-01-01

Additive manufacturing (also known as 3D printing) technology has the potential to drastically reduce costs and lead times associated with the development of complex liquid rocket engine systems. NASA is using 3D printing to manufacture rocket engine components including augmented spark igniters, injectors, turbopumps, and valves. NASA is advancing the process to certify these components for flight. Success Story: MSFC has been developing rocket 3D-printing technology using the Selective Laser Melting (SLM) process. Over the last several years, NASA has built and tested several injectors and combustion chambers. Recently, MSFC has 3D printed an augmented spark igniter for potential use the RS-25 engines that will be used on the Space Launch System. The new design is expected to reduce the cost of the igniter by a factor of four. MSFC has also 3D printed and tested a liquid hydrogen turbopump for potential use on an Upper Stage Engine. Additive manufacturing of the turbopump resulted in a 45% part count reduction. To understanding how the 3D printed parts perform and to certify them for flight, MSFC built a breadboard liquid rocket engine using additive manufactured components including injectors, turbomachinery, and valves. The liquid rocket engine was tested seven times in 2016 using liquid oxygen and liquid hydrogen. In addition to exposing the hardware to harsh environments, engineers learned to design for the new manufacturing technique, taking advantage of its capabilities and gaining awareness of its limitations. Benefit: The 3D-printing technology promises reduced cost and schedule for rocket engines. Cost is a function of complexity, and the most complicated features provide the largest opportunities for cost reductions. This is especially true where brazes or welds can be eliminated. The drastic reduction in part count achievable with 3D printing creates a waterfall effect that reduces the number of processes and drawings, decreases the amount of touch labor required, and increases reliability. When certification is achieved, NASA missions will be able to realize these benefits.

Experimental Methodology for Measuring Combustion and Injection-Coupled Responses

NASA Technical Reports Server (NTRS)

Cavitt, Ryan C.; Frederick, Robert A.; Bazarov, Vladimir G.

2006-01-01

A Russian scaling methodology for liquid rocket engines utilizing a single, full scale element is reviewed. The scaling methodology exploits the supercritical phase of the full scale propellants to simplify scaling requirements. Many assumptions are utilized in the derivation of the scaling criteria. A test apparatus design is presented to implement the Russian methodology and consequently verify the assumptions. This test apparatus will allow researchers to assess the usefulness of the scaling procedures and possibly enhance the methodology. A matrix of the apparatus capabilities for a RD-170 injector is also presented. Several methods to enhance the methodology have been generated through the design process.

Assessing Spontaneous Combustion Instability with Recurrence Quantification Analysis

NASA Technical Reports Server (NTRS)

Eberhart, Chad J.; Casiano, Matthew J.

2016-01-01

Spontaneous instabilities can pose a significant challenge to verification of combustion stability, and characterizing its onset is an important avenue of improvement for stability assessments of liquid propellant rocket engines. Recurrence Quantification Analysis (RQA) is used here to explore nonlinear combustion dynamics that might give insight into instability. Multiple types of patterns representative of different dynamical states are identified within fluctuating chamber pressure data, and markers for impending instability are found. A class of metrics which describe these patterns is also calculated. RQA metrics are compared with and interpreted against another metric from nonlinear time series analysis, the Hurst exponent, to help better distinguish between stable and unstable operation.

Advanced cooling techniques for high-pressure hydrocarbon-fueled engines

NASA Technical Reports Server (NTRS)

Cook, R. T.

1979-01-01

The regenerative cooling limits (maximum chamber pressure) for 02/hydrocarbon gas generator and staged combustion cycle rocket engines over a thrust range of 89,000 N (20,000lbf) to 2,669,000 N (600,000 lbf) for a reusable life of 250 missions were defined. Maximum chamber pressure limits were first determined for the three propellant combinations (O2/CH4, O2/C3H8, and O2/RP-1 without a carbon layer (unenhanced designs). Chamber pressure cooling enhancement limits were then established for seven thermal barriers. The thermal barriers evaluated for these designs were: carbon layer, ceramic coating, graphite liner, film cooling, transpiration cooling, zoned combustion, and a combination of two of the above. All fluid barriers were assessed a 3 percent performance loss. Sensitivity studies were then conducted to determine the influence of cycle life and RP-1 decomposition temperature on chamber pressure limits. Chamber and nozzle design parameters are presented for the unenahanced and enhanced designs. The maximum regenerative cooled chamber pressure limits were attained with the O2/CH4 propellant combination. The O2/RP-1 designs relied on a carbon layer and liquid gas injection chamber contours, short chamber, to be competitive with the other two propellant combinations. This was attributed to the low decomposition temperature of RP-1.

Hybrid Rocket Motor Test

NASA Technical Reports Server (NTRS)

1994-01-01

A 10,000-pound thrust hybrid rocket motor is tested at Stennis Space Center's E-1 test facility. A hybrid rocket motor is a cross between a solid rocket and a liquid-fueled engine. It uses environmentally safe solid fuel and liquid oxygen.

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

NASA Technical Reports Server (NTRS)

Ballard, Richard O.; Brown, Kendall K.

2005-01-01

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.

Development of life prediction capabilities for liquid propellant rocket engines. Post-fire diagnostic system for the SSME system architecture study

NASA Technical Reports Server (NTRS)

Gage, Mark; Dehoff, Ronald

1991-01-01

This system architecture task (1) analyzed the current process used to make an assessment of engine and component health after each test or flight firing of an SSME, (2) developed an approach and a specific set of objectives and requirements for automated diagnostics during post fire health assessment, and (3) listed and described the software applications required to implement this system. The diagnostic system described is a distributed system with a database management system to store diagnostic information and test data, a CAE package for visual data analysis and preparation of plots of hot-fire data, a set of procedural applications for routine anomaly detection, and an expert system for the advanced anomaly detection and evaluation.

Space Shuttle Projects

NASA Image and Video Library

1989-06-03

The Marshall Space Flight Center (MSFC) engineers test fired a 26-foot long, 100,000-pound-thrust solid rocket motor for 30 seconds at the MSFC east test area, the first test firing of the Modified NASA Motor (M-NASA Motor). The M-NASA Motor was fired in a newly constructed stand. The motor is 48-inches in diameter and was loaded with two propellant cartridges weighing a total of approximately 12,000 pounds. The purpose of the test was to learn more about solid rocket motor insulation and nozzle materials and to provide young engineers additional hands-on expertise in solid rocket motor technology. The test is a part of NASA's Solid Propulsion Integrity Program, that is to provide NASA engineers with the techniques, engineering tools, and computer programs to be able to better design, build, and verify solid rocket motors.

Study of solid rocket motors for a space shuttle booster. Volume 2, book 3: Cost estimating data

NASA Technical Reports Server (NTRS)

Vanderesch, A. H.

1972-01-01

Cost estimating data for the 156 inch diameter, parallel burn solid rocket propellant engine selected for the space shuttle booster are presented. The costing aspects on the baseline motor are initially considered. From the baseline, sufficient data is obtained to provide cost estimates of alternate approaches.

Experimental Study on an Unsteady Pressure Gain Combustion Hypergolic Rocket Engine Concept

NASA Astrophysics Data System (ADS)

Kan, Brandon K.

An experimental study is conducted to investigate pulsed combustion in a lab-scale bipropellant rocket engine using hypergolic propellants. The propellant combination is high concentration hydrogen peroxide and a catalyst-laced triglyme fuel. A total of 50 short duration firings have been conducted; the vast majority in an open-chamber configuration. High amplitude pulsations were evident in nearly all cases and have been assessed with high frequency pressure measurements. Both pintle and unlike impinging quadlet injector types have been evaluated although the bulk of the testing was with the latter configuration. Several firings were conducted with a transparent chamber in an attempt to gain understanding using a high-speed camera in the visible spectrum. Peak chamber pressures in excess of 5000 psi have been recorded with surface mounted high frequency gages with pulsation frequencies exceeding 600 Hz. A characterization of time-averaged performance is made for the unsteady system, where time-resolved thrust and pressure measurements were attempted. While prior literature describes this system as a pulse detonation rocket engine, the combustion appears to be more "constant volume" in nature.

On fundamentally new sources of energy for rockets in the early works of the pioneers of astronautics

NASA Technical Reports Server (NTRS)

Melkumov, T. M.

1977-01-01

The research for more efficient methods of propelling a spacecraft, than can be achieved with chemical energy, was studied. During a time when rockets for space flight had not actually been built pioneers in rocket technology were already concerned with this problem. Alternative sources proposed at that time, were nuclear and solar energy. Basic engineering problems of each source were investigated.

Catalytic Microtube Rocket Igniter

NASA Technical Reports Server (NTRS)

Schneider, Steven J.; Deans, Matthew C.

2011-01-01

Devices that generate both high energy and high temperature are required to ignite reliably the propellant mixtures in combustion chambers like those present in rockets and other combustion systems. This catalytic microtube rocket igniter generates these conditions with a small, catalysis-based torch. While traditional spark plug systems can require anywhere from 50 W to multiple kW of power in different applications, this system has demonstrated ignition at less than 25 W. Reactants are fed to the igniter from the same tanks that feed the reactants to the rest of the rocket or combustion system. While this specific igniter was originally designed for liquid methane and liquid oxygen rockets, it can be easily operated with gaseous propellants or modified for hydrogen use in commercial combustion devices. For the present cryogenic propellant rocket case, the main propellant tanks liquid oxygen and liquid methane, respectively are regulated and split into different systems for the individual stages of the rocket and igniter. As the catalyst requires a gas phase for reaction, either the stored boil-off of the tanks can be used directly or one stream each of fuel and oxidizer can go through a heat exchanger/vaporizer that turns the liquid propellants into a gaseous form. For commercial applications, where the reactants are stored as gases, the system is simplified. The resulting gas-phase streams of fuel and oxidizer are then further divided for the individual components of the igniter. One stream each of the fuel and oxidizer is introduced to a mixing bottle/apparatus where they are mixed to a fuel-rich composition with an O/F mass-based mixture ratio of under 1.0. This premixed flow then feeds into the catalytic microtube device. The total flow is on the order of 0.01 g/s. The microtube device is composed of a pair of sub-millimeter diameter platinum tubes connected only at the outlet so that the two outlet flows are parallel to each other. The tubes are each approximately 10 cm long and are heated via direct electric resistive heating. This heating brings the gasses to their minimum required ignition temperature, which is lower than the auto-thermal ignition temperature, and causes the onset of both surface and gas phase ignition producing hot temperatures and a highly reacting flame. The combustion products from the catalytic tubes, which are below the melting point of platinum, are injected into the center of another combustion stage, called the primary augmenter. The reactants for this combustion stage come from the same source but the flows of non-premixed methane and oxygen gas are split off to a secondary mixing apparatus and can be mixed in a near-stoichiometric to highly lean mixture ratio. The primary augmenter is a component that has channels venting this mixed gas to impinge on each other in the center of the augmenter, perpendicular to the flow from the catalyst. The total crosssectional area of these channels is on a similar order as that of the catalyst. The augmenter has internal channels that act as a manifold to distribute equally the gas to the inward-venting channels. This stage creates a stable flame kernel as its flows, which are on the order of 0.01 g/s, are ignited by the combustion products of the catalyst. This stage is designed to produce combustion products in the flame kernel that exceed the autothermal ignition temperature of oxygen and methane.

Deep Space 1 is encapsulated on launch pad

NASA Technical Reports Server (NTRS)

1998-01-01

On Launch Pad 17A at Cape Canaveral Air Station, released from its protective payload transportation container, Deep Space 1 waits to have the fairing attached before launch. Targeted for launch aboard a Boeing Delta 7326 rocket on Oct. 25, Deep Space 1 is the first flight in NASA's New Millennium Program, and is designed to validate 12 new technologies for scientific space missions of the next century, including the engine. Propelled by the gas xenon, the engine is being flight-tested for future deep space and Earth-orbiting missions. Deceptively powerful, the ion drive emits only an eerie blue glow as ionized atoms of xenon are pushed out of the engine. While slow to pick up speed, over the long haul it can deliver 10 times as much thrust per pound of fuel as liquid or solid fuel rockets. Other onboard experiments include software that tracks celestial bodies so the spacecraft can make its own navigation decisions without the intervention of ground controllers. Deep Space 1 will complete most of its mission objectives within the first two months, but will also do a flyby of a near-Earth asteroid, 1992 KD, in July 1999.

Deep Space 1 is prepared for transport to launch pad

NASA Technical Reports Server (NTRS)

1998-01-01

Workers in the Defense Satellite Communication Systems Processing Facility (DPF), Cape Canaveral Air Station (CCAS), move to the workstand the second conical section leaf of the payload transportation container for Deep Space 1. The first flight in NASA's New Millennium Program, Deep Space 1 is designed to validate 12 new technologies for scientific space missions of the next century, including the engine. Propelled by the gas xenon, the engine is being flight-tested for future deep space and Earth-orbiting missions. Deceptively powerful, the ion drive emits only an eerie blue glow as ionized atoms of xenon are pushed out of the engine. While slow to pick up speed, over the long haul it can deliver 10 times as much thrust per pound of fuel as liquid or solid fuel rockets. Other onboard experiments include software that tracks celestial bodies so the spacecraft can make its own navigation decisions without the intervention of ground controllers. Deep Space 1 will complete most of its mission objectives within the first two months, but will also do a flyby of a near-Earth asteroid, 1992 KD, in July 1999. Deep Space 1 will be launched aboard a Boeing Delta 7326 rocket from Launch Pad 17A, CCAS.

Deep Space 1 is prepared for transport to launch pad

NASA Technical Reports Server (NTRS)

1998-01-01

Workers in the Defense Satellite Communication Systems Processing Facility (DPF), Cape Canaveral Air Station (CCAS), begin attaching the conical section leaves of the payload transportation container on Deep Space 1 before launch, targeted for Oct. 25 aboard a Boeing Delta 7326 rocket from Launch Pad 17A. The first flight in NASA's New Millennium Program, Deep Space 1 is designed to validate 12 new technologies for scientific space missions of the next century, including the engine. Propelled by the gas xenon, the engine is being flight- tested for future deep space and Earth-orbiting missions. Deceptively powerful, the ion drive emits only an eerie blue glow as ionized atoms of xenon are pushed out of the engine. While slow to pick up speed, over the long haul it can deliver 10 times as much thrust per pound of fuel as liquid or solid fuel rockets. Other onboard experiments include software that tracks celestial bodies so the spacecraft can make its own navigation decisions without the intervention of ground controllers. Deep Space 1 will complete most of its mission objectives within the first two months, but will also do a flyby of a near-Earth asteroid, 1992 KD, in July 1999.

KSC-98pc1328

NASA Image and Video Library

1998-10-12

KENNEDY SPACE CENTER, FLA. -- Wrapped in an anti-static blanket for protection, Deep Space 1 is moved out of the Defense Satellite Communications Systems Processing Facility (DPF) at Cape Canaveral Air Station (CCAS) for its trip to Launch Pad 17A. The spacecraft will be launched aboard a Boeing Delta 7326 rocket on Oct. 25. Deep Space 1 is the first flight in NASA's New Millennium Program, and is designed to validate 12 new technologies for scientific space missions of the next century, including the engine. Propelled by the gas xenon, the engine is being flight-tested for future deep space and Earth-orbiting missions. Deceptively powerful, the ion drive emits only an eerie blue glow as ionized atoms of xenon are pushed out of the engine. While slow to pick up speed, over the long haul it can deliver 10 times as much thrust per pound of fuel as liquid or solid fuel rockets. Other onboard experiments include software that tracks celestial bodies so the spacecraft can make its own navigation decisions without the intervention of ground controllers. Deep Space 1 will complete most of its mission objectives within the first two months, but will also do a flyby of a near-Earth asteroid, 1992 KD, in July 1999

KSC-98pc1345

NASA Image and Video Library

1998-10-16

KENNEDY SPACE CENTER, FLA. -- On Launch Pad 17A at Cape Canaveral Air Station, released from its protective payload transportation container, Deep Space 1 waits to have the fairing attached before launch. Targeted for launch aboard a Boeing Delta 7326 rocket on Oct. 25, Deep Space 1 is the first flight in NASA's New Millennium Program, and is designed to validate 12 new technologies for scientific space missions of the next century, including the engine. Propelled by the gas xenon, the engine is being flight-tested for future deep space and Earth-orbiting missions. Deceptively powerful, the ion drive emits only an eerie blue glow as ionized atoms of xenon are pushed out of the engine. While slow to pick up speed, over the long haul it can deliver 10 times as much thrust per pound of fuel as liquid or solid fuel rockets. Other onboard experiments include software that tracks celestial bodies so the spacecraft can make its own navigation decisions without the intervention of ground controllers. Deep Space 1 will complete most of its mission objectives within the first two months, but will also do a flyby of a near-Earth asteroid, 1992 KD, in July 1999

KSC-98pc1329

NASA Image and Video Library

1998-10-12

KENNEDY SPACE CENTER, FLA. -- Wrapped in an anti-static blanket for protection, Deep Space 1 is lifted out of the transporter that carried it to Launch Pad 17A at Cape Canaveral Air Station. The spacecraft will be launched aboard a Boeing Delta 7326 rocket on Oct. 25. Deep Space 1 is the first flight in NASA's New Millennium Program, and is designed to validate 12 new technologies for scientific space missions of the next century, including the engine. Propelled by the gas xenon, the engine is being flight-tested for future deep space and Earth-orbiting missions. Deceptively powerful, the ion drive emits only an eerie blue glow as ionized atoms of xenon are pushed out of the engine. While slow to pick up speed, over the long haul it can deliver 10 times as much thrust per pound of fuel as liquid or solid fuel rockets. Other onboard experiments include software that tracks celestial bodies so the spacecraft can make its own navigation decisions without the intervention of ground controllers. Deep Space 1 will complete most of its mission objectives within the first two months, but will also do a flyby of a near-Earth asteroid, 1992 KD, in July 1999

KSC-98pc1318

NASA Image and Video Library

1998-10-10

KENNEDY SPACE CENTER, FLA. - Wrapped in an antistatic blanket for protection, Deep Space 1 is moved out of the Defense Satellite Communications System Processing Facility (DPF) at Cape Canaveral Air Station (CCAS) for its trip to Launch Pad 17A. The spacecraft will be launched aboard Boeing's Delta 7326 rocket in October. The first flight in NASA's New Millennium Program, Deep Space 1 is designed to validate 12 new technologies for scientific space missions of the next century, including an ion propulsion engine. Propelled by the gas xenon, the engine is being flight tested for future deep space and Earth-orbiting missions. Deceptively powerful, the ion drive emits only an eerie blue glow as ionized atoms of xenon are pushed out of the engine. While slow to pick up speed, over the long haul it can deliver 10 times as much thrust per pound of fuel as liquid or solid fuel rockets. Other onboard experiments include softwre that tracks celestial bodies so the spacecraft can make its own navigation decisions without the intervention of ground controllers. Deep Space 1 will complete most of its mission objectives within the firs two months, but will also make a flyby of a near-Earth asteroid, 1992 KD, in July 1999.

KSC-98pc1313

NASA Image and Video Library

1998-10-10

KENNEDY SPACE CENTER, FLA. -- Workers in the Defense Satellite Communication Systems Processing Facility (DPF), Cape Canaveral Air Station (CCAS), begin attaching the conical section leaves of the payload transportation container on Deep Space 1 before launch, targeted for Oct. 25 aboard a Boeing Delta 7326 rocket from Launch Pad 17A. The first flight in NASA's New Millennium Program, Deep Space 1 is designed to validate 12 new technologies for scientific space missions of the next century, including the engine. Propelled by the gas xenon, the engine is being flight-tested for future deep space and Earth-orbiting missions. Deceptively powerful, the ion drive emits only an eerie blue glow as ionized atoms of xenon are pushed out of the engine. While slow to pick up speed, over the long haul it can deliver 10 times as much thrust per pound of fuel as liquid or solid fuel rockets. Other onboard experiments include software that tracks celestial bodies so the spacecraft can make its own navigation decisions without the intervention of ground controllers. Deep Space 1 will complete most of its mission objectives within the first two months, but will also do a flyby of a near-Earth asteroid, 1992 KD, in July 1999

KSC-98pc1332

NASA Image and Video Library

1998-10-12

KENNEDY SPACE CENTER, FLA. -- On Launch Pad 17A at Cape Canaveral Air Station, Deep Space 1 is lowered toward the second stage of a Boeing Delta 7326 rocket. The adapter on the spacecraft can be seen surrounding the booster motor. Targeted for launch on Oct. 25, Deep Space 1 is the first flight in NASA's New Millennium Program, and is designed to validate 12 new technologies for scientific space missions of the next century, including the engine. Propelled by the gas xenon, the engine is being flight-tested for future deep space and Earth-orbiting missions. Deceptively powerful, the ion drive emits only an eerie blue glow as ionized atoms of xenon are pushed out of the engine. While slow to pick up speed, over the long haul it can deliver 10 times as much thrust per pound of fuel as liquid or solid fuel rockets. Other onboard experiments include software that tracks celestial bodies so the spacecraft can make its own navigation decisions without the intervention of ground controllers. Deep Space 1 will complete most of its mission objectives within the first two months, but will also do a flyby of a near-Earth asteroid, 1992 KD, in July 1999

KSC-98pc1347

NASA Image and Video Library

1998-10-16

KENNEDY SPACE CENTER, FLA. -- On Launch Pad 17A at Cape Canaveral Air Station, workers maneuver part of the fairing (viewed from the inside) to encapsulate Deep Space 1. Targeted for launch aboard a Boeing Delta 7326 rocket on Oct. 25, Deep Space 1 is the first flight in NASA's New Millennium Program, and is designed to validate 12 new technologies for scientific space missions of the next century, including the engine. Propelled by the gas xenon, the engine is being flight-tested for future deep space and Earth-orbiting missions. Deceptively powerful, the ion drive emits only an eerie blue glow as ionized atoms of xenon are pushed out of the engine. While slow to pick up speed, over the long haul it can deliver 10 times as much thrust per pound of fuel as liquid or solid fuel rockets. Other onboard experiments include software that tracks celestial bodies so the spacecraft can make its own navigation decisions without the intervention of ground controllers. Deep Space 1 will complete most of its mission objectives within the first two months, but will also do a flyby of a near-Earth asteroid, 1992 KD, in July 1999

KSC-98pc1330

NASA Image and Video Library

1998-10-12

KENNEDY SPACE CENTER, FLA. -- Just before sunrise, on Launch Pad 17A at Cape Canaveral Air Station, Deep Space 1 is hoisted up the mobile service tower for installation on a Boeing Delta 7326 rocket . The spacecraft is targeted for launch on Oct. 25. Deep Space 1 is the first flight in NASA's New Millennium Program, and is designed to validate 12 new technologies for scientific space missions of the next century, including the engine. Propelled by the gas xenon, the engine is being flight-tested for future deep space and Earth-orbiting missions. Deceptively powerful, the ion drive emits only an eerie blue glow as ionized atoms of xenon are pushed out of the engine. While slow to pick up speed, over the long haul it can deliver 10 times as much thrust per pound of fuel as liquid or solid fuel rockets. Other onboard experiments include software that tracks celestial bodies so the spacecraft can make its own navigation decisions without the intervention of ground controllers. Deep Space 1 will complete most of its mission objectives within the first two months, but will also do a flyby of a near-Earth asteroid, 1992 KD, in July 1999

KSC-98pc1314

NASA Image and Video Library

1998-10-10

KENNEDY SPACE CENTER, FLA. -- Workers in the Defense Satellite Communication Systems Processing Facility (DPF), Cape Canaveral Air Station (CCAS), move to the workstand the second conical section leaf of the payload transportation container for Deep Space 1. The first flight in NASA's New Millennium Program, Deep Space 1 is designed to validate 12 new technologies for scientific space missions of the next century, including the engine. Propelled by the gas xenon, the engine is being flight-tested for future deep space and Earth-orbiting missions. Deceptively powerful, the ion drive emits only an eerie blue glow as ionized atoms of xenon are pushed out of the engine. While slow to pick up speed, over the long haul it can deliver 10 times as much thrust per pound of fuel as liquid or solid fuel rockets. Other onboard experiments include software that tracks celestial bodies so the spacecraft can make its own navigation decisions without the intervention of ground controllers. Deep Space 1 will complete most of its mission objectives within the first two months, but will also do a flyby of a near-Earth asteroid, 1992 KD, in July 1999. Deep Space 1 will be launched aboard a Boeing Delta 7326 rocket from Launch Pad 17A, CCAS

Particle size reduction of propellants by cryocycling

DOE Office of Scientific and Technical Information (OSTI.GOV)

Whinnery, L.; Griffiths, S.; Lipkin, J.

1995-05-01

Repeated exposure of a propellant to liquid nitrogen causes thermal stress gradients within the material resulting in cracking and particle size reduction. This process is termed cryocycling. The authors conducted a feasibility study, combining experiments on both inert and live propellants with three modeling approaches. These models provided optimized cycle times, predicted ultimate particle size, and allowed crack behavior to be explored. Process safety evaluations conducted separately indicated that cryocycling does not increase the sensitivity of the propellants examined. The results of this study suggest that cryocycling is a promising technology for the demilitarization of tactical rocket motors.

Current status of free radicals and electronically excited metastable species as high energy propellants

NASA Technical Reports Server (NTRS)

Rosen, G.

1973-01-01

A survey is presented of free radicals and electronically excited metastable species as high energy propellants for rocket engines. Nascent or atomic forms of diatomic gases are considered free radicals as well as the highly reactive diatomic triatomic molecules that posess unpaired electrons. Manufacturing and storage problems are described, and a review of current experimental work related to the manufacture of atomic hydrogen propellants is presented.