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.

The prediction of nozzle performance and heat transfer in hydrogen/oxygen rocket engines with transpiration cooling, film cooling, and high area ratios

NASA Technical Reports Server (NTRS)

Kacynski, Kenneth J.; Hoffman, Joe D.

1993-01-01

An advanced engineering computational model has been developed to aid in the analysis and design of hydrogen/oxygen chemical rocket engines. The complete multi-species, chemically reacting and diffusing Navier-Stokes equations are modelled, finite difference approach that is tailored to be conservative in an axisymmetric coordinate system for both the inviscid and viscous terms. Demonstration cases are presented for a 1030:1 area ratio nozzle, a 25 lbf film cooled nozzle, and transpiration cooled plug-and-spool rocket engine. The results indicate that the thrust coefficient predictions of the 1030:1 nozzle and the film cooled nozzle are within 0.2 to 0.5 percent, respectively, of experimental measurements when all of the chemical reaction and diffusion terms are considered. Further, the model's predictions agree very well with the heat transfer measurements made in all of the nozzle test cases. The Soret thermal diffusion term is demonstrated to have a significant effect on the predicted mass fraction of hydrogen along the wall of the nozzle in both the laminar flow 1030:1 nozzle and the turbulent plug-and-spool rocket engine analysis cases performed. Further, the Soret term was shown to represent a significant fraction of the diffusion fluxes occurring in the transpiration cooled rocket engine.

Project-based introduction to aerospace engineering course: A model rocket

NASA Astrophysics Data System (ADS)

Jayaram, Sanjay; Boyer, Lawrence; George, John; Ravindra, K.; Mitchell, Kyle

2010-05-01

In this paper, a model rocket project suitable for sophomore aerospace engineering students is described. This project encompasses elements of drag estimation, thrust determination and analysis using digital data acquisition, statistical analysis of data, computer aided drafting, programming, team work and written communication skills. The student built rockets are launched in the university baseball field with the objective of carrying a specific amount of payload so that the rocket achieves a specific altitude before the parachute is deployed. During the course of the project, the students are introduced to real-world engineering practice through written report submission of their designs. Over the years, the project has proven to enhance the learning objectives, yet cost effective and has provided good outcome measures.

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...

Simplified Analysis of Pulse Detonation Rocket Engine Blowdown Gasdynamics and Performance

NASA Technical Reports Server (NTRS)

Morris, C. I.; Rodgers, Stephen L. (Technical Monitor)

2002-01-01

Pulse detonation rocket engines (PDREs) offer potential performance improvements over conventional designs, but represent a challenging modellng task. A simplified model for an idealized, straight-tube, single-shot PDRE blowdown process and thrust determination is described and implemented. In order to form an assessment of the accuracy of the model, the flowfield time history is compared to experimental data from Stanford University. Parametric Studies of the effect of mixture stoichiometry, initial fill temperature, and blowdown pressure ratio on the performance of a PDRE are performed using the model. PDRE performance is also compared with a conventional steady-state rocket engine over a range of pressure ratios using similar gasdynamic assumptions.

Hybrid rocket engine, theoretical model and experiment

NASA Astrophysics Data System (ADS)

Chelaru, Teodor-Viorel; Mingireanu, Florin

2011-06-01

The purpose of this paper is to build a theoretical model for the hybrid rocket engine/motor and to validate it using experimental results. The work approaches the main problems of the hybrid motor: the scalability, the stability/controllability of the operating parameters and the increasing of the solid fuel regression rate. At first, we focus on theoretical models for hybrid rocket motor and compare the results with already available experimental data from various research groups. A primary computation model is presented together with results from a numerical algorithm based on a computational model. We present theoretical predictions for several commercial hybrid rocket motors, having different scales and compare them with experimental measurements of those hybrid rocket motors. Next the paper focuses on tribrid rocket motor concept, which by supplementary liquid fuel injection can improve the thrust controllability. A complementary computation model is also presented to estimate regression rate increase of solid fuel doped with oxidizer. Finally, the stability of the hybrid rocket motor is investigated using Liapunov theory. Stability coefficients obtained are dependent on burning parameters while the stability and command matrixes are identified. The paper presents thoroughly the input data of the model, which ensures the reproducibility of the numerical results by independent researchers.

Damage-mitigating control of a reusable rocket engine for high performance and extended life

NASA Technical Reports Server (NTRS)

Ray, Asok; Dai, Xiaowen

1995-01-01

The goal of damage mitigating control in reusable rocket engines is to achieve high performance with increased durability of mechanical structures such that functional lives of the critical components are increased. The major benefit is an increase in structural durability with no significant loss of performance. This report investigates the feasibility of damage mitigating control of reusable rocket engines. Phenomenological models of creep and thermo-mechanical fatigue damage have been formulated in the state-variable setting such that these models can be combined with the plant model of a reusable rocket engine, such as the Space Shuttle Main Engine (SSME), for synthesizing an optimal control policy. Specifically, a creep damage model of the main thrust chamber wall is analytically derived based on the theories of sandwich beam and viscoplasticity. This model characterizes progressive bulging-out and incremental thinning of the coolant channel ligament leading to its eventual failure by tensile rupture. The objective is to generate a closed form solution of the wall thin-out phenomenon in real time where the ligament geometry is continuously updated to account for the resulting deformation. The results are in agreement with those obtained from the finite element analyses and experimental observation for both Oxygen Free High Conductivity (OFHC) copper and a copper-zerconium-silver alloy called NARloy-Z. Due to its computational efficiency, this damage model is suitable for on-line applications of life prediction and damage mitigating control, and also permits parametric studies for off-line synthesis of damage mitigating control systems. The results are presented to demonstrate the potential of life extension of reusable rocket engines via damage mitigating control. The control system has also been simulated on a testbed to observe how the damage at different critical points can be traded off without any significant loss of engine performance. The research work reported here is built upon concepts derived from the disciplines of Controls, Thermo-fluids, Structures, and Materials. The concept of damage mitigation, as presented in this report, is not restricted to control of rocket engines. It can be applied to any system where structural durability is an important issue.

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.

Transient Mathematical Modeling for Liquid Rocket Engine Systems: Methods, Capabilities, and Experience

NASA Technical Reports Server (NTRS)

Seymour, David C.; Martin, Michael A.; Nguyen, Huy H.; Greene, William D.

2005-01-01

The subject of mathematical modeling of the transient operation of liquid rocket engines is presented in overview form from the perspective of engineers working at the NASA Marshall Space Flight Center. The necessity of creating and utilizing accurate mathematical models as part of liquid rocket engine development process has become well established and is likely to increase in importance in the future. The issues of design considerations for transient operation, development testing, and failure scenario simulation are discussed. An overview of the derivation of the basic governing equations is presented along with a discussion of computational and numerical issues associated with the implementation of these equations in computer codes. Also, work in the field of generating usable fluid property tables is presented along with an overview of efforts to be undertaken in the future to improve the tools use for the mathematical modeling process.

Transient Mathematical Modeling for Liquid Rocket Engine Systems: Methods, Capabilities, and Experience

NASA Technical Reports Server (NTRS)

Martin, Michael A.; Nguyen, Huy H.; Greene, William D.; Seymout, David C.

2003-01-01

The subject of mathematical modeling of the transient operation of liquid rocket engines is presented in overview form from the perspective of engineers working at the NASA Marshall Space Flight Center. The necessity of creating and utilizing accurate mathematical models as part of liquid rocket engine development process has become well established and is likely to increase in importance in the future. The issues of design considerations for transient operation, development testing, and failure scenario simulation are discussed. An overview of the derivation of the basic governing equations is presented along with a discussion of computational and numerical issues associated with the implementation of these equations in computer codes. Also, work in the field of generating usable fluid property tables is presented along with an overview of efforts to be undertaken in the future to improve the tools use for the mathematical modeling process.

Probabilistic failure assessment with application to solid rocket motors

NASA Technical Reports Server (NTRS)

Jan, Darrell L.; Davidson, Barry D.; Moore, Nicholas R.

1990-01-01

A quantitative methodology is being developed for assessment of risk of failure of solid rocket motors. This probabilistic methodology employs best available engineering models and available information in a stochastic framework. The framework accounts for incomplete knowledge of governing parameters, intrinsic variability, and failure model specification error. Earlier case studies have been conducted on several failure modes of the Space Shuttle Main Engine. Work in progress on application of this probabilistic approach to large solid rocket boosters such as the Advanced Solid Rocket Motor for the Space Shuttle is described. Failure due to debonding has been selected as the first case study for large solid rocket motors (SRMs) since it accounts for a significant number of historical SRM failures. Impact of incomplete knowledge of governing parameters and failure model specification errors is expected to be important.

Performance of a RBCC Engine in Rocket-Operation

NASA Astrophysics Data System (ADS)

Tomioka, Sadatake; Kubo, Takahiro; Noboru Sakuranaka; Tani, Koichiro

Combination of a scramjet (supersonic combustion ramjet) flow-pass with embedded rocket engines (the combined system termed as Rocket-based Combined Cycle engine) are expected to be the most effective propulsion system for space launch vehicles. Either SSTO (Single Stage To Orbit) system or TSTO (Two Stage To Orbit) system with separation at high altitude needs final stage acceleration in space, so that the RBCC (Rocket Based Combined Cycle) engine should be operated as rocket engines. Performance of the scramjet combustor as the extension to the rocket nozzle, was experimentally evaluated by injecting inert gas at various pressure through the embedded rocket chamber while the whole sub-scaled model was placed in a low pressure chamber connected to an air-driven ejector system. The results showed that the thrust coefficient was about 1.2, the low value being found to mainly due to the friction force on the scramjet combustor wall, while blocking the scramjet flow pass’s opening to increase nozzle extension thrust surface, was found to have little effects on the thrust performance. The combustor was shortened to reduce the friction loss, however, degree of reduction was limited as friction decreased rapidly with distance from the onset of the scramjet combustor.

Nuclear Thermal Rocket Simulation in NPSS

NASA Technical Reports Server (NTRS)

Belair, Michael L.; Sarmiento, Charles J.; Lavelle, Thomas M.

2013-01-01

Four nuclear thermal rocket (NTR) models have been created in the Numerical Propulsion System Simulation (NPSS) framework. The models are divided into two categories. One set is based upon the ZrC-graphite composite fuel element and tie tube-style reactor developed during the Nuclear Engine for Rocket Vehicle Application (NERVA) project in the late 1960s and early 1970s. The other reactor set is based upon a W-UO2 ceramic-metallic (CERMET) fuel element. Within each category, a small and a large thrust engine are modeled. The small engine models utilize RL-10 turbomachinery performance maps and have a thrust of approximately 33.4 kN (7,500 lbf ). The large engine models utilize scaled RL-60 turbomachinery performance maps and have a thrust of approximately 111.2 kN (25,000 lbf ). Power deposition profiles for each reactor were obtained from a detailed Monte Carlo N-Particle (MCNP5) model of the reactor cores. Performance factors such as thermodynamic state points, thrust, specific impulse, reactor power level, and maximum fuel temperature are analyzed for each engine design.

Nuclear Thermal Rocket Simulation in NPSS

NASA Technical Reports Server (NTRS)

Belair, Michael L.; Sarmiento, Charles J.; Lavelle, Thomas L.

2013-01-01

Four nuclear thermal rocket (NTR) models have been created in the Numerical Propulsion System Simulation (NPSS) framework. The models are divided into two categories. One set is based upon the ZrC-graphite composite fuel element and tie tube-style reactor developed during the Nuclear Engine for Rocket Vehicle Application (NERVA) project in the late 1960s and early 1970s. The other reactor set is based upon a W-UO2 ceramic- metallic (CERMET) fuel element. Within each category, a small and a large thrust engine are modeled. The small engine models utilize RL-10 turbomachinery performance maps and have a thrust of approximately 33.4 kN (7,500 lbf ). The large engine models utilize scaled RL-60 turbomachinery performance maps and have a thrust of approximately 111.2 kN (25,000 lbf ). Power deposition profiles for each reactor were obtained from a detailed Monte Carlo N-Particle (MCNP5) model of the reactor cores. Performance factors such as thermodynamic state points, thrust, specific impulse, reactor power level, and maximum fuel temperature are analyzed for each engine design.

Propulsion Technology Lifecycle Operational Analysis

NASA Technical Reports Server (NTRS)

Robinson, John W.; Rhodes, Russell E.

2010-01-01

The paper presents the results of a focused effort performed by the members of the Space Propulsion Synergy Team (SPST) Functional Requirements Sub-team to develop propulsion data to support Advanced Technology Lifecycle Analysis System (ATLAS). This is a spreadsheet application to analyze the impact of technology decisions at a system-of-systems level. Results are summarized in an Excel workbook we call the Technology Tool Box (TTB). The TTB provides data for technology performance, operations, and programmatic parameters in the form of a library of technical information to support analysis tools and/or models. The lifecycle of technologies can be analyzed from this data and particularly useful for system operations involving long running missions. The propulsion technologies in this paper are listed against Chemical Rocket Engines in a Work Breakdown Structure (WBS) format. The overall effort involved establishing four elements: (1) A general purpose Functional System Breakdown Structure (FSBS). (2) Operational Requirements for Rocket Engines. (3) Technology Metric Values associated with Operating Systems (4) Work Breakdown Structure (WBS) of Chemical Rocket Engines The list of Chemical Rocket Engines identified in the WBS is by no means complete. It is planned to update the TTB with a more complete list of available Chemical Rocket Engines for United States (US) engines and add the Foreign rocket engines to the WBS which are available to NASA and the Aerospace Industry. The Operational Technology Metric Values were derived by the SPST Sub-team in the form of the TTB and establishes a database for users to help evaluate and establish the technology level of each Chemical Rocket Engine in the database. The Technology Metric Values will serve as a guide to help determine which rocket engine to invest technology money in for future development.

Hydrocarbon Fuel Thermal Performance Modeling based on Systematic Measurement and Comprehensive Chromatographic Analysis

DTIC Science & Technology

2016-07-31

fueled liquid rocket engine, enthalpy is removed from the combustion chamber by a regenerative cooling system comprising a series of passages through... rocket engine, enthalpy is removed from the combustion chamber by a regenerative cooling system comprising a series of passages through which fuel flows...the unprecedented correlation of comprehensive two-dimensional gas chromatographic (GC×GC) rocket fuel data with physical and thermochemical

Rocket engine diagnostics using qualitative modeling techniques

NASA Technical Reports Server (NTRS)

Binder, Michael; Maul, William; Meyer, Claudia; Sovie, Amy

1992-01-01

Researchers at NASA Lewis Research Center are presently developing qualitative modeling techniques for automated rocket engine diagnostics. A qualitative model of a turbopump interpropellant seal system has been created. The qualitative model describes the effects of seal failures on the system steady-state behavior. This model is able to diagnose the failure of particular seals in the system based on anomalous temperature and pressure values. The anomalous values input to the qualitative model are generated using numerical simulations. Diagnostic test cases include both single and multiple seal failures.

Rocket engine diagnostics using qualitative modeling techniques

NASA Technical Reports Server (NTRS)

Binder, Michael; Maul, William; Meyer, Claudia; Sovie, Amy

1992-01-01

Researchers at NASA Lewis Research Center are presently developing qualitative modeling techniques for automated rocket engine diagnostics. A qualitative model of a turbopump interpropellant seal system was created. The qualitative model describes the effects of seal failures on the system steady state behavior. This model is able to diagnose the failure of particular seals in the system based on anomalous temperature and pressure values. The anomalous values input to the qualitative model are generated using numerical simulations. Diagnostic test cases include both single and multiple seal failures.

Liquid rocket combustion computer model with distributed energy release. DER computer program documentation and user's guide, volume 1

NASA Technical Reports Server (NTRS)

Combs, L. P.

1974-01-01

A computer program for analyzing rocket engine performance was developed. The program is concerned with the formation, distribution, flow, and combustion of liquid sprays and combustion product gases in conventional rocket combustion chambers. The capabilities of the program to determine the combustion characteristics of the rocket engine are described. Sample data code sheets show the correct sequence and formats for variable values and include notes concerning options to bypass the input of certain data. A seperate list defines the variables and indicates their required dimensions.

Linear quadratic servo control of a reusable rocket engine

NASA Technical Reports Server (NTRS)

Musgrave, Jeffrey L.

1991-01-01

A design method for a servo compensator is developed in the frequency domain using singular values. The method is applied to a reusable rocket engine. An intelligent control system for reusable rocket engines was proposed which includes a diagnostic system, a control system, and an intelligent coordinator which determines engine control strategies based on the identified failure modes. The method provides a means of generating various linear multivariable controllers capable of meeting performance and robustness specifications and accommodating failure modes identified by the diagnostic system. Command following with set point control is necessary for engine operation. A Kalman filter reconstructs the state while loop transfer recovery recovers the required degree of robustness while maintaining satisfactory rejection of sensor noise from the command error. The approach is applied to the design of a controller for a rocket engine satisfying performance constraints in the frequency domain. Simulation results demonstrate the performance of the linear design on a nonlinear engine model over all power levels during mainstage operation.

Modeling Primary Atomization Processes

DTIC Science & Technology

1999-02-01

consumable , catalytic igniter has shown to provide reliable, reproducible ignition in hydrogen peroxide/polyethylene hybrid engines. Currently, a...verified in a hybrid rocket using hydrogen peroxide as oxidizer and polyethylene as fuel. The engine made use of a unique Consumable Catalytic Bed (CCB...interest to the liquid and hybrid rocket engine community. TECHNOLOGY TRANSFER Performer Customer Result Application 1 S. D. Heister Purdue University

Design considerations for a pressure-driven multi-stage rocket

NASA Astrophysics Data System (ADS)

Sauerwein, Steven Craig

2002-01-01

The purpose of this study was to examine the feasibility of using propellant tank pressurization to eliminate the use of high-pressure turbopumps in multi-stage liquid-fueled satellite launchers. Several new technologies were examined to reduce the mass of such a rocket. Composite materials have a greater strength-to-weight ratio than metals and can be used to reduce the weight of rocket propellant tanks and structure. Catalytically combined hydrogen and oxygen can be used to heat pressurization gas, greatly reducing the amount of gas required. Ablatively cooled rocket engines can reduce the complexity and cost of the rocket. Methods were derived to estimate the mass of the various rocket components. These included a method to calculate the amount of gas needed to pressurize a propellant tank by modeling the behavior of the pressurization gas as the liquid propellant flows out of the tank. A way to estimate the mass and size of a ablatively cooled composite cased rocket engine. And a method to model the flight of such a rocket through the atmosphere in conjunction with optimization of the rockets trajectory. The results show that while a liquid propellant rocket using tank pressurization are larger than solid propellant rockets and turbopump driven liquid propellant rockets, they are not impractically large.

Outbrief - Long Life Rocket Engine Panel

NASA Technical Reports Server (NTRS)

Quinn, Jason Eugene

2004-01-01

This white paper is an overview of the JANNAF Long Life Rocket Engine (LLRE) Panel results from the last several years of activity. The LLRE Panel has met over the last several years in order to develop an approach for the development of long life rocket engines. Membership for this panel was drawn from a diverse set of the groups currently working on rocket engines (Le. government labs, both large and small companies and university members). The LLRE Panel was formed in order to determine the best way to enable the design of rocket engine systems that have life capability greater than 500 cycles while meeting or exceeding current performance levels (Specific Impulse and Thrust/Weight) with a 1/1,OOO,OOO likelihood of vehicle loss due to rocket system failure. After several meetings and much independent work the panel reached a consensus opinion that the primary issues preventing LLRE are a lack of: physics based life prediction, combined loads prediction, understanding of material microphysics, cost effective system level testing. and the inclusion of fabrication process effects into physics based models. With the expected level of funding devoted to LLRE development, the panel recommended that fundamental research efforts focused on these five areas be emphasized.

A transient model of the RL10A-3-3A rocket engine

NASA Technical Reports Server (NTRS)

Binder, Michael P.

1995-01-01

RL10A-3-3A rocket engines have served as the main propulsion system for Centaur upper stage vehicles since the early 1980's. This hydrogen/oxygen expander cycle engine continues to play a major role in the American launch industry. The Space Propulsion Technology Division at the NASA Lewis Research Center has created a computer model of the RL10 engine, based on detailed component analyses and available test data. This RL10 engine model can predict the performance of the engine over a wide range of operating conditions. The model may also be used to predict the effects of any proposed design changes and anticipated failure scenarios. In this paper, the results of the component analyses are discussed. Simulation results from the new system model are compared with engine test and flight data, including the start and shut-down transient characteristics.

An Eight-Parameter Function for Simulating Model Rocket Engine Thrust Curves

ERIC Educational Resources Information Center

Dooling, Thomas A.

2007-01-01

The toy model rocket is used extensively as an example of a realistic physical system. Teachers from grade school to the university level use them. Many teachers and students write computer programs to investigate rocket physics since the problem involves nonlinear functions related to air resistance and mass loss. This paper describes a nonlinear…

Space Shuttle Model in the 10- by 10-Foot Supersonic Wind Tunnel

NASA Image and Video Library

1975-07-21

Ken Baskin, an engineer from the Facilities and Engineering Branch at the National Aeronautics and Space Administration’s (NASA) Lewis Research Center checks a complete 2.25-scale model of the shuttle in the 10- by 10-Foot Supersonic Wind Tunnel. Baskin’s space shuttle project began in July 1976 during the run-up to the shuttle’s first lift-off scheduled for 1979. The space shuttle was expected to experience multifaceted heating and pressure distributions during the first and second stages of its launch. Rockwell International engineers needed to understand these issues in order to design proper thermal protection. The 10- by 10 tests evaluated the base heating and pressure. The test’s specific objectives were to measure heat transfer and pressure distributions around the orbiter’s external tank and solid rocket booster afterbody caused by rocket exhaust recirculation and impingement, to measure the heat transfer and pressure distributions due to rocket exhaust-induced flow separation, and determine gas recovery temperatures using gas temperature probes and heated model base components. The shuttle model’s main engines and solid rockets were fired during the tests, then just the main engines in an effort to simulate a launch. The researchers conducted 163 runs in the 10- by 10 during the test program.

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.

Vertical Landing Aerodynamics of Reusable Rocket Vehicle

NASA Astrophysics Data System (ADS)

Nonaka, Satoshi; Nishida, Hiroyuki; Kato, Hiroyuki; Ogawa, Hiroyuki; Inatani, Yoshifumi

The aerodynamic characteristics of a vertical landing rocket are affected by its engine plume in the landing phase. The influences of interaction of the engine plume with the freestream around the vehicle on the aerodynamic characteristics are studied experimentally aiming to realize safe landing of the vertical landing rocket. The aerodynamic forces and surface pressure distributions are measured using a scaled model of a reusable rocket vehicle in low-speed wind tunnels. The flow field around the vehicle model is visualized using the particle image velocimetry (PIV) method. Results show that the aerodynamic characteristics, such as the drag force and pitching moment, are strongly affected by the change in the base pressure distributions and reattachment of a separation flow around the vehicle.

Pulse Detonation Rocket Engine Research at NASA Marshall

NASA Technical Reports Server (NTRS)

Morris, Christopher I.

2003-01-01

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

Mean Flow Augmented Acoustics in Rocket Systems

NASA Technical Reports Server (NTRS)

Fischbach, Sean R.

2015-01-01

Combustion instability in solid rocket motors and liquid engines is a complication that continues to plague designers and engineers. Many rocket systems experience violent fluctuations in pressure, velocity, and temperature originating from the complex interactions between the combustion process and gas dynamics. During sever cases of combustion instability fluctuation amplitudes can reach values equal to or greater than the average chamber pressure. Large amplitude oscillations lead to damaged injectors, loss of rocket performance, damaged payloads, and in some cases breach of case/loss of mission. Historic difficulties in modeling and predicting combustion instability has reduced most rocket systems experiencing instability into a costly fix through testing paradigm or to scrap the system entirely.

Evaluation of innovative rocket engines for single-stage earth-to-orbit vehicles

NASA Astrophysics Data System (ADS)

Manski, Detlef; Martin, James A.

1988-07-01

Computer models of rocket engines and single-stage-to-orbit vehicles that were developed by the authors at DFVLR and NASA have been combined. The resulting code consists of engine mass, performance, trajectory and vehicle sizing models. The engine mass model includes equations for each subsystem and describes their dependences on various propulsion parameters. The engine performance model consists of multidimensional sets of theoretical propulsion properties and a complete thermodynamic analysis of the engine cycle. The vehicle analyses include an optimized trajectory analysis, mass estimation, and vehicle sizing. A vertical-takeoff, horizontal-landing, single-stage, winged, manned, fully reusable vehicle with a payload capability of 13.6 Mg (30,000 lb) to low earth orbit was selected. Hydrogen, methane, propane, and dual-fuel engines were studied with staged-combustion, gas-generator, dual bell, and the dual-expander cycles. Mixture ratio, chamber pressure, nozzle exit pressure liftoff acceleration, and dual fuel propulsive parameters were optimized.

Evaluation of innovative rocket engines for single-stage earth-to-orbit vehicles

NASA Technical Reports Server (NTRS)

Manski, Detlef; Martin, James A.

1988-01-01

Computer models of rocket engines and single-stage-to-orbit vehicles that were developed by the authors at DFVLR and NASA have been combined. The resulting code consists of engine mass, performance, trajectory and vehicle sizing models. The engine mass model includes equations for each subsystem and describes their dependences on various propulsion parameters. The engine performance model consists of multidimensional sets of theoretical propulsion properties and a complete thermodynamic analysis of the engine cycle. The vehicle analyses include an optimized trajectory analysis, mass estimation, and vehicle sizing. A vertical-takeoff, horizontal-landing, single-stage, winged, manned, fully reusable vehicle with a payload capability of 13.6 Mg (30,000 lb) to low earth orbit was selected. Hydrogen, methane, propane, and dual-fuel engines were studied with staged-combustion, gas-generator, dual bell, and the dual-expander cycles. Mixture ratio, chamber pressure, nozzle exit pressure liftoff acceleration, and dual fuel propulsive parameters were optimized.

Rocket Engines Displayed for 1966 Inspection at Lewis Research Center

NASA Image and Video Library

1966-10-21

An array of rocket engines displayed in the Propulsion Systems Laboratory for the 1966 Inspection held at the National Aeronautics and Space Administration (NASA) Lewis Research Center. Lewis engineers had been working on chemical, nuclear, and solid rocket engines throughout the 1960s. The engines on display are from left to right: two scale models of the Aerojet M-1, a Rocketdyne J-2, a Pratt and Whitney RL-10, and a Rocketdyne throttleable engine. Also on display are several ejector plates and nozzles. The Chemical Rocket Division resolved issues such as combustion instability and screech, and improved operation of cooling systems and turbopumps. The 1.5-million pound thrust M-1 engine was the largest hydrogen-fueled rocket engine ever created. It was a joint project between NASA Lewis and Aerojet-General. Although much larger in size, the M-1 used technology developed for the RL-10 and J-2. The M-1 program was cancelled in late 1965 due to budget cuts and the lack of a post-Apollo mission. The October 1966 Inspection was the culmination of almost a year of events held to mark the centers’ 25th anniversary. The three‐day Inspection, Lewis’ first since 1957, drew 2000 business, industry, and government executives and included an employee open house. The visitors witnessed presentations at the major facilities and viewed the Gemini VII spacecraft, a Centaur rocket, and other displays in the hangar. In addition, Lewis’ newest facility, the Zero Gravity Facility, was shown off for the first time.

Research on shock wave characteristics in the isolator of central strut rocket-based combined cycle engine under Ma5.5

NASA Astrophysics Data System (ADS)

Wei, Xianggeng; Xue, Rui; Qin, Fei; Hu, Chunbo; He, Guoqiang

2017-11-01

A numerical calculation of shock wave characteristics in the isolator of central strut rocket-based combined cycle (RBCC) engine fueled by kerosene was carried out in this paper. A 3D numerical model was established by the DES method. The kerosene chemical kinetic model used the 9-component and 12-step simplified mechanism model. Effects of fuel equivalence ratio, inflow total temperature and central strut rocket on-off on shock wave characteristics were studied under Ma5.5. Results demonstrated that with the increase of equivalence ratio, the leading shock wave moves toward upstream, accompanied with higher possibility of the inlet unstart. However, the leading shock wave moves toward downstream as the inflow total temperature rises. After the central strut rocket is closed, the leading shock wave moves toward downstream, which can reduce risks of the inlet unstart. State of the shear layer formed by the strut rocket jet flow and inflow can influence the shock train structure significantly.

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.

Potential Climate and Ozone Impacts From Hybrid Rocket Engine Emissions

NASA Astrophysics Data System (ADS)

Ross, M.

2009-12-01

Hybrid rocket engines that use N2O as an oxidizer and a solid hydrocarbon (such as rubber) as a fuel are relatively new. Little is known about the composition of such hybrid engine emissions. General principles and visual inspection of hybrid plumes suggest significant soot and possibly NO emissions. Understanding hybrid rocket emissions is important because of the possibility that a fleet of hybrid powered suborbital rockets will be flying on the order of 1000 flights per year by 2020. The annual stratospheric emission for these rockets would be about 10 kilotons, equal to present day solid rocket motor (SRM) emissions. We present a preliminary analysis of the magnitude of (1) the radiative forcing from soot emissions and (2) the ozone depletion from soot and NO emissions associated with such a fleet of suborbital hybrid rockets. Because the details of the composition of hybrid emissions are unknown, it is not clear if the ozone depletion caused by these hybrid rockets would be more or less than the ozone depletion from SRMs. We also consider the climate implications associated with the N2O production and use requirements for hybrid rockets. Finally, we identify the most important data collection and modeling needs that are required to reliably assess the complete range of environmental impacts of a fleet of hybrid rockets.

Analytical concepts for health management systems of liquid rocket engines

NASA Technical Reports Server (NTRS)

Williams, Richard; Tulpule, Sharayu; Hawman, Michael

1990-01-01

Substantial improvement in health management systems performance can be realized by implementing advanced analytical methods of processing existing liquid rocket engine sensor data. In this paper, such techniques ranging from time series analysis to multisensor pattern recognition to expert systems to fault isolation models are examined and contrasted. The performance of several of these methods is evaluated using data from test firings of the Space Shuttle main engines.

RL10A-3-3A Rocket Engine Modeling Project

NASA Technical Reports Server (NTRS)

Binder, Michael; Tomsik, Thomas; Veres, Joseph P.

1997-01-01

Two RL10A-3-3A rocket engines comprise the main propulsion system for the Centaur upper stage vehicle. Centaur is used with bod Titan and Atlas launch vehicles, carrying military and civilian payloads from high altitudes into orbit and beyond. The RL10 has delivered highly reliable service for the past 30 years. Recently, however, there have been two in-flight failures which have refocused attention on the RL10. This heightened interest has sparked a desire for an independent RL10 modeling capability within NASA and th Air Force. Pratt & Whitney, which presently has the most detailed model of the RL10, also sees merit in having an independent model which could be used as a cross-check with their own simulations. The Space Propulsion Technology Division (SPTD) at the NASA Lewis Research Center has developed a computer model of the RL10A-3-3A. A project team was formed, consisting of experts in the areas of turbomachinery, combustion, and heat transfer. The overall goal of the project was to provide a model of the entire RL10 rocket engine for government use. In the course of the project, the major engine components have been modeled using a combination of simple correlations and detailed component analysis tools (computer codes). The results of these component analyses were verified with data provided by Pratt & Whitney. Select modeling results and test data curves were then integrated to form the RL10 engine system model The purpose of this report is to introduce the reader to the RL10 rocket engine and to describe the engine system model. The RL10 engine and its application to U.S. launch vehicles are described first, followed by a summary of the SPTD project organization, goals, and accomplishments. Simulated output from the system model are shown in comparison with test and flight data for start transient, steady state, and shut-down transient operations. Detailed descriptions of all component analyses, including those not selected for integration with the system model, are included as appendices.

Modal Survey of ETM-3, A 5-Segment Derivative of the Space Shuttle Solid Rocket Booster

NASA Technical Reports Server (NTRS)

Nielsen, D.; Townsend, J.; Kappus, K.; Driskill, T.; Torres, I.; Parks, R.

2005-01-01

The complex interactions between internal motor generated pressure oscillations and motor structural vibration modes associated with the static test configuration of a Reusable Solid Rocket Motor have potential to generate significant dynamic thrust loads in the 5-segment configuration (Engineering Test Motor 3). Finite element model load predictions for worst-case conditions were generated based on extrapolation of a previously correlated 4-segment motor model. A modal survey was performed on the largest rocket motor to date, Engineering Test Motor #3 (ETM-3), to provide data for finite element model correlation and validation of model generated design loads. The modal survey preparation included pretest analyses to determine an efficient analysis set selection using the Effective Independence Method and test simulations to assure critical test stand component loads did not exceed design limits. Historical Reusable Solid Rocket Motor modal testing, ETM-3 test analysis model development and pre-test loads analyses, as well as test execution, and a comparison of results to pre-test predictions are discussed.

Engine Data Interpretation System (EDIS)

NASA Technical Reports Server (NTRS)

Cost, Thomas L.; Hofmann, Martin O.

1990-01-01

A prototype of an expert system was developed which applies qualitative or model-based reasoning to the task of post-test analysis and diagnosis of data resulting from a rocket engine firing. A combined component-based and process theory approach is adopted as the basis for system modeling. Such an approach provides a framework for explaining both normal and deviant system behavior in terms of individual component functionality. The diagnosis function is applied to digitized sensor time-histories generated during engine firings. The generic system is applicable to any liquid rocket engine but was adapted specifically in this work to the Space Shuttle Main Engine (SSME). The system is applied to idealized data resulting from turbomachinery malfunction in the SSME.

JSC engineers visit area schools for National Engineers Week

NASA Image and Video Library

1996-02-28

Johnson Space Center (JSC) engineers visit Houston area schools for National Engineers Week. Students examine a machine that generates static electricity (4296-7). Students examine model rockets (4298).

Transpiration cooled throat for hydrocarbon rocket engines

NASA Technical Reports Server (NTRS)

May, Lee R.; Burkhardt, Wendel M.

1991-01-01

The objective for the Transpiration Cooled Throat for Hydrocarbon Rocket Engines Program was to characterize the use of hydrocarbon fuels as transpiration coolants for rocket nozzle throats. The hydrocarbon fuels investigated in this program were RP-1 and methane. To adequately characterize the above transpiration coolants, a program was planned which would (1) predict engine system performance and life enhancements due to transpiration cooling of the throat region using analytical models, anchored with available data; (2) a versatile transpiration cooled subscale rocket thrust chamber was designed and fabricated; (3) the subscale thrust chamber was tested over a limited range of conditions, e.g., coolant type, chamber pressure, transpiration cooled length, and coolant flow rate; and (4) detailed data analyses were conducted to determine the relationship between the key performance and life enhancement variables.

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.

Numerical Modeling of Pulse Detonation Rocket Engine Gasdynamics and Performance

NASA Technical Reports Server (NTRS)

Morris, C. I.

2003-01-01

Pulse detonation engines (PDB) have generated considerable research interest in recent years as a chemical propulsion system potentially offering improved performance and reduced complexity compared to conventional gas turbines and rocket engines. The detonative mode of combustion employed by these devices offers a theoretical thermodynamic advantage over the constant-pressure deflagrative combustion mode used in conventional engines. However, the unsteady blowdown process intrinsic to all pulse detonation devices has made realistic estimates of the actual propulsive performance of PDES problematic. The recent review article by Kailasanath highlights some of the progress that has been made in comparing the available experimental measurements with analytical and numerical models.

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.

Large liquid rocket engine transient performance simulation system

NASA Technical Reports Server (NTRS)

Mason, J. R.; Southwick, R. D.

1991-01-01

A simulation system, ROCETS, was designed and developed to allow cost-effective computer predictions of liquid rocket engine transient performance. The system allows a user to generate a simulation of any rocket engine configuration using component modules stored in a library through high-level input commands. The system library currently contains 24 component modules, 57 sub-modules and maps, and 33 system routines and utilities. FORTRAN models from other sources can be operated in the system upon inclusion of interface information on comment cards. Operation of the simulation is simplified for the user by run, execution, and output processors. The simulation system makes available steady-state trim balance, transient operation, and linear partial generation. The system utilizes a modern equation solver for efficient operation of the simulations. Transient integration methods include integral and differential forms for the trapezoidal, first order Gear, and second order Gear corrector equations. A detailed technology test bed engine (TTBE) model was generated to be used as the acceptance test of the simulation system. The general level of model detail was that reflected in the Space Shuttle Main Engine DTM. The model successfully obtained steady-state balance in main stage operation and simulated throttle transients, including engine starts and shutdown. A NASA FORTRAN control model was obtained, ROCETS interface installed in comment cards, and operated with the TTBE model in closed-loop transient mode.

Development of CFD model for augmented core tripropellant rocket engine

NASA Astrophysics Data System (ADS)

Jones, Kenneth M.

1994-10-01

The Space Shuttle era has made major advances in technology and vehicle design to the point that the concept of a single-stage-to-orbit (SSTO) vehicle appears more feasible. NASA presently is conducting studies into the feasibility of certain advanced concept rocket engines that could be utilized in a SSTO vehicle. One such concept is a tripropellant system which burns kerosene and hydrogen initially and at altitude switches to hydrogen. This system will attain a larger mass fraction because LOX-kerosene engines have a greater average propellant density and greater thrust-to-weight ratio. This report describes the investigation to model the tripropellant augmented core engine. The physical aspects of the engine, the CFD code employed, and results of the numerical model for a single modular thruster are discussed.

A demonstration of an intelligent control system for a reusable rocket engine

NASA Technical Reports Server (NTRS)

Musgrave, Jeffrey L.; Paxson, Daniel E.; Litt, Jonathan S.; Merrill, Walter C.

1992-01-01

An Intelligent Control System for reusable rocket engines is under development at NASA Lewis Research Center. The primary objective is to extend the useful life of a reusable rocket propulsion system while minimizing between flight maintenance and maximizing engine life and performance through improved control and monitoring algorithms and additional sensing and actuation. This paper describes current progress towards proof-of-concept of an Intelligent Control System for the Space Shuttle Main Engine. A subset of identifiable and accommodatable engine failure modes is selected for preliminary demonstration. Failure models are developed retaining only first order effects and included in a simplified nonlinear simulation of the rocket engine for analysis under closed loop control. The engine level coordinator acts as an interface between the diagnostic and control systems, and translates thrust and mixture ratio commands dictated by mission requirements, and engine status (health) into engine operational strategies carried out by a multivariable control. Control reconfiguration achieves fault tolerance if the nominal (healthy engine) control cannot. Each of the aforementioned functionalities is discussed in the context of an example to illustrate the operation of the system in the context of a representative failure. A graphical user interface allows the researcher to monitor the Intelligent Control System and engine performance under various failure modes selected for demonstration.

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.

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.

Engine System Loads Analysis Compared to Hot-Fire Data

NASA Technical Reports Server (NTRS)

Frady, Gregory P.; Jennings, John M.; Mims, Katherine; Brunty, Joseph; Christensen, Eric R.; McConnaughey, Paul R. (Technical Monitor)

2002-01-01

Early implementation of structural dynamics finite element analyses for calculation of design loads is considered common design practice for high volume manufacturing industries such as automotive and aeronautical industries. However with the rarity of rocket engine development programs starts, these tools are relatively new to the design of rocket engines. In the NASA MC-1 engine program, the focus was to reduce the cost-to-weight ratio. The techniques for structural dynamics analysis practices, were tailored in this program to meet both production and structural design goals. Perturbation of rocket engine design parameters resulted in a number of MC-1 load cycles necessary to characterize the impact due to mass and stiffness changes. Evolution of loads and load extraction methodologies, parametric considerations and a discussion of load path sensitivities are important during the design and integration of a new engine system. During the final stages of development, it is important to verify the results of an engine system model to determine the validity of the results. During the final stages of the MC-1 program, hot-fire test results were obtained and compared to the structural design loads calculated by the engine system model. These comparisons are presented in this paper.

Analysis of liquid-propellant rocket engines designed by F. A. Tsander

NASA Technical Reports Server (NTRS)

Dushkin, L. S.; Moshkin, Y. K.

1977-01-01

The development of the oxygen-gasoline OR-2 engines and the oxygen-alcohol GIRD-10 rocket engine is described. A result of Tsander's rocket research was an engineering method for propellant calculation of oxygen-propellant rocket engines that determined the basic parameters of the engine and the structural elements.

Rocket Ejector Studies for Application to RBCC Engines: An Integrated Experimental/CFD Approach

NASA Technical Reports Server (NTRS)

Pal, S.; Merkle, C. L.; Anderson, W. E.; Santoro, R. J.

1997-01-01

Recent interest in low cost, reliable access to space has generated increased interest in advanced technology approaches to space transportation systems. A key to the success of such programs lies in the development of advanced propulsion systems capable of achieving the performance and operations goals required for the next generation of space vehicles. One extremely promising approach involves the combination of rocket and air- breathing engines into a rocket-based combined-cycle engine (RBCC). A key element of that engine is the rocket ejector which is utilized in the zero to Mach two operating regime. Studies of RBCC engine concepts are not new and studies dating back thirty years are well documented in the literature. However, studies focused on the rocket ejector mode of the RBCC cycle are lacking. The present investigation utilizes an integrated experimental and computation fluid dynamics (CFD) approach to examine critical rocket ejector performance issues. In particular, the development of a predictive methodology capable of performance prediction is a key objective in order to analyze thermal choking and its control, primary/secondary pressure matching considerations, and effects of nozzle expansion ratio. To achieve this objective, the present study emphasizes obtaining new data using advanced optical diagnostics such as Raman spectroscopy and CFD techniques to investigate mixing in the rocket ejector mode. A new research facility for the study of the rocket ejector mode is described along with the diagnostic approaches to be used. The CFD modeling approach is also described along with preliminary CFD predictions obtained to date.

Development Status of Reusable Rocket Engine

NASA Astrophysics Data System (ADS)

Yoshida, Makoto; Takada, Satoshi; Naruo, Yoshihiro; Niu, Kenichi

A 30-kN rocket engine, a pilot engine, is being developed in Japan. Development of this pilot engine has been initiated in relation to a reusable sounding rocket, which is also being developed in Japan. This rocket takes off vertically, reaches an altitude of 100 km, lands vertically at the launch site, and is launched again within several days. Due to advantage of reusability, successful development of this rocket will mean that observation missions can be carried out more frequently and economically. In order to realize this rocket concept, the engines installed on the rocket should be characterized by reusability, long life, deep throttling and health monitoring, features which have not yet been established in Japanese rocket engines. To solve the engineering factors entitled by those features, a new design methodology, advanced engine simulations and engineering testing are being focused on in the pilot engine development stage. Especially in engineering testing, limit condition data is acquired to facilitate development of new diagnostic techniques, which can be applied by utilizing the mobility of small-size hardware. In this paper, the development status of the pilot engine is described, including fundamental design and engineering tests of the turbopump bearing and seal, turbine rig, injector and combustion chamber, and operation and maintenance concepts for one hundred flights by a reusable rocket are examined.

Overview of rocket engine control

NASA Technical Reports Server (NTRS)

Lorenzo, Carl F.; Musgrave, Jeffrey L.

1991-01-01

The issues of Chemical Rocket Engine Control are broadly covered. The basic feedback information and control variables used in expendable and reusable rocket engines, such as Space Shuttle Main Engine, are discussed. The deficiencies of current approaches are considered and a brief introduction to Intelligent Control Systems for rocket engines (and vehicles) is presented.

Fabry-Perot interferometer development for rocket engine plume spectroscopy

NASA Astrophysics Data System (ADS)

Bickford, R. L.; Madzsar, G.

1990-07-01

This paper describes a new rugged high-resolution Fabry-Perot interferometer (FPI) designed for rocket engine plume spectroscopy, which is capable of detecting spectral signatures of eroding engine components during rocket engine tests and/or flight operations. The FPI system will make it possible to predict and to respond to the incipient rocket engine failures and to indicate the presence of rocket components degradation. The design diagram of the FPI spectrometer is presented.

Fabry-Perot interferometer development for rocket engine plume spectroscopy

NASA Technical Reports Server (NTRS)

Bickford, R. L.; Madzsar, G.

1990-01-01

This paper describes a new rugged high-resolution Fabry-Perot interferometer (FPI) designed for rocket engine plume spectroscopy, which is capable of detecting spectral signatures of eroding engine components during rocket engine tests and/or flight operations. The FPI system will make it possible to predict and to respond to the incipient rocket engine failures and to indicate the presence of rocket components degradation. The design diagram of the FPI spectrometer is presented.

A reliability as an independent variable (RAIV) methodology for optimizing test planning for liquid rocket engines

NASA Astrophysics Data System (ADS)

Strunz, Richard; Herrmann, Jeffrey W.

2011-12-01

The hot fire test strategy for liquid rocket engines has always been a concern of space industry and agency alike because no recognized standard exists. Previous hot fire test plans focused on the verification of performance requirements but did not explicitly include reliability as a dimensioning variable. The stakeholders are, however, concerned about a hot fire test strategy that balances reliability, schedule, and affordability. A multiple criteria test planning model is presented that provides a framework to optimize the hot fire test strategy with respect to stakeholder concerns. The Staged Combustion Rocket Engine Demonstrator, a program of the European Space Agency, is used as example to provide the quantitative answer to the claim that a reduced thrust scale demonstrator is cost beneficial for a subsequent flight engine development. Scalability aspects of major subsystems are considered in the prior information definition inside the Bayesian framework. The model is also applied to assess the impact of an increase of the demonstrated reliability level on schedule and affordability.

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.

Dynamic Loads Generation for Multi-Point Vibration Excitation Problems

NASA Technical Reports Server (NTRS)

Shen, Lawrence

2011-01-01

A random-force method has been developed to predict dynamic loads produced by rocket-engine random vibrations for new rocket-engine designs. The method develops random forces at multiple excitation points based on random vibration environments scaled from accelerometer data obtained during hot-fire tests of existing rocket engines. This random-force method applies random forces to the model and creates expected dynamic response in a manner that simulates the way the operating engine applies self-generated random vibration forces (random pressure acting on an area) with the resulting responses that we measure with accelerometers. This innovation includes the methodology (implementation sequence), the computer code, two methods to generate the random-force vibration spectra, and two methods to reduce some of the inherent conservatism in the dynamic loads. This methodology would be implemented to generate the random-force spectra at excitation nodes without requiring the use of artificial boundary conditions in a finite element model. More accurate random dynamic loads than those predicted by current industry methods can then be generated using the random force spectra. The scaling method used to develop the initial power spectral density (PSD) environments for deriving the random forces for the rocket engine case is based on the Barrett Criteria developed at Marshall Space Flight Center in 1963. This invention approach can be applied in the aerospace, automotive, and other industries to obtain reliable dynamic loads and responses from a finite element model for any structure subject to multipoint random vibration excitations.

Reentry aerodynamic characteristics of a space shuttle solid rocket booster model 449 tested in MSFC 14 by 14 inch TWT (SA26F)

NASA Technical Reports Server (NTRS)

Johnson, J. D.; Braddock, W. F.

1974-01-01

Force tests of a 0.563 percent scale space shuttle solid rocket booster (SRB) model, MSFC Model 449, were conducted at the Marshall Space Flight Center 14 x 14 inch Trisonic Wind Tunnel. There were a total of 134 runs (pitch polars) made. Test Mach numbers were 0.6, 0.9, 1.2, 1.96, 2.74, 3.48, 4.00, 4.45, and 4.96; test angles of attack ranged from minus 10 degrees to 190 degrees; test Reynolds numbers ranged from 4.9 million per foot to 7.1 million per foot; and test roll angles were 0, 45, 90, and 135 degrees. The model was tested with three different engine nozzle/skirts. Two of these engine configurations differed from each other in the magnitude of the volume inside the nozzle and skirt. The third engine configuration had part of the nozzle removed. The model was tested with an electrical tunnel in combination with separation rockets of two different heights.

A hydrogen-oxygen rocket engine coolant passage design program (RECOP) for fluid-cooled thrust chambers and nozzles

NASA Technical Reports Server (NTRS)

Tomsik, Thomas M.

1994-01-01

The design of coolant passages in regeneratively cooled thrust chambers is critical to the operation and safety of a rocket engine system. Designing a coolant passage is a complex thermal and hydraulic problem requiring an accurate understanding of the heat transfer between the combustion gas and the coolant. Every major rocket engine company has invested in the development of thrust chamber computer design and analysis tools; two examples are Rocketdyne's REGEN code and Aerojet's ELES program. In an effort to augment current design capabilities for government and industry, the NASA Lewis Research Center is developing a computer model to design coolant passages for advanced regeneratively cooled thrust chambers. The RECOP code incorporates state-of-the-art correlations, numerical techniques and design methods, certainly minimum requirements for generating optimum designs of future space chemical engines. A preliminary version of the RECOP model was recently completed and code validation work is in progress. This paper introduces major features of RECOP and compares the analysis to design points for the first test case engine; the Pratt & Whitney RL10A-3-3A thrust chamber.

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.

Robot Rocket Rally

NASA Image and Video Library

2014-03-14

CAPE CANAVERAL, Fla. – Two young visitors get an up-close look at an engineering model of Robonaut 2, complete with a set of legs, during the Robot Rocket Rally. The three-day event at Florida's Kennedy Space Center Visitor Complex is highlighted by exhibits, games and demonstrations of a variety of robots, with exhibitors ranging from school robotics clubs to veteran NASA scientists and engineers. Photo credit: NASA/Kim Shiflett

Astro STARS Camp

NASA Image and Video Library

2012-06-15

Summer is a time of educational activity at Stennis Space Center. In June 2012, 25 young people age 13-15 attended the annual Astro STARS (Spaceflight, Technology, Astronomy and Robotics at Stennis) camp at the rocket engine test facility. During the five-day camp, participants engaged in hands-on experiences in a variety of areas, including engineering and robotics. On the final day, campers launched model rockets they had assembled.

Pegasus delivers SLS engine section

NASA Image and Video Library

2017-03-03

NASA engineers install test hardware for the agency's new heavy lift rocket, the Space Launch System, into a newly constructed 50-foot structural test stand at NASA's Marshall Space Flight Center. In the stand, hydraulic cylinders will be electronically controlled to push, pull, twist and bend the test article with millions of pounds of force. Engineers will record and analyze over 3,000 channels of data for each test case to verify the capabilities of the engine section and validate that the design and analysis models accurately predict the amount of loads the core stage can withstand during launch and ascent. The engine section, recently delivered via NASA's barge Pegasus from NASA's Michoud Assembly Facility, is the first of four core stage structural test articles scheduled to be delivered to Marshall for testing. The engine section, located at the bottom of SLS's massive core stage, will house the rocket's four RS-25 engines and be an attachment point for the two solid rocket boosters.

Pegasus delivers SLS engine section

NASA Image and Video Library

2017-05-18

NASA engineers install test hardware for the agency's new heavy lift rocket, the Space Launch System, into a newly constructed 50-foot structural test stand at NASA's Marshall Space Flight Center. In the stand, hydraulic cylinders will be electronically controlled to push, pull, twist and bend the test article with millions of pounds of force. Engineers will record and analyze over 3,000 channels of data for each test case to verify the capabilities of the engine section and validate that the design and analysis models accurately predict the amount of loads the core stage can withstand during launch and ascent. The engine section, recently delivered via NASA's barge Pegasus from NASA's Michoud Assembly Facility, is the first of four core stage structural test articles scheduled to be delivered to Marshall for testing. The engine section, located at the bottom of SLS's massive core stage, will house the rocket's four RS-25 engines and be an attachment point for the two solid rocket boosters.

Solid rocket booster performance evaluation model. Volume 1: Engineering description

NASA Technical Reports Server (NTRS)

1974-01-01

The space shuttle solid rocket booster performance evaluation model (SRB-II) is made up of analytical and functional simulation techniques linked together so that a single pass through the model will predict the performance of the propulsion elements of a space shuttle solid rocket booster. The available options allow the user to predict static test performance, predict nominal and off nominal flight performance, and reconstruct actual flight and static test performance. Options selected by the user are dependent on the data available. These can include data derived from theoretical analysis, small scale motor test data, large motor test data and motor configuration data. The user has several options for output format that include print, cards, tape and plots. Output includes all major performance parameters (Isp, thrust, flowrate, mass accounting and operating pressures) as a function of time as well as calculated single point performance data. The engineering description of SRB-II discusses the engineering and programming fundamentals used, the function of each module, and the limitations of each module.

RTE: A computer code for Rocket Thermal Evaluation

NASA Technical Reports Server (NTRS)

Naraghi, Mohammad H. N.

1995-01-01

The numerical model for a rocket thermal analysis code (RTE) is discussed. RTE is a comprehensive thermal analysis code for thermal analysis of regeneratively cooled rocket engines. The input to the code consists of the composition of fuel/oxidant mixture and flow rates, chamber pressure, coolant temperature and pressure. dimensions of the engine, materials and the number of nodes in different parts of the engine. The code allows for temperature variation in axial, radial and circumferential directions. By implementing an iterative scheme, it provides nodal temperature distribution, rates of heat transfer, hot gas and coolant thermal and transport properties. The fuel/oxidant mixture ratio can be varied along the thrust chamber. This feature allows the user to incorporate a non-equilibrium model or an energy release model for the hot-gas-side. The user has the option of bypassing the hot-gas-side calculations and directly inputting the gas-side fluxes. This feature is used to link RTE to a boundary layer module for the hot-gas-side heat flux calculations.

Efficiency of the rocket engines with a supersonic afterburner

NASA Astrophysics Data System (ADS)

Sergienko, A. A.

1992-08-01

The paper is concerned with the problem of regenerative cooling of the liquid-propellant rocket engine combustion chamber at high pressures of the working fluid. It is shown that high combustion product pressures can be achieved in the liquid-propellant rocket engine with a supersonic afterburner than in a liquid-propellant rocket engine with a conventional subsonic combustion chamber for the same allowable heat flux density. However, the liquid-propellant rocket engine with a supersonic afterburner becomes more economical than the conventional engine only at generator gas temperatures of 1700 K and higher.

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.

Zone radiometer measurements on a model rocket exhaust plume

NASA Technical Reports Server (NTRS)

1972-01-01

Radiometer for analytical prediction of rocket plume-to-booster thermal radiation and convective heating is described. Applications for engine combustion analysis, incineration, and pollution control by high temperature processing are discussed. Illustrations of equipment are included.

A minimum cost tolerance allocation method for rocket engines and robust rocket engine design

NASA Technical Reports Server (NTRS)

Gerth, Richard J.

1993-01-01

Rocket engine design follows three phases: systems design, parameter design, and tolerance design. Systems design and parameter design are most effectively conducted in a concurrent engineering (CE) environment that utilize methods such as Quality Function Deployment and Taguchi methods. However, tolerance allocation remains an art driven by experience, handbooks, and rules of thumb. It was desirable to develop and optimization approach to tolerancing. The case study engine was the STME gas generator cycle. The design of the major components had been completed and the functional relationship between the component tolerances and system performance had been computed using the Generic Power Balance model. The system performance nominals (thrust, MR, and Isp) and tolerances were already specified, as were an initial set of component tolerances. However, the question was whether there existed an optimal combination of tolerances that would result in the minimum cost without any degradation in system performance.

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.

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.

Comparison of Engine Cycle Codes for Rocket-Based Combined Cycle Engines

NASA Technical Reports Server (NTRS)

Waltrup, Paul J.; Auslender, Aaron H.; Bradford, John E.; Carreiro, Louis R.; Gettinger, Christopher; Komar, D. R.; McDonald, J.; Snyder, Christopher A.

2002-01-01

This paper summarizes the results from a one day workshop on Rocket-Based Combined Cycle (RBCC) Engine Cycle Codes held in Monterey CA in November of 2000 at the 2000 JANNAF JPM with the authors as primary participants. The objectives of the workshop were to discuss and compare the merits of existing Rocket-Based Combined Cycle (RBCC) engine cycle codes being used by government and industry to predict RBCC engine performance and interpret experimental results. These merits included physical and chemical modeling, accuracy and user friendliness. The ultimate purpose of the workshop was to identify the best codes for analyzing RBCC engines and to document any potential shortcomings, not to demonstrate the merits or deficiencies of any particular engine design. Five cases representative of the operating regimes of typical RBCC engines were used as the basis of these comparisons. These included Mach 0 sea level static and Mach 1.0 and Mach 2.5 Air-Augmented-Rocket (AAR), Mach 4 subsonic combustion ramjet or dual-mode scramjet, and Mach 8 scramjet operating modes. Specification of a generic RBCC engine geometry and concomitant component operating efficiencies, bypass ratios, fuel/oxidizer/air equivalence ratios and flight dynamic pressures were provided. The engine included an air inlet, isolator duct, axial rocket motor/injector, axial wall fuel injectors, diverging combustor, and exit nozzle. Gaseous hydrogen was used as the fuel with the rocket portion of the system using a gaseous H2/O2 propellant system to avoid cryogenic issues. The results of the workshop, even after post-workshop adjudication of differences, were surprising. They showed that the codes predicted essentially the same performance at the Mach 0 and I conditions, but progressively diverged from a common value (for example, for fuel specific impulse, Isp) as the flight Mach number increased, with the largest differences at Mach 8. The example cases and results are compared and discussed in this paper.

Life extending control for rocket engines

NASA Technical Reports Server (NTRS)

Lorenzo, C. F.; Saus, J. R.; Ray, A.; Carpino, M.; Wu, M.-K.

1992-01-01

The concept of life extending control is defined. A brief discussion of current fatigue life prediction methods is given and the need for an alternative life prediction model based on a continuous functional relationship is established. Two approaches to life extending control are considered: (1) the implicit approach which uses cyclic fatigue life prediction as a basis for control design; and (2) the continuous life prediction approach which requires a continuous damage law. Progress on an initial formulation of a continuous (in time) fatigue model is presented. Finally, nonlinear programming is used to develop initial results for life extension for a simplified rocket engine (model).

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.

SRB-3D Solid Rocket Booster performance prediction program. Volume 1: Engineering description/users information manual

NASA Technical Reports Server (NTRS)

Winkler, J. C.

1976-01-01

The modified Solid Rocket Booster Performance Evaluation Model (SRB-3D) was developed as an extension to the internal ballistics module of the SRB-2 performance program. This manual contains the engineering description of SRB-3D which describes the approach used to develop the 3D concept and an explanation of the modifications which were necessary to implement these concepts.

Fluidized-Solid-Fuel Injection Process

NASA Technical Reports Server (NTRS)

Taylor, William

1992-01-01

Report proposes development of rocket engines burning small grains of solid fuel entrained in gas streams. Main technical discussion in report divided into three parts: established fluidization technology; variety of rockets and rocket engines used by nations around the world; and rocket-engine equation. Discusses significance of specific impulse and ratio between initial and final masses of rocket. Concludes by stating three important reasons to proceed with new development: proposed engines safer; fluidized-solid-fuel injection process increases variety of solid-fuel formulations used; and development of fluidized-solid-fuel injection process provides base of engineering knowledge.

Heat transfer to throat tubes in a square-chambered rocket engine at the NASA Lewis Research Center

NASA Technical Reports Server (NTRS)

Nesbitt, James A.; Brindley, William J.

1989-01-01

A gaseous H2/O2 rocket engine was constructed at the NASA-Lewis to provide a high heat flux source representative of the heat flux to the blades in the high pressure fuel turbopump (HPFTP) during startup of the space shuttle main engines. The high heat flux source was required to evaluate the durability of thermal barrier coatings being investigated for use on these blades. The heat transfer, and specifically, the heat flux to tubes located at the throat of the test rocket engine was evaluated and compared to the heat flux to the blades in the HPFTP during engine startup. Gas temperatures, pressures and heat transfer coefficients in the test rocket engine were measured. Near surface metal temperatures below thin thermal barrier coatings were also measured at various angular orientations around the throat tube to indicate the angular dependence of the heat transfer coefficients. A finite difference model for a throat tube was developed and a thermal analysis was performed using the measured gas temperatures and the derived heat transfer coefficients to predict metal temperatures in the tube. Near surface metal temperatures of an uncoated throat tube were measured at the stagnation point and showed good agreement with temperatures predicted by the thermal model. The maximum heat flux to the throat tube was calculated and compared to that predicted for the leading edge of an HPFTP blade. It is shown that the heat flux to an uncooled throat tube is slightly greater than the heat flux to an HPFTP blade during engine startup.

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.

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.

Performance of an Axisymmetric Rocket Based Combined Cycle Engine During Rocket Only Operation Using Linear Regression Analysis

NASA Technical Reports Server (NTRS)

Smith, Timothy D.; Steffen, Christopher J., Jr.; Yungster, Shaye; Keller, Dennis J.

1998-01-01

The all rocket mode of operation is shown to be a critical factor in the overall performance of a rocket based combined cycle (RBCC) vehicle. An axisymmetric RBCC engine was used to determine specific impulse efficiency values based upon both full flow and gas generator configurations. Design of experiments methodology was used to construct a test matrix and multiple linear regression analysis was used to build parametric models. The main parameters investigated in this study were: rocket chamber pressure, rocket exit area ratio, injected secondary flow, mixer-ejector inlet area, mixer-ejector area ratio, and mixer-ejector length-to-inlet diameter ratio. A perfect gas computational fluid dynamics analysis, using both the Spalart-Allmaras and k-omega turbulence models, was performed with the NPARC code to obtain values of vacuum specific impulse. Results from the multiple linear regression analysis showed that for both the full flow and gas generator configurations increasing mixer-ejector area ratio and rocket area ratio increase performance, while increasing mixer-ejector inlet area ratio and mixer-ejector length-to-diameter ratio decrease performance. Increasing injected secondary flow increased performance for the gas generator analysis, but was not statistically significant for the full flow analysis. Chamber pressure was found to be not statistically significant.

Electrostatic propulsion beam divergence effects on spacecraft surfaces, volume 2

NASA Technical Reports Server (NTRS)

Hall, D. F.

1973-01-01

The third phase of a program to develop understanding of and tolerance-level criteria for the deleterious effects of electrostatic rocket exhaust (Cs, Cs(+), Hg, Hg(+)) and materials of rocket construction impinging on typical classes of spacecraft (S/C) surfaces was completed. Models of ion engine effluents and models describing the degradation of S/C surfaces by these effluents are presented. Experimental data from previous phases are summarized and Phase 2 data and analysis are presented in detail. The spacecraft design implications of ion engine contaminants are discussed.

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.

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.

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

Space engine safety system

NASA Technical Reports Server (NTRS)

Maul, William A.; Meyer, Claudia M.

1991-01-01

A rocket engine safety system was designed to initiate control procedures to minimize damage to the engine or vehicle or test stand in the event of an engine failure. The features and the implementation issues associated with rocket engine safety systems are discussed, as well as the specific concerns of safety systems applied to a space-based engine and long duration space missions. Examples of safety system features and architectures are given, based on recent safety monitoring investigations conducted for the Space Shuttle Main Engine and for future liquid rocket engines. Also, the general design and implementation process for rocket engine safety systems is presented.

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.

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

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.

Droplet-turbulence interactions in subcritical and supercritical evaporating sprays

NASA Technical Reports Server (NTRS)

Santavicca, Domenic A.; Coy, Edward; Greenfield, Stuart; Song, Young-Hoon

1991-01-01

The objective of this research is to obtain an improved understanding of droplet turbulence interactions in vaporizing liquid sprays under conditions typical of those encountered in liquid fueled rocket engines. The interaction between liquid droplets and the surrounding turbulent gas flow affects droplet dispersion, droplet collisions, droplet vaporization and gas-phase, fuel-oxidant mixing, and therefore has a significant effect on the engine's combustion characteristics. An example of this is the role which droplet-turbulence interactions are believed to play in combustion instabilities. Despite their importance, droplet-turbulence interactions and their effect on liquid fueled rocket engine performance are not well understood. This is particularly true under supercritical conditions, where many conventional concepts, such as surface tension, no longer apply. Our limited understanding of droplet-turbulence interactions, under both subcritical conditions, represents a major limitation in our ability to design improved liquid previously unavailable information and valuable new insights which will directly impact the design of future liquid fueled rocket engines, as well as, allow for the development of significantly improved spray combustion models, making such models useful design tools.

Application of High Speed Digital Image Correlation in Rocket Engine Hot Fire Testing

NASA Technical Reports Server (NTRS)

Gradl, Paul R.; Schmidt, Tim

2016-01-01

Hot fire testing of rocket engine components and rocket engine systems is a critical aspect of the development process to understand performance, reliability and system interactions. Ground testing provides the opportunity for highly instrumented development testing to validate analytical model predictions and determine necessary design changes and process improvements. To properly obtain discrete measurements for model validation, instrumentation must survive in the highly dynamic and extreme temperature application of hot fire testing. Digital Image Correlation has been investigated and being evaluated as a technique to augment traditional instrumentation during component and engine testing providing further data for additional performance improvements and cost savings. The feasibility of digital image correlation techniques were demonstrated in subscale and full scale hotfire testing. This incorporated a pair of high speed cameras to measure three-dimensional, real-time displacements and strains installed and operated under the extreme environments present on the test stand. The development process, setup and calibrations, data collection, hotfire test data collection and post-test analysis and results are presented in this paper.

The technique for Simulation of Transient Combustion Processes in the Rocket Engine Operating with Gaseous Fuel “Hydrogen and Oxygen”

NASA Astrophysics Data System (ADS)

Zubanov, V. M.; Stepanov, D. V.; Shabliy, L. S.

2017-01-01

The article describes the method for simulation of transient combustion processes in the rocket engine. The engine operates on gaseous propellant: oxygen and hydrogen. Combustion simulation was performed using the ANSYS CFX software. Three reaction mechanisms for the stationary mode were considered and described in detail. Reactions mechanisms have been taken from several sources and verified. The method for converting ozone properties from the Shomate equation to the NASA-polynomial format was described in detail. The way for obtaining quick CFD-results with intermediate combustion components using an EDM model was found. Modeling difficulties with combustion model Finite Rate Chemistry, associated with a large scatter of reference data were identified and described. The way to generate the Flamelet library with CFX-RIF is described. Formulated adequate reaction mechanisms verified at a steady state have also been tested for transient simulation. The Flamelet combustion model was recognized as adequate for the transient mode. Integral parameters variation relates to the values obtained during stationary simulation. A cyclic irregularity of the temperature field, caused by precession of the vortex core, was detected in the chamber with the proposed simulation technique. Investigations of unsteady processes of rocket engines including the processes of ignition were proposed as the area for application of the described simulation technique.

Dual Expander Cycle Rocket Engine with an Intermediate, Closed-cycle Heat Exchanger

NASA Technical Reports Server (NTRS)

Greene, William D. (Inventor)

2008-01-01

A dual expander cycle (DEC) rocket engine with an intermediate closed-cycle heat exchanger is provided. A conventional DEC rocket engine has a closed-cycle heat exchanger thermally coupled thereto. The heat exchanger utilizes heat extracted from the engine's fuel circuit to drive the engine's oxidizer turbomachinery.

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.

Technician Works on a Shuttle Model in the 10- by 10-Foot Supersonic Wind Tunnel

NASA Image and Video Library

1977-02-21

A technician prepares a 2.25 percent scale model of the space shuttle for a base heat study in the 10- by 10-Foot Supersonic Wind Tunnel at the National Aeronautics and Space Administration (NASA) Lewis Research Center. This space shuttle project, begun here in July 1976, was aimed at evaluating base heating and pressure prior to the Shuttle’s first lift-off scheduled for 1979. The space shuttle was expected to experience multifaceted heating and pressure distributions during the first and second stages of its launch. Engineers needed to understand these issues in order to design proper thermal protection. The test’s specific objectives were to measure the heat transfer and pressure distributions around the orbiter’s external tank and solid rocket afterbody caused by rocket exhaust recirculation and impingement, to measure the heat transfer and pressure distributions caused by rocket exhaust-induced separation, and determine gas recovery temperatures using gas temperature probes and heated base components. The shuttle model’s main engines and solid rockets were first fired and then just the main engines to simulate a launch during the testing. Lewis researchers conducted 163 runs in the 10- by 10 during the test program.

Reusable rocket engine optical condition monitoring

NASA Technical Reports Server (NTRS)

Wyett, L.; Maram, J.; Barkhoudarian, S.; Reinert, J.

1987-01-01

Plume emission spectrometry and optical leak detection are described as two new applications of optical techniques to reusable rocket engine condition monitoring. Plume spectrometry has been used with laboratory flames and reusable rocket engines to characterize both the nominal combustion spectra and anomalous spectra of contaminants burning in these plumes. Holographic interferometry has been used to identify leaks and quantify leak rates from reusable rocket engine joints and welds.

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.

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.

Replacement of chemical rocket launchers by beamed energy propulsion.

PubMed

Fukunari, Masafumi; Arnault, Anthony; Yamaguchi, Toshikazu; Komurasaki, Kimiya

2014-11-01

Microwave Rocket is a beamed energy propulsion system that is expected to reach space at drastically lower cost. This cost reduction is estimated by replacing the first-stage engine and solid rocket boosters of the Japanese H-IIB rocket with Microwave Rocket, using a recently developed thrust model in which thrust is generated through repetitively pulsed microwave detonation with a reed-valve air-breathing system. Results show that Microwave Rocket trajectory, in terms of velocity versus altitude, can be designed similarly to the current H-IIB first stage trajectory. Moreover, the payload ratio can be increased by 450%, resulting in launch-cost reduction of 74%.

Heat transfer in rocket engine combustion chambers and regeneratively cooled nozzles

NASA Technical Reports Server (NTRS)

1993-01-01

A conjugate heat transfer computational fluid dynamics (CFD) model to describe regenerative cooling in the main combustion chamber and nozzle and in the injector faceplate region for a launch vehicle class liquid rocket engine was developed. An injector model for sprays which treats the fluid as a variable density, single-phase media was formulated, incorporated into a version of the FDNS code, and used to simulate the injector flow typical of that in the Space Shuttle Main Engine (SSME). Various chamber related heat transfer analyses were made to verify the predictive capability of the conjugate heat transfer analysis provided by the FDNS code. The density based version of the FDNS code with the real fluid property models developed was successful in predicting the streamtube combustion of individual injector elements.

Experimental research and design planning in the field of liquid-propellant rocket engines conducted between 1934 - 1944 by the followers of F. A. Tsander

NASA Technical Reports Server (NTRS)

Dushkin, L. S.

1977-01-01

The development of the following Liquid-Propellant Rocket Engines (LPRE) is reviewed: (1) an alcohol-oxygen single-firing LPRE for use in wingless and winged rockets, (2) a similar multifiring LPRE for use in rocket gliders, (3) a combined solid-liquid propellant rocket engine, and (4) an aircraft LPRE operating on nitric acid and kerosene.

1. ROCKET ENGINE TEST STAND, LOCATED IN THE NORTHEAST ¼ ...

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

1. ROCKET ENGINE TEST STAND, LOCATED IN THE NORTHEAST ¼ OF THE X-15 ENGINE TEST COMPLEX. Looking northeast. - Edwards Air Force Base, X-15 Engine Test Complex, Rocket Engine & Complete X-15 Vehicle Test Stands, Rogers Dry Lake, east of runway between North Base & South Base, Boron, Kern County, CA

Investigation of Cooling Water Injection into Supersonic Rocket Engine Exhaust

NASA Astrophysics Data System (ADS)

Jones, Hansen; Jeansonne, Christopher; Menon, Shyam

2017-11-01

Water spray cooling of the exhaust plume from a rocket undergoing static testing is critical in preventing thermal wear of the test stand structure, and suppressing the acoustic noise signature. A scaled test facility has been developed that utilizes non-intrusive diagnostic techniques including Focusing Color Schlieren (FCS) and Phase Doppler Particle Anemometry (PDPA) to examine the interaction of a pressure-fed water jet with a supersonic flow of compressed air. FCS is used to visually assess the interaction of the water jet with the strong density gradients in the supersonic air flow. PDPA is used in conjunction to gain statistical information regarding water droplet size and velocity as the jet is broken up. Measurement results, along with numerical simulations and jet penetration models are used to explain the observed phenomena. Following the cold flow testing campaign a scaled hybrid rocket engine will be constructed to continue tests in a combusting flow environment similar to that generated by the rocket engines tested at NASA facilities. LaSPACE.

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

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.

Rocket propulsion elements - An introduction to the engineering of rockets (6th revised and enlarged edition)

NASA Astrophysics Data System (ADS)

Sutton, George P.

The subject of rocket propulsion is treated with emphasis on the basic technology, performance, and design rationale. Attention is given to definitions and fundamentals, nozzle theory and thermodynamic relations, heat transfer, flight performance, chemical rocket propellant performance analysis, and liquid propellant rocket engine fundamentals. The discussion also covers solid propellant rocket fundamentals, hybrid propellant rockets, thrust vector control, selection of rocket propulsion systems, electric propulsion, and rocket testing.

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.

Design and Evaluation of a Turbojet Exhaust Simulator, Utilizing a Solid-Propellant Rocket Motor, for use in Free-Flight Aerodynamic Research Models

NASA Technical Reports Server (NTRS)

deMoraes, Carlos A.; Hagginbothom, William K., Jr.; Falanga, Ralph A.

1954-01-01

A method has been developed for modifying a rocket motor so that its exhaust characteristics simulate those of a turbojet engine. The analysis necessary to the design is presented along with tests from which the designs are evaluated. Simulation was found to be best if the exhaust characteristics to be duplicated were those of a turbojet engine at high altitudes and with the afterburner operative.

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.

Effect of engine shroud configuration on the static aerodynamic characteristics of a 0.00563 scale 142-inch diameter solid rocket booster (SA10F)

NASA Technical Reports Server (NTRS)

Johnson, J. D.; Braddock, W. F.

1974-01-01

A test of a 0.563 percent scale space shuttle Solid Rocket Booster (SRB) model, MSFC Model 449, was conducted in a trisonic wind tunnel. Test Mach numbers were 0.4, 0.6, 0.9, 1.2, 1.96, 3.48, 4.0, 4.45, and 4.96. Test angles-of-attack ranged from minus 10 degrees to 190 degrees. Test Reynolds numbers ranged from 3.0 million per foot to 8.6 million per foot. Test roll angles were 0, 11.25, 22.5, 45, and 90 degrees. In addition to the static stability evaluation of the primary SRB configuration, five parametric investigations were made: (1) effect of Reynolds number, (2) effect of engine shroud flare angle, (3) effect of engine shroud length, (4) effect of engine shroud strakes, and (5) effect of engine shroud strakes and trust vector control bottles.

Software for Collaborative Engineering of Launch Rockets

NASA Technical Reports Server (NTRS)

Stanley, Thomas Troy

2003-01-01

The Rocket Evaluation and Cost Integration for Propulsion and Engineering software enables collaborative computing with automated exchange of information in the design and analysis of launch rockets and other complex systems. RECIPE can interact with and incorporate a variety of programs, including legacy codes, that model aspects of a system from the perspectives of different technological disciplines (e.g., aerodynamics, structures, propulsion, trajectory, aeroheating, controls, and operations) and that are used by different engineers on different computers running different operating systems. RECIPE consists mainly of (1) ISCRM a file-transfer subprogram that makes it possible for legacy codes executed in their original operating systems on their original computers to exchange data and (2) CONES an easy-to-use filewrapper subprogram that enables the integration of legacy codes. RECIPE provides a tightly integrated conceptual framework that emphasizes connectivity among the programs used by the collaborators, linking these programs in a manner that provides some configuration control while facilitating collaborative engineering tradeoff studies, including design to cost studies. In comparison with prior collaborative-engineering schemes, one based on the use of RECIPE enables fewer engineers to do more in less time.

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.

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.

Comparison of FDNS liquid rocket engine plume computations with SPF/2

NASA Technical Reports Server (NTRS)

Kumar, G. N.; Griffith, D. O., II; Warsi, S. A.; Seaford, C. M.

1993-01-01

Prediction of a plume's shape and structure is essential to the evaluation of base region environments. The JANNAF standard plume flowfield analysis code SPF/2 predicts plumes well, but cannot analyze base regions. Full Navier-Stokes CFD codes can calculate both zones; however, before they can be used, they must be validated. The CFD code FDNS3D (Finite Difference Navier-Stokes Solver) was used to analyze the single plume of a Space Transportation Main Engine (STME) and comparisons were made with SPF/2 computations. Both frozen and finite rate chemistry models were employed as well as two turbulence models in SPF/2. The results indicate that FDNS3D plume computations agree well with SPF/2 predictions for liquid rocket engine plumes.

Nitrous Oxide/Paraffin Hybrid Rocket Engines

NASA Technical Reports Server (NTRS)

Zubrin, Robert; Snyder, Gary

2010-01-01

Nitrous oxide/paraffin (N2OP) hybrid rocket engines have been invented as alternatives to other rocket engines especially those that burn granular, rubbery solid fuels consisting largely of hydroxyl- terminated polybutadiene (HTPB). Originally intended for use in launching spacecraft, these engines would also be suitable for terrestrial use in rocket-assisted takeoff of small airplanes. The main novel features of these engines are (1) the use of reinforced paraffin as the fuel and (2) the use of nitrous oxide as the oxidizer. Hybrid (solid-fuel/fluid-oxidizer) rocket engines offer advantages of safety and simplicity over fluid-bipropellant (fluid-fuel/fluid-oxidizer) rocket en - gines, but the thrusts of HTPB-based hybrid rocket engines are limited by the low regression rates of the fuel grains. Paraffin used as a solid fuel has a regression rate about 4 times that of HTPB, but pure paraffin fuel grains soften when heated; hence, paraffin fuel grains can, potentially, slump during firing. In a hybrid engine of the present type, the paraffin is molded into a 3-volume-percent graphite sponge or similar carbon matrix, which supports the paraffin against slumping during firing. In addition, because the carbon matrix material burns along with the paraffin, engine performance is not appreciably degraded by use of the matrix.

Performance potential of gas-core and fusion rockets - A mission applications survey.

NASA Technical Reports Server (NTRS)

Fishbach, L. H.; Willis, E. A., Jr.

1971-01-01

This paper reports an evaluation of the performance potential of five nuclear rocket engines for four mission classes. These engines are: the regeneratively cooled gas-core nuclear rocket; the light bulb gas-core nuclear rocket; the space-radiator cooled gas-core nuclear rocket; the fusion rocket; and an advanced solid-core nuclear rocket which is included for comparison. The missions considered are: earth-to-orbit launch; near-earth space missions; close interplanetary missions; and distant interplanetary missions. For each of these missions, the capabilities of each rocket engine type are compared in terms of payload ratio for the earth launch mission or by the initial vehicle mass in earth orbit for space missions (a measure of initial cost). Other factors which might determine the engine choice are discussed. It is shown that a 60 day manned round trip to Mars is conceivable.-

Multiobjective Optimization of Rocket Engine Pumps Using Evolutionary Algorithm

NASA Technical Reports Server (NTRS)

Oyama, Akira; Liou, Meng-Sing

2001-01-01

A design optimization method for turbopumps of cryogenic rocket engines has been developed. Multiobjective Evolutionary Algorithm (MOEA) is used for multiobjective pump design optimizations. Performances of design candidates are evaluated by using the meanline pump flow modeling method based on the Euler turbine equation coupled with empirical correlations for rotor efficiency. To demonstrate the feasibility of the present approach, a single stage centrifugal pump design and multistage pump design optimizations are presented. In both cases, the present method obtains very reasonable Pareto-optimal solutions that include some designs outperforming the original design in total head while reducing input power by one percent. Detailed observation of the design results also reveals some important design criteria for turbopumps in cryogenic rocket engines. These results demonstrate the feasibility of the EA-based design optimization method in this field.

DataRocket: Interactive Visualisation of Data Structures

NASA Astrophysics Data System (ADS)

Parkes, Steve; Ramsay, Craig

2010-08-01

CodeRocket is a software engineering tool that provides cognitive support to the software engineer for reasoning about a method or procedure and for documenting the resulting code [1]. DataRocket is a software engineering tool designed to support visualisation and reasoning about program data structures. DataRocket is part of the CodeRocket family of software tools developed by Rapid Quality Systems [2] a spin-out company from the Space Technology Centre at the University of Dundee. CodeRocket and DataRocket integrate seamlessly with existing architectural design and coding tools and provide extensive documentation with little or no effort on behalf of the software engineer. Comprehensive, abstract, detailed design documentation is available early on in a project so that it can be used for design reviews with project managers and non expert stakeholders. Code and documentation remain fully synchronised even when changes are implemented in the code without reference to the existing documentation. At the end of a project the press of a button suffices to produce the detailed design document. Existing legacy code can be easily imported into CodeRocket and DataRocket to reverse engineer detailed design documentation making legacy code more manageable and adding substantially to its value. This paper introduces CodeRocket. It then explains the rationale for DataRocket and describes the key features of this new tool. Finally the major benefits of DataRocket for different stakeholders are considered.

A detailed description of the uncertainty analysis for high area ratio rocket nozzle tests at the NASA Lewis Research Center

NASA Technical Reports Server (NTRS)

Davidian, Kenneth J.; Dieck, Ronald H.; Chuang, Isaac

1987-01-01

A preliminary uncertainty analysis was performed for the High Area Ratio Rocket Nozzle test program which took place at the altitude test capsule of the Rocket Engine Test Facility at the NASA Lewis Research Center. Results from the study establish the uncertainty of measured and calculated parameters required for the calculation of rocket engine specific impulse. A generalized description of the uncertainty methodology used is provided. Specific equations and a detailed description of the analysis is presented. Verification of the uncertainty analysis model was performed by comparison with results from the experimental program's data reduction code. Final results include an uncertainty for specific impulse of 1.30 percent. The largest contributors to this uncertainty were calibration errors from the test capsule pressure and thrust measurement devices.

A detailed description of the uncertainty analysis for High Area Ratio Rocket Nozzle tests at the NASA Lewis Research Center

NASA Technical Reports Server (NTRS)

Davidian, Kenneth J.; Dieck, Ronald H.; Chuang, Isaac

1987-01-01

A preliminary uncertainty analysis has been performed for the High Area Ratio Rocket Nozzle test program which took place at the altitude test capsule of the Rocket Engine Test Facility at the NASA Lewis Research Center. Results from the study establish the uncertainty of measured and calculated parameters required for the calculation of rocket engine specific impulse. A generalized description of the uncertainty methodology used is provided. Specific equations and a detailed description of the analysis are presented. Verification of the uncertainty analysis model was performed by comparison with results from the experimental program's data reduction code. Final results include an uncertainty for specific impulse of 1.30 percent. The largest contributors to this uncertainty were calibration errors from the test capsule pressure and thrust measurement devices.

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.

2. ROCKET ENGINE TEST STAND, SHOWING TANK (BUILDING 1929) AND ...

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

2. ROCKET ENGINE TEST STAND, SHOWING TANK (BUILDING 1929) AND GARAGE (BUILDING 1930) AT LEFT REAR. Looking to west. - Edwards Air Force Base, X-15 Engine Test Complex, Rocket Engine & Complete X-15 Vehicle Test Stands, Rogers Dry Lake, east of runway between North Base & South Base, Boron, Kern County, CA

7. Historic aerial photo of rocket engine test facility complex, ...

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

7. Historic aerial photo of rocket engine test facility complex, June 1962. On file at NASA Plumbrook Research Center, Sandusky, Ohio. NASA GRC photo number C-60674. - Rocket Engine Testing Facility, NASA Glenn Research Center, Cleveland, Cuyahoga County, OH

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.

Space Shuttle Five-Segment Booster (Short Course)

NASA Technical Reports Server (NTRS)

Graves, Stanley R.; Rudolphi, Michael (Technical Monitor)

2002-01-01

NASA is considering upgrading the Space Shuttle by adding a fifth segment (FSB) to the current four-segment solid rocket booster. Course materials cover design and engineering issues related to the Reusable Solid Rocket Motor (RSRM) raised by the addition of a fifth segment to the rocket booster. Topics cover include: four segment vs. five segment booster, abort modes, FSB grain design, erosive burning, enhanced propellant burn rate, FSB erosive burning model development and hardware configuration.

Rocket engine exhaust plume diagnostics and health monitoring/management during ground testing

NASA Technical Reports Server (NTRS)

Chenevert, D. J.; Meeks, G. R.; Woods, E. G.; Huseonica, H. F.

1992-01-01

The current status of a rocket exhaust plume diagnostics program sponsored by NASA is reviewed. The near-term objective of the program is to enhance test operation efficiency and to provide for safe cutoff of rocket engines prior to incipient failure, thereby avoiding the destruction of the engine and the test complex and preventing delays in the national space program. NASA programs that will benefit from the nonintrusive remote sensed rocket plume diagnostics and related vehicle health management and nonintrusive measurement program are Space Shuttle Main Engine, National Launch System, National Aero-Space Plane, Space Exploration Initiative, Advanced Solid Rocket Motor, and Space Station Freedom. The role of emission spectrometry and other types of remote sensing in rocket plume diagnostics is discussed.

12. Historic plot plan and drawings index for rocket engine ...

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

12. Historic plot plan and drawings index for rocket engine test facility, June 28, 1956. NASA GRC drawing number CE-101810. On file at NASA Glenn Research Center. - Rocket Engine Testing Facility, NASA Glenn Research Center, Cleveland, Cuyahoga County, OH

9. Historic aerial photo of rocket engine test facility complex, ...

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

9. Historic aerial photo of rocket engine test facility complex, June 11, 1965. On file at NASA Plumbrook Research Center, Sandusky, Ohio. NASA GRC photo number C-65-1270. - Rocket Engine Testing Facility, NASA Glenn Research Center, Cleveland, Cuyahoga County, OH

10. Historic photo of rendering of rocket engine test facility ...

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

10. Historic photo of rendering of rocket engine test facility complex, April 28, 1964. On file at NASA Plumbrook Research Center, Sandusky, Ohio. NASA GRC photo number C-69472. - Rocket Engine Testing Facility, NASA Glenn Research Center, Cleveland, Cuyahoga County, OH

8. Historic aerial photo of rocket engine test facility complex, ...

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

8. Historic aerial photo of rocket engine test facility complex, June 11, 1965. On file at NASA Plumbrook Research Center, Sandusky, Ohio. NASA GRC photo number C-65-1271. - Rocket Engine Testing Facility, NASA Glenn Research Center, Cleveland, Cuyahoga County, OH

Hydrocarbon-Fueled Rocket Engine Plume Diagnostics: Analytical Developments and Experimental Results

NASA Technical Reports Server (NTRS)

Tejwani, Gopal D.; McVay, Gregory P.; Langford, Lester A.; St. Cyr, William W.

2006-01-01

A viewgraph presentation describing experimental results and analytical developments about plume diagnostics for hydrocarbon-fueled rocket engines is shown. The topics include: 1) SSC Plume Diagnostics Background; 2) Engine Health Monitoring Approach; 3) Rocket Plume Spectroscopy Simulation Code; 4) Spectral Simulation for 10 Atomic Species and for 11 Diatomic Molecular Electronic Bands; 5) "Best" Lines for Plume Diagnostics for Hydrocarbon-Fueled Rocket Engines; 6) Experimental Set Up for the Methane Thruster Test Program and Experimental Results; and 7) Summary and Recommendations.

NASA Tests RS-25 Flight Engine for Space Launch System

NASA Image and Video Library

2017-10-19

Engineers at NASA’s Stennis Space Center in Mississippi on Oct. 19 completed a hot-fire test of RS-25 rocket engine E2063, a flight engine for NASA’s new Space Launch System (SLS) rocket. Engine E2063 is scheduled to help power SLS on its Exploration Mission-2 (EM-2), the first flight of the new rocket to carry humans.

Significant Climate Changes Caused by Soot Emitted From Rockets in the Stratosphere

NASA Astrophysics Data System (ADS)

Mills, M. J.; Ross, M.; Toohey, D. W.

2010-12-01

A new type of hydrocarbon rocket engine with a larger soot emission index than current kerosene rockets is expected to power a fleet of suborbital rockets for commercial and scientific purposes in coming decades. At projected launch rates, emissions from these rockets will create a persistent soot layer in the northern middle stratosphere that would disproportionally affect the Earth’s atmosphere and cryosphere. A global climate model predicts that thermal forcing in the rocket soot layer will cause significant changes in the global atmospheric circulation and distributions of ozone and temperature. Tropical ozone columns decline as much as 1%, while polar ozone columns increase by up to 6%. Polar surface temperatures rise one Kelvin regionally and polar summer sea ice fractions shrink between 5 - 15%. After 20 years of suborbital rocket fleet operation, globally averaged radiative forcing (RF) from rocket soot exceeds the RF from rocket CO_{2} by six orders of magnitude, but remains small, comparable to the global RF from aviation. The response of the climate system is surprising given the small forcing, and should be investigated further with different climate models.

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

Air-Breathing Rocket Engine Test

NASA Technical Reports Server (NTRS)

2000-01-01

This photograph depicts an air-breathing rocket engine that completed an hour or 3,600 seconds of testing at the General Applied Sciences Laboratory in Ronkonkoma, New York. Referred to as ARGO by its design team, the engine is named after the mythological Greek ship that bore Jason and the Argonauts on their epic voyage of discovery. Air-breathing engines, known as rocket based, combined-cycle engines, get their initial take-off power from specially designed rockets, called air-augmented rockets, that boost performance about 15 percent over conventional rockets. When the vehicle's velocity reaches twice the speed of sound, the rockets are turned off and the engine relies totally on oxygen in the atmosphere to burn hydrogen fuel, as opposed to a rocket that must carry its own oxygen, thus reducing weight and flight costs. Once the vehicle has accelerated to about 10 times the speed of sound, the engine converts to a conventional rocket-powered system to propel the craft into orbit or sustain it to suborbital flight speed. NASA's Advanced SpaceTransportation Program at Marshall Space Flight Center, along with several industry partners and collegiate forces, is developing this technology to make space transportation affordable for everyone from business travelers to tourists. The goal is to reduce launch costs from today's price tag of $10,000 per pound to only hundreds of dollars per pound. NASA's series of hypersonic flight demonstrators currently include three air-breathing vehicles: the X-43A, X-43B and X-43C.

11. Historic photo of cutaway rendering of rocket engine test ...

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

11. Historic photo of cutaway rendering of rocket engine test facility complex, June 11, 1965. On file at NASA Plumbrook Research Center, Sandusky, Ohio. NASA GRC photo number C-74433. - Rocket Engine Testing Facility, NASA Glenn Research Center, Cleveland, Cuyahoga County, OH

Injector element characterization methodology

NASA Technical Reports Server (NTRS)

Cox, George B., Jr.

1988-01-01

Characterization of liquid rocket engine injector elements is an important part of the development process for rocket engine combustion devices. Modern nonintrusive instrumentation for flow velocity and spray droplet size measurement, and automated, computer-controlled test facilities allow rapid, low-cost evaluation of injector element performance and behavior. Application of these methods in rocket engine development, paralleling their use in gas turbine engine development, will reduce rocket engine development cost and risk. The Alternate Turbopump (ATP) Hot Gas Systems (HGS) preburner injector elements were characterized using such methods, and the methodology and some of the results obtained will be shown.

Capitol Day

NASA Image and Video Library

2009-02-19

Stennis Space Center Director Gene Goldman visits with Mississippi Gov. Haley Barbour during NASA Day at the Capitol activities on Feb. 19. During the visit, Goldman presented the governor with a model of the J-2X rocket engine currently in development. Stennis engineers did early component testing for the new engine.

Centrifugal and Axial Pump Design and Off-Design Performance Prediction

NASA Technical Reports Server (NTRS)

Veres, Joseph P.

1995-01-01

A meanline pump-flow modeling method has been developed to provide a fast capability for modeling pumps of cryogenic rocket engines. Based on this method, a meanline pump-flow code PUMPA was written that can predict the performance of pumps at off-design operating conditions, given the loss of the diffusion system at the design point. The design-point rotor efficiency and slip factors are obtained from empirical correlations to rotor-specific speed and geometry. The pump code can model axial, inducer, mixed-flow, and centrifugal pumps and can model multistage pumps in series. The rapid input setup and computer run time for this meanline pump flow code make it an effective analysis and conceptual design tool. The map-generation capabilities of the code provide the information needed for interfacing with a rocket engine system modeling code. The off-design and multistage modeling capabilities of PUMPA permit the user to do parametric design space exploration of candidate pump configurations and to provide head-flow maps for engine system evaluation.

Axisymmetric Numerical Modeling of Pulse Detonation Rocket Engines

NASA Technical Reports Server (NTRS)

Morris, Christopher I.

2005-01-01

Pulse detonation rocket engines (PDREs) have generated research interest in recent years as a chemical propulsion system potentially offering improved performance and reduced complexity compared to conventional rocket engines. The detonative mode of combustion employed by these devices offers a thermodynamic advantage over the constant-pressure deflagrative combustion mode used in conventional rocket engines and gas turbines. However, while this theoretical advantage has spurred considerable interest in building PDRE devices, the unsteady blowdown process intrinsic to the PDRE has made realistic estimates of the actual propulsive performance problematic. The recent review article by Kailasanath highlights some of the progress that has been made in comparing the available experimental measurements with analytical and numerical models. In recent work by the author, a quasi-one-dimensional, finite rate chemistry CFD model was utilized to study the gasdynamics and performance characteristics of PDREs over a range of blowdown pressure ratios from 1-1000. Models of this type are computationally inexpensive, and enable first-order parametric studies of the effect of several nozzle and extension geometries on PDRE performance over a wide range of conditions. However, the quasi-one-dimensional approach is limited in that it cannot properly capture the multidimensional blast wave and flow expansion downstream of the PDRE, nor can it resolve nozzle flow separation if present. Moreover, the previous work was limited to single-pulse calculations. In this paper, an axisymmetric finite rate chemistry model is described and utilized to study these issues in greater detail. Example Mach number contour plots showing the multidimensional blast wave and nozzle exhaust plume are shown. The performance results are compared with the quasi-one-dimensional results from the previous paper. Both Euler and Navier-Stokes solutions are calculated in order to determine the effect of viscous effects in the nozzle flowfield. Additionally, comparisons of the model results to performance data from CalTech, as well as experimental flowfield measurements from Stanford University, are also reported.

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.

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.

Computational Fluid Dynamic Modeling of Rocket Based Combined Cycle Engine Flowfields

NASA Technical Reports Server (NTRS)

Daines, Russell L.; Merkle, Charles L.

1994-01-01

Computational Fluid Dynamic techniques are used to study the flowfield of a fixed geometry Rocket Based Combined Cycle engine operating in rocket ejector mode. Heat addition resulting from the combustion of injected fuel causes the subsonic engine flow to choke and go supersonic in the slightly divergent combustor-mixer section. Reacting flow computations are undertaken to predict the characteristics of solutions where the heat addition is determined by the flowfield. Here, adaptive gridding is used to improve resolution in the shear layers. Results show that the sonic speed is reached in the unheated portions of the flow first, while the heated portions become supersonic later. Comparison with results from another code show reasonable agreement. The coupled solutions show that the character of the combustion-based thermal choking phenomenon can be controlled reasonably well such that there is opportunity to optimize the length and expansion ratio of the combustor-mixer.

6. Historic photo of rocket engine test facility Building 202 ...

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

6. Historic photo of rocket engine test facility Building 202 complex in operation at night, September 12, 1957. On file at NASA Plumbrook Research Center, Sandusky, Ohio. NASA GRC photo number C-45924. - Rocket Engine Testing Facility, NASA Glenn Research Center, Cleveland, Cuyahoga County, OH

13. Historic drawing of rocket engine test facility layout, including ...

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

13. Historic drawing of rocket engine test facility layout, including Buildings 202, 205, 206, and 206A, February 3, 1984. NASA GRC drawing number CF-101539. On file at NASA Glenn Research Center. - Rocket Engine Testing Facility, NASA Glenn Research Center, Cleveland, Cuyahoga County, OH

Prediction of pressure and flow transients in a gaseous bipropellant reaction control rocket engine

NASA Technical Reports Server (NTRS)

Markowsky, J. J.; Mcmanus, H. N., Jr.

1974-01-01

An analytic model is developed to predict pressure and flow transients in a gaseous hydrogen-oxygen reaction control rocket engine feed system. The one-dimensional equations of momentum and continuity are reduced by the method of characteristics from partial derivatives to a set of total derivatives which describe the state properties along the feedline. System components, e.g., valves, manifolds, and injectors are represented by pseudo steady-state relations at discrete junctions in the system. Solutions were effected by a FORTRAN IV program on an IBM 360/65. The results indicate the relative effect of manifold volume, combustion lag time, feedline pressure fluctuations, propellant temperature, and feedline length on the chamber pressure transient. The analytical combustion model is verified by good correlation between predicted and observed chamber pressure transients. The developed model enables a rocket designer to vary the design parameters analytically to obtain stable combustion for a particular mode of operation which is prescribed by mission objectives.

Free-Spinning-Tunnel Investigation to Determine the Effect of Spin-Recovery Rockets and Thrust Simulation on the Recovery Characteristics of a 1/21-Scale Model of the Chance Vought F7U-3 Airplane, TED No. NACA AD 3103

NASA Technical Reports Server (NTRS)

Burk, Sanger H., Jr.; Healy, Frederick M.

1955-01-01

An investigation of a l/21-scale model of the Chance Vought F7U-3 airplane in the co&at-load- condition has been conducted in the Langley 20-foot free-spinning tunnel, The recovery characteristics of the model were determined by use of spin-recovery rockets for the erect and inverted spinning condition. The rockets were so placed as to provide either a yawing or rolling moment about the model center of gravity. Also included in the investigation were tests to determine the effect of simulated engine thrust on the recovery characteristics of the model. On the basis of model tests, recoveries from erect and inverted spins were satisfactory when a yawing moment of 22,200 foot-pounds (full scale) was provided against the spin by rockets attached to the wing tips; the anti-spin yawing moment was applied for approximately 9 seconds, (full scale). Satisfactory recoveries were obtained from erect spins when a rolling moment of 22,200 foot-pounds (full scale) was provided with the spin (rolls right wing down in right spin). Although the inverted spin was satisfactorily terminated when a rolling moment of equal magnitude was provided, a roll rocket was not considered to be an optimum spin-recovery device to effect recoveries from inverted spins for this airplane because of resulting gyrations during spin recovery. Simulation of engine thrust had no apparent effect on the spin recovery characteristics.

RS-25 Rocket Engine Test

NASA Image and Video Library

2017-08-09

The 8.5-minute test conducted at NASA’s Stennis Space Center is part of a series of tests designed to put the upgraded former space shuttle engines through the rigorous temperature and pressure conditions they will experience during a launch. The tests also support the development of a new controller, or “brain,” for the engine, which monitors engine status and communicates between the rocket and the engine, relaying commands to the engine and transmitting data back to the 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.

State Machine Modeling of the Space Launch System Solid Rocket Boosters

NASA Technical Reports Server (NTRS)

Harris, Joshua A.; Patterson-Hine, Ann

2013-01-01

The Space Launch System is a Shuttle-derived heavy-lift vehicle currently in development to serve as NASA's premiere launch vehicle for space exploration. The Space Launch System is a multistage rocket with two Solid Rocket Boosters and multiple payloads, including the Multi-Purpose Crew Vehicle. Planned Space Launch System destinations include near-Earth asteroids, the Moon, Mars, and Lagrange points. The Space Launch System is a complex system with many subsystems, requiring considerable systems engineering and integration. To this end, state machine analysis offers a method to support engineering and operational e orts, identify and avert undesirable or potentially hazardous system states, and evaluate system requirements. Finite State Machines model a system as a finite number of states, with transitions between states controlled by state-based and event-based logic. State machines are a useful tool for understanding complex system behaviors and evaluating "what-if" scenarios. This work contributes to a state machine model of the Space Launch System developed at NASA Ames Research Center. The Space Launch System Solid Rocket Booster avionics and ignition subsystems are modeled using MATLAB/Stateflow software. This model is integrated into a larger model of Space Launch System avionics used for verification and validation of Space Launch System operating procedures and design requirements. This includes testing both nominal and o -nominal system states and command sequences.

Prediction of Acoustic Loads Generated by Propulsion Systems

NASA Technical Reports Server (NTRS)

Perez, Linamaria; Allgood, Daniel C.

2011-01-01

NASA Stennis Space Center is one of the nation's premier facilities for conducting large-scale rocket engine testing. As liquid rocket engines vary in size, so do the acoustic loads that they produce. When these acoustic loads reach very high levels they may cause damages both to humans and to actual structures surrounding the testing area. To prevent these damages, prediction tools are used to estimate the spectral content and levels of the acoustics being generated by the rocket engine plumes and model their propagation through the surrounding atmosphere. Prior to the current work, two different acoustic prediction tools were being implemented at Stennis Space Center, each having their own advantages and disadvantages depending on the application. Therefore, a new prediction tool was created, using NASA SP-8072 handbook as a guide, which would replicate the same prediction methods as the previous codes, but eliminate any of the drawbacks the individual codes had. Aside from replicating the previous modeling capability in a single framework, additional modeling functions were added thereby expanding the current modeling capability. To verify that the new code could reproduce the same predictions as the previous codes, two verification test cases were defined. These verification test cases also served as validation cases as the predicted results were compared to actual test data.

XLR-11 - X-1 rocket engine display

NASA Technical Reports Server (NTRS)

1996-01-01

What started as a hobby for four rocket fanatics went on to break the sound barrier: Lovell Lawrence, Hugh Franklin Pierce, John Shesta, and Jimmy Wyld the four founders of Reaction Motors, Inc. that built the XLR-11 Rocket Engine. The XLR-11 engine is shown on display in the NASA Exchange Gift Shop, NASA Hugh L. Dryden Flight Research Center at Edwards, California. This engine, familiarly known as Black Betsy, a 4-chamber rocket that ignited diluted ethyl alcohol and liquid oxygen into 6000 pounds or more of thrust powered the X-1 series airplanes.

Rocket University at KSC

NASA Technical Reports Server (NTRS)

Sullivan, Steven J.

2014-01-01

"Rocket University" is an exciting new initiative at Kennedy Space Center led by NASA's Engineering and Technology Directorate. This hands-on experience has been established to develop, refine & maintain targeted flight engineering skills to enable the Agency and KSC strategic goals. Through "RocketU", KSC is developing a nimble, rapid flight engineering life cycle systems knowledge base. Ongoing activities in RocketU develop and test new technologies and potential customer systems through small scale vehicles, build and maintain flight experience through balloon and small-scale rocket missions, and enable a revolving fresh perspective of engineers with hands on expertise back into the large scale NASA programs, providing a more experienced multi-disciplined set of systems engineers. This overview will define the Program, highlight aspects of the training curriculum, and identify recent accomplishments and activities.

Improving of Hybrid Rocket Engine on the Basis of Optimizing Design Fuel Grain

NASA Astrophysics Data System (ADS)

Oriekov, K. M.; Ushkin, M. P.

2015-09-01

This article examines the processes intrachamber in hybrid rocket engine (HRE) and the comparative assessment of the use of solid rocket motors (SRM) and HRE for meteorological rockets with a mass of payload of the 364 kg. Results of the research showed the possibility of a significant increase in the ballistic effectiveness of meteorological rocket.

Quasi-One-Dimensional Modeling of Pulse Detonation Rocket Engines

NASA Technical Reports Server (NTRS)

Morris, Christopher I.

2002-01-01

Pulsed detonation rocket engines (PDREs) have generated considerable research interest in recent years as a chemical propulsion system potentially offering improved performance and reduced complexity compared to conventional rocket engines. The detonative mode of combustion employed by these devices offers a thermodynamic advantage over the constant-pressure deflagrative combustion mode used in conventional rocket engines and gas turbines. However, while this theoretical advantage has spurred a great deal of interest in building PDRE devices, the unsteady blowdown process intrinsic to the PDRE has made realistic estimates of the actual propulsive performance problematic. The recent review article by Kailasanath highlights some of the difficulties in comparing the available experimental measurements with numerical models. In a previous paper by the author, parametric studies of the performance of a single, straight-tube PDRE were reported. A 1-D, unsteady method of characteristics code, employing a constant-gamma assumption behind the detonation front, was developed for that study. Models of this type are computationally inexpensive, and are particularly useful for parametric performance comparisons. For example, a plot showing the specific impulse of various PDRE and steady-state rocket engine (SSRE) configurations as a function of blowdown pressure ratio. The performance curves clearly indicate that a straight-tube PDRE is superior in specific impulse to a SSRE with a sonic nozzle over the entire range of pressure ratios. Note, however, that a straight-tube PDRE in general does not compare favorably to a SSRE fitted with an optimized de Laval supersonic nozzle, particularly at the high pressure ratios typical for boost or in-space rocket applications. However, the calculations also show that if a dynamically optimized, supersonic de Laval nozzle could be could be fitted to a PDRE, then the specific impulse of the device would exceed that of a comparable SSRE. While such a nozzle is a considerable idealization, it is clear that nozzle design and optimization will play a critical role in whether the performance potential of PDREs can be effectively realized in practice. In order to study PDRE nozzle issues with greater accuracy, a quasi-one-dimensional, finite-rate chemistry CFD code has been developed by the author. Comparisons of the code with both the previous MOC model and experimental data from Stanford University are reported. The effect of constant-gamma and finite-rate chemistry assumptions on the flowfield and performance is examined. Parametric studies of the effect of nozzle throat size and expansion ratio, at various blowdown pressure ratios, are reported.

Software for Estimating Costs of Testing Rocket Engines

NASA Technical Reports Server (NTRS)

Hines, Merlon M.

2004-01-01

A high-level parametric mathematical model for estimating the costs of testing rocket engines and components at Stennis Space Center has been implemented as a Microsoft Excel program that generates multiple spreadsheets. The model and the program are both denoted, simply, the Cost Estimating Model (CEM). The inputs to the CEM are the parameters that describe particular tests, including test types (component or engine test), numbers and duration of tests, thrust levels, and other parameters. The CEM estimates anticipated total project costs for a specific test. Estimates are broken down into testing categories based on a work-breakdown structure and a cost-element structure. A notable historical assumption incorporated into the CEM is that total labor times depend mainly on thrust levels. As a result of a recent modification of the CEM to increase the accuracy of predicted labor times, the dependence of labor time on thrust level is now embodied in third- and fourth-order polynomials.

Software for Estimating Costs of Testing Rocket Engines

NASA Technical Reports Server (NTRS)

Hines, Merion M.

2002-01-01

A high-level parametric mathematical model for estimating the costs of testing rocket engines and components at Stennis Space Center has been implemented as a Microsoft Excel program that generates multiple spreadsheets. The model and the program are both denoted, simply, the Cost Estimating Model (CEM). The inputs to the CEM are the parameters that describe particular tests, including test types (component or engine test), numbers and duration of tests, thrust levels, and other parameters. The CEM estimates anticipated total project costs for a specific test. Estimates are broken down into testing categories based on a work-breakdown structure and a cost-element structure. A notable historical assumption incorporated into the CEM is that total labor times depend mainly on thrust levels. As a result of a recent modification of the CEM to increase the accuracy of predicted labor times, the dependence of labor time on thrust level is now embodied in third- and fourth-order polynomials.

Software for Estimating Costs of Testing Rocket Engines

NASA Technical Reports Server (NTRS)

Hines, Merlon M.

2003-01-01

A high-level parametric mathematical model for estimating the costs of testing rocket engines and components at Stennis Space Center has been implemented as a Microsoft Excel program that generates multiple spreadsheets. The model and the program are both denoted, simply, the Cost Estimating Model (CEM). The inputs to the CEM are the parameters that describe particular tests, including test types (component or engine test), numbers and duration of tests, thrust levels, and other parameters. The CEM estimates anticipated total project costs for a specific test. Estimates are broken down into testing categories based on a work-breakdown structure and a cost-element structure. A notable historical assumption incorporated into the CEM is that total labor times depend mainly on thrust levels. As a result of a recent modification of the CEM to increase the accuracy of predicted labor times, the dependence of labor time on thrust level is now embodied in third- and fourth-order polynomials.

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.

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

NASA Tests 2nd RS-25 Flight Engine for Space Launch System

NASA Image and Video Library

2017-10-19

Engineers at NASA’s Stennis Space Center in Mississippi on Oct. 19 completed a hot-fire test of RS-25 rocket engine E2063, a flight engine for NASA’s new Space Launch System (SLS) rocket. Engine E2063 is scheduled to help power SLS on its Exploration Mission-2 (EM-2), the first flight of the new rocket to carry humans. Flight engine E2059 was tested on March 10, 2016, also for use on the EM-2 flight.

NASA Tests 2nd RS-25 Flight Engine For Space Launch System

NASA Image and Video Library

2017-10-19

Engineers at NASA’s Stennis Space Center in Mississippi on Oct. 19 completed a hot-fire test of RS-25 rocket engine E2063, a flight engine for NASA’s new Space Launch System (SLS) rocket. Engine E2063 is scheduled to help power SLS on its Exploration Mission-2 (EM-2), the first flight of the new rocket to carry humans. Flight engine E2059 was tested on March 10, 2016, also for use on the EM-2 flight.

Video File - NASA Tests 2nd RS-25 Flight Engine for Space Launch System

NASA Image and Video Library

2017-10-19

Engineers at NASA’s Stennis Space Center in Mississippi on Oct. 19 completed a hot-fire test of RS-25 rocket engine E2063, a flight engine for NASA’s new Space Launch System (SLS) rocket. Engine E2063 is scheduled to help power SLS on its Exploration Mission-2 (EM-2), the first flight of the new rocket to carry humans. Flight engine E2059 was tested on March 10, 2016, also for use on the EM-2 flight.

Teaching Engineering Design Through Paper Rockets

ERIC Educational Resources Information Center

Welling, Jonathan; Wright, Geoffrey A.

2018-01-01

The paper rocket activity described in this article effectively teaches the engineering design process (EDP) by engaging students in a problem-based learning activity that encourages iterative design. For example, the first rockets the students build typically only fly between 30 and 100 feet. As students test and evaluate their rocket designs,…

Thrust control system design of ducted rockets

NASA Astrophysics Data System (ADS)

Chang, Juntao; Li, Bin; Bao, Wen; Niu, Wenyu; Yu, Daren

2011-07-01

The investigation of the thrust control system is aroused by the need for propulsion system of ducted rockets. Firstly the dynamic mathematical models of gas flow regulating system, pneumatic servo system and ducted rocket engine were established and analyzed. Then, to conquer the discussed problems of thrust control, the idea of information fusion was proposed to construct a new feedback variable. With this fused feedback variable, the thrust control system was designed. According to the simulation results, the introduction of the new fused feedback variable is valid in eliminating the contradiction between rapid response and stability for the thrust control system of ducted rockets.

Optimal trajectories for an aerospace plane. Part 1: Formulation, results, and analysis

NASA Technical Reports Server (NTRS)

Miele, Angelo; Lee, W. Y.; Wu, G. D.

1990-01-01

The optimization of the trajectories of an aerospace plane is discussed. This is a hypervelocity vehicle capable of achieving orbital speed, while taking off horizontally. The vehicle is propelled by four types of engines: turbojet engines for flight at subsonic speeds/low supersonic speeds; ramjet engines for flight at moderate supersonic speeds/low hypersonic speeds; scramjet engines for flight at hypersonic speeds; and rocket engines for flight at near-orbital speeds. A single-stage-to-orbit (SSTO) configuration is considered, and the transition from low supersonic speeds to orbital speeds is studied under the following assumptions: the turbojet portion of the trajectory has been completed; the aerospace plane is controlled via the angle of attack and the power setting; the aerodynamic model is the generic hypersonic aerodynamics model example (GHAME). Concerning the engine model, three options are considered: (EM1), a ramjet/scramjet combination in which the scramjet specific impulse tends to a nearly-constant value at large Mach numbers; (EM2), a ramjet/scramjet combination in which the scramjet specific impulse decreases monotonically at large Mach numbers; and (EM3), a ramjet/scramjet/rocket combination in which, owing to stagnation temperature limitations, the scramjet operates only at M approx. less than 15; at higher Mach numbers, the scramjet is shut off and the aerospace plane is driven only by the rocket engines. Under the above assumptions, four optimization problems are solved using the sequential gradient-restoration algorithm for optimal control problems: (P1) minimization of the weight of fuel consumed; (P2) minimization of the peak dynamic pressure; (P3) minimization of the peak heating rate; and (P4) minimization of the peak tangential acceleration.

Numerical investigations of hybrid rocket engines

NASA Astrophysics Data System (ADS)

Betelin, V. B.; Kushnirenko, A. G.; Smirnov, N. N.; Nikitin, V. F.; Tyurenkova, V. V.; Stamov, L. I.

2018-03-01

Paper presents the results of numerical studies of hybrid rocket engines operating cycle including unsteady-state transition stage. A mathematical model is developed accounting for the peculiarities of diffusion combustion of fuel in the flow of oxidant, which is composed of oxygen-nitrogen mixture. Three dimensional unsteady-state simulations of chemically reacting gas mixture above thermochemically destructing surface are performed. The results show that the diffusion combustion brings to strongly non-uniform fuel mass regression rate in the flow direction. Diffusive deceleration of chemical reaction brings to the decrease of fuel regression rate in the longitudinal direction.

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.

Development of a 12-Thrust Chamber Kerosene /Oxygen Primary Rocket Sub-System for an Early (1964) Air-Augmented Rocket Ground-Test System

NASA Technical Reports Server (NTRS)

Pryor, D.; Hyde, E. H.; Escher, W. J. D.

1999-01-01

Airbreathing/Rocket combined-cycle, and specifically rocket-based combined- cycle (RBCC), propulsion systems, typically employ an internal engine flow-path installed primary rocket subsystem. To achieve acceptably short mixing lengths in effecting the "air augmentation" process, a large rocket-exhaust/air interfacial mixing surface is needed. This leads, in some engine design concepts, to a "cluster" of small rocket units, suitably arrayed in the flowpath. To support an early (1964) subscale ground-test of a specific RBCC concept, such a 12-rocket cluster was developed by NASA's Marshall Space Flight Center (MSFC). The small primary rockets used in the cluster assembly were modified versions of an existing small kerosene/oxygen water-cooled rocket engine unit routinely tested at MSFC. Following individual thrust-chamber tests and overall subsystem qualification testing, the cluster assembly was installed at the U. S. Air Force's Arnold Engineering Development Center (AEDC) for RBCC systems testing. (The results of the special air-augmented rocket testing are not covered here.) While this project was eventually successfully completed, a number of hardware integration problems were met, leading to catastrophic thrust chamber failures. The principal "lessons learned" in conducting this early primary rocket subsystem experimental effort are documented here as a basic knowledge-base contribution for the benefit of today's RBCC research and development community.

29. Historic view of twentythousandpound rocket test stand with engine ...

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

29. Historic view of twenty-thousand-pound rocket test stand with engine installation in test cell of Building 202, September 1957. On file at NASA Plumbrook Research Center, Sandusky, Ohio. NASA GRC photo number C-45870. - Rocket Engine Testing Facility, GRC Building No. 202, NASA Glenn Research Center, Cleveland, Cuyahoga County, OH

Calculation of Propulsive Nozzle Flowfields in Multidiffusing Chemically Reacting Environments. Ph.D. Thesis - Purdue Univ.

NASA Technical Reports Server (NTRS)

Kacynski, Kenneth John

1994-01-01

An advanced engineering model has been developed to aid in the analysis and design of hydrogen/oxygen chemical rocket engines. The complete multispecies, chemically reacting and multidiffusing Navier-Stokes equations are modelled, including the Soret thermal diffusion and the Dufour energy transfer terms. In addition to the spectrum of multispecies aspects developed, the model developed in this study is also conservative in axisymmetric flow for both inviscid and viscous flow environments and the boundary conditions employ a viscous, chemically reacting, reference plane characteristics method. Demonstration cases are presented for a 1030:1 area ratio nozzle, a 25 lbf film cooled nozzle, and a transpiration cooled plug and spool rocket engine. The results indicate that the thrust coefficient predictions of the 1030:1 and the 25 lbf film cooled nozzle are within 0.2 to 0.5 percent, respectively, of experimental measurements when all of the chemical reaction and diffusion terms are considered. Further, the model's predictions agree very well with the heat transfer measurements made in all of the nozzle test cases. The Soret thermal diffusion term is demonstrated to have a significant effect on the predicted mass fraction of hydrogen along the wall of the nozzle in both the laminar flow 1030:1 nozzle and the turbulent flow plug and spool nozzle analysis cases performed. Further, the Soret term was shown to represent an important fraction of the diffusion fluxes occurring in a transpiration cooled rocket engine.

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.

Mercury Capsule Separation Tests

NASA Image and Video Library

1960-04-01

Mercury capsule separation from Redstone booster in the Altitude Wind Tunnel (AWT): NASA Lewis conducted full-scale separation tests of the posigrade rockets that were fired after the Redstone rockets burned out. The researchers studied the effect of the posigrade rockets firing on the Redstone booster and retrograde package. This film shows the Mercury capsule being mounted to the Redstone missile model in the Altitude Wind Tunnel. The capsule's engines are fired and it horizontally separates from the Atlas. After firing the capsule swings from an overhead crane.

Scale-Up of GRCop: From Laboratory to Rocket Engines

NASA Technical Reports Server (NTRS)

Ellis, David L.

2016-01-01

GRCop is a high temperature, high thermal conductivity copper-based series of alloys designed primarily for use in regeneratively cooled rocket engine liners. It began with laboratory-level production of a few grams of ribbon produced by chill block melt spinning and has grown to commercial-scale production of large-scale rocket engine liners. Along the way, a variety of methods of consolidating and working the alloy were examined, a database of properties was developed and a variety of commercial and government applications were considered. This talk will briefly address the basic material properties used for selection of compositions to scale up, the methods used to go from simple ribbon to rocket engines, the need to develop a suitable database, and the issues related to getting the alloy into a rocket engine or other application.

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.

Study of solid rocket motor for a space shuttle booster. Appendix A: SRM water entry loads

NASA Technical Reports Server (NTRS)

1972-01-01

An analysis of the water entry loads imposed on the reusable solid propellant rocket engine of the space shuttle following parachute descent is presented. The cases discussed are vertical motion, horizontal motion, and motion after penetration. Mathematical models, diagrams, and charts are included to support the theoretical considerations.

Integrated model development for liquid fueled rocket propulsion systems

NASA Technical Reports Server (NTRS)

Santi, L. Michael

1993-01-01

As detailed in the original statement of work, the objective of phase two of this research effort was to develop a general framework for rocket engine performance prediction that integrates physical principles, a rigorous mathematical formalism, component level test data, system level test data, and theory-observation reconciliation. Specific phase two development tasks are defined.

KSC-2013-2757

NASA Image and Video Library

2013-06-13

In the Mojave Desert in California, students and engineers checkout the Garvey Spacecraft Corporation's Prospector P-18D rocket engine. The rocket is scheduled for launch June 15 with the RUBICS-1 payload on a high-altitude, suborbital flight. The rocket will carry four satellites made from four-inch cube section. Collectively known as CubeSats, the satellites will record shock, vibrations and heat inside the rocket. They will not be released during the test flight, but the results will be used to prove or strengthen their designs before they are carried into orbit in 2014 on a much larger rocket. A new, lightweight carrier is also being tested for use on future missions to deploy the small spacecraft. The flight also is being watched closely as a model for trying out new or off-the-shelf technologies quickly before putting them in the pipeline for use on NASA's largest launchers. Built by several different organizations, including a university, a NASA field center and a high school, the spacecraft are four-inch cubes designed to fly on their own eventually, but will remain firmly attached to the rocket during the upcoming mission. For more information, visit http://www.nasa.gov/mission_pages/smallsats/elana/cubesatlaunchpreview.html Photo credit: NASA/Dimitri Gerondidakis

KSC-2013-2758

NASA Image and Video Library

2013-06-13

In the Mojave Desert in California, students and engineers checkout the Garvey Spacecraft Corporation's Prospector P-18D rocket engine. The rocket is scheduled for launch June 15 with the RUBICS-1 payload on a high-altitude, suborbital flight. The rocket will carry four satellites made from four-inch cube section. Collectively known as CubeSats, the satellites will record shock, vibrations and heat inside the rocket. They will not be released during the test flight, but the results will be used to prove or strengthen their designs before they are carried into orbit in 2014 on a much larger rocket. A new, lightweight carrier is also being tested for use on future missions to deploy the small spacecraft. The flight also is being watched closely as a model for trying out new or off-the-shelf technologies quickly before putting them in the pipeline for use on NASA's largest launchers. Built by several different organizations, including a university, a NASA field center and a high school, the spacecraft are four-inch cubes designed to fly on their own eventually, but will remain firmly attached to the rocket during the upcoming mission. For more information, visit http://www.nasa.gov/mission_pages/smallsats/elana/cubesatlaunchpreview.html Photo credit: NASA/Dimitri Gerondidakis

KSC-2013-2759

NASA Image and Video Library

2013-06-13

In the Mojave Desert in California, students and engineers checkout the Garvey Spacecraft Corporation's Prospector P-18D rocket engine. The rocket is scheduled for launch June 15 with the RUBICS-1 payload on a high-altitude, suborbital flight. The rocket will carry four satellites made from four-inch cube section. Collectively known as CubeSats, the satellites will record shock, vibrations and heat inside the rocket. They will not be released during the test flight, but the results will be used to prove or strengthen their designs before they are carried into orbit in 2014 on a much larger rocket. A new, lightweight carrier is also being tested for use on future missions to deploy the small spacecraft. The flight also is being watched closely as a model for trying out new or off-the-shelf technologies quickly before putting them in the pipeline for use on NASA's largest launchers. Built by several different organizations, including a university, a NASA field center and a high school, the spacecraft are four-inch cubes designed to fly on their own eventually, but will remain firmly attached to the rocket during the upcoming mission. For more information, visit http://www.nasa.gov/mission_pages/smallsats/elana/cubesatlaunchpreview.html Photo credit: NASA/Dimitri Gerondidakis

The Strutjet Rocket Based Combined Cycle Engine

NASA Technical Reports Server (NTRS)

Siebenhaar, A.; Bulman, M. J.; Bonnar, D. K.

1998-01-01

The multi stage chemical rocket has been established over many years as the propulsion System for space transportation vehicles, while, at the same time, there is increasing concern about its continued affordability and rather involved reusability. Two broad approaches to addressing this overall launch cost problem consist in one, the further development of the rocket motor, and two, the use of airbreathing propulsion to the maximum extent possible as a complement to the limited use of a conventional rocket. In both cases, a single-stage-to-orbit (SSTO) vehicle is considered a desirable goal. However, neither the "all-rocket" nor the "all-airbreathing" approach seems realizable and workable in practice without appreciable advances in materials and manufacturing. An affordable system must be reusable with minimal refurbishing on-ground, and large mean time between overhauls, and thus with high margins in design. It has been suggested that one may use different engine cycles, some rocket and others airbreathing, in a combination over a flight trajectory, but this approach does not lead to a converged solution with thrust-to-mass, specific impulse, and other performance and operational characteristics that can be obtained in the different engines. The reason is this type of engine is simply a combination of different engines with no commonality of gas flowpath or components, and therefore tends to have the deficiencies of each of the combined engines. A further development in this approach is a truly combined cycle that incorporates a series of cycles for different modes of propulsion along a flight path with multiple use of a set of components and an essentially single gas flowpath through the engine. This integrated approach is based on realizing the benefits of both a rocket engine and airbreathing engine in various combinations by a systematic functional integration of components in an engine class usually referred to as a rocket-based combined cycle (RBCC) engine. RBCC engines exhibit a high potential for lowering the operating cost of launching payloads into orbit. Two sources of cost reductions can be identified. First, RBCC powered vehicles require only 20% takeoff thrust compared to conventional rockets, thereby lowering the thrust requirements and the replacement cost of the engines. Second, due to the higher structural and thermal margins achievable with RBCC engines coupled with a higher degree of subsystem redundance lower maintenance and operating cost are obtainable.

The pasty propellant rocket engine development

NASA Astrophysics Data System (ADS)

Kukushkin, V. I.; Ivanchenko, A. N.

1993-06-01

The paper describes a newly developed pasty propellant rocket engine (PPRE) and the combustion process and presents results of performance tests. It is shown that, compared with liquid propellant rocket engines, the PPREs can regulate the thrust level within a wider range, are safer ecologically, and have better weight characteristics. Compared with solid propellant rocket engines, the PPREs may be produced with lower costs and more safely, are able to regulate thrust performance within a wider range, and are able to offer a greater scope for the variation of the formulation components and propellant characteristics. Diagrams of the PPRE are included.

Boundary layer integral matrix procedure: Verification of models

NASA Technical Reports Server (NTRS)

Bonnett, W. S.; Evans, R. M.

1977-01-01

The three turbulent models currently available in the JANNAF version of the Aerotherm Boundary Layer Integral Matrix Procedure (BLIMP-J) code were studied. The BLIMP-J program is the standard prediction method for boundary layer effects in liquid rocket engine thrust chambers. Experimental data from flow fields with large edge-to-wall temperature ratios are compared to the predictions of the three turbulence models contained in BLIMP-J. In addition, test conditions necessary to generate additional data on a flat plate or in a nozzle are given. It is concluded that the Cebeci-Smith turbulence model be the recommended model for the prediction of boundary layer effects in liquid rocket engines. In addition, the effects of homogeneous chemical reaction kinetics were examined for a hydrogen/oxygen system. Results show that for most flows, kinetics are probably only significant for stoichiometric mixture ratios.

Numerical analysis of the hot-gas-side and coolant-side heat transfer in liquid rocket engine combustors

NASA Technical Reports Server (NTRS)

Wang, Ten-See; Van, Luong

1992-01-01

The objective of this paper are to develop a multidisciplinary computational methodology to predict the hot-gas-side and coolant-side heat transfer and to use it in parametric studies to recommend optimized design of the coolant channels for a regeneratively cooled liquid rocket engine combustor. An integrated numerical model which incorporates CFD for the hot-gas thermal environment, and thermal analysis for the liner and coolant channels, was developed. This integrated CFD/thermal model was validated by comparing predicted heat fluxes with those of hot-firing test and industrial design methods for a 40 k calorimeter thrust chamber and the Space Shuttle Main Engine Main Combustion Chamber. Parametric studies were performed for the Advanced Main Combustion Chamber to find a strategy for a proposed combustion chamber coolant channel design.

-----SPACE TRANSPORTATION

NASA Image and Video Library

1998-10-07

This photograph depicts an air-breathing rocket engine prototype in the test bay at the General Applied Science Lab facility in Ronkonkoma, New York. Air-breathing engines, known as rocket based, combined-cycle engines, get their initial take-off power from specially designed rockets, called air-augmented rockets, that boost performance about 15 percent over conventional rockets. When the vehicle's velocity reaches twice the speed of sound, the rockets are turned off and the engine relies totally on oxygen in the atmosphere to burn hydrogen fuel, as opposed to a rocket that must carry its own oxygen, thus reducing weight and flight costs. Once the vehicle has accelerated to about 10 times the speed of sound, the engine converts to a conventional rocket-powered system to propel the craft into orbit or sustain it to suborbital flight speed. NASA's Advanced Space Transportation Program at Marshall Space Flight Center, along with several industry partners and collegiate forces, is developing this technology to make space transportation affordable for everyone from business travelers to tourists. The goal is to reduce launch costs from today's price tag of $10,000 per pound to only hundreds of dollars per pound. NASA's series of hypersonic flight demonstrators currently include three air-breathing vehicles: the X-43A, X-43B and X-43C.

30. Historic view of twentythousandpound rocket test stand with engine ...

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

30. Historic view of twenty-thousand-pound rocket test stand with engine installation in test cell of Building 202, looking down from elevated location, September 1957. On file at NASA Plumbrook Research Center, Sandusky, Ohio. NASA GRC photo number C-45872. - Rocket Engine Testing Facility, GRC Building No. 202, NASA Glenn Research Center, Cleveland, Cuyahoga County, OH

14 CFR Appendix E to Part 25 - Appendix E to Part 25

Code of Federal Regulations, 2013 CFR

2013-01-01

... certificated takeoff and landing weights of an airplane equipped with a type-certificated standby power rocket engine may obtain an increase as specified in paragraph (b) if— (1) The installation of the rocket engine has been approved and it has been established by flight test that the rocket engine and its controls...

14 CFR Appendix E to Part 25 - Appendix E to Part 25

Code of Federal Regulations, 2011 CFR

2011-01-01

... certificated takeoff and landing weights of an airplane equipped with a type-certificated standby power rocket engine may obtain an increase as specified in paragraph (b) if— (1) The installation of the rocket engine has been approved and it has been established by flight test that the rocket engine and its controls...

14 CFR Appendix E to Part 25 - Appendix E to Part 25

Code of Federal Regulations, 2014 CFR

2014-01-01

... certificated takeoff and landing weights of an airplane equipped with a type-certificated standby power rocket engine may obtain an increase as specified in paragraph (b) if— (1) The installation of the rocket engine has been approved and it has been established by flight test that the rocket engine and its controls...

14 CFR Appendix E to Part 25 - Appendix E to Part 25

Code of Federal Regulations, 2012 CFR

2012-01-01

... certificated takeoff and landing weights of an airplane equipped with a type-certificated standby power rocket engine may obtain an increase as specified in paragraph (b) if— (1) The installation of the rocket engine has been approved and it has been established by flight test that the rocket engine and its controls...

14 CFR Appendix E to Part 25 - Appendix E to Part 25

Code of Federal Regulations, 2010 CFR

2010-01-01

... certificated takeoff and landing weights of an airplane equipped with a type-certificated standby power rocket engine may obtain an increase as specified in paragraph (b) if— (1) The installation of the rocket engine has been approved and it has been established by flight test that the rocket engine and its controls...

Mean Flow Augmented Acoustics in Rocket Systems

NASA Technical Reports Server (NTRS)

Fischbach, Sean

2014-01-01

Combustion instability in solid rocket motors and liquid engines has long been a subject of concern. Many rockets display violent fluctuations in pressure, velocity, and temperature originating from the complex interactions between the combustion process and gas dynamics. Recent advances in energy based modeling of combustion instabilities require accurate determination of acoustic frequencies and mode shapes. Of particular interest is the acoustic mean flow interactions within the converging section of a rocket nozzle, where gradients of pressure, density, and velocity become large. The expulsion of unsteady energy through the nozzle of a rocket is identified as the predominate source of acoustic damping for most rocket systems. Recently, an approach to address nozzle damping with mean flow effects was implemented by French [1]. This new approach extends the work originated by Sigman and Zinn [2] by solving the acoustic velocity potential equation (AVPE) formulated by perturbing the Euler equations [3]. The present study aims to implement the French model within the COMSOL Multiphysiscs framework and analyzes one of the author's presented test cases.

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.

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.

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

Bolonkin, A.

A first-hand account of developments in the Soviet rocket industry is presented. The organization and leadership of the rocket and missile industry are traced from its beginning in the 1920s. The development of the Glushko Experimental Design Bureau, where the majority of Soviet rocket engines were created, is related. The evolution of Soviet rocket engines is traced in regard to both their technical improvement and their application in missiles and space vehicles. Improved Glushko engines and specialized Isaev and Kosberg engines are discussed. The difficulties faced by the Soviet missile and space program, such as the pre-Sputnik failures, the oscillationmore » problem of 1965/1966, which exposed a weakness in Soviet ICBM missiles, and the Nedelin disaster of 1960, which cost the lives of more than 200 scientists and engineers, as well as the Commander-in-Chief of the Strategic Rocket Forces, Marshall Nedelin, are examined. 122 refs.« less

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.

Nuclear Engine System Simulation (NESS) version 2.0

NASA Technical Reports Server (NTRS)

Pelaccio, Dennis G.; Scheil, Christine M.; Petrosky, Lyman J.

1993-01-01

The topics are presented in viewgraph form and include the following; nuclear thermal propulsion (NTP) engine system analysis program development; nuclear thermal propulsion engine analysis capability requirements; team resources used to support NESS development; expanded liquid engine simulations (ELES) computer model; ELES verification examples; NESS program development evolution; past NTP ELES analysis code modifications and verifications; general NTP engine system features modeled by NESS; representative NTP expander, gas generator, and bleed engine system cycles modeled by NESS; NESS program overview; NESS program flow logic; enabler (NERVA type) nuclear thermal rocket engine; prismatic fuel elements and supports; reactor fuel and support element parameters; reactor parameters as a function of thrust level; internal shield sizing; and reactor thermal model.

Thermal stratification potential in rocket engine coolant channels

NASA Technical Reports Server (NTRS)

Kacynski, Kenneth J.

1992-01-01

The potential for rocket engine coolant channel flow stratification was computationally studied. A conjugate, 3-D, conduction/advection analysis code (SINDA/FLUINT) was used. Core fluid temperatures were predicted to vary by over 360 K across the coolant channel, at the throat section, indicating that the conventional assumption of a fully mixed fluid may be extremely inaccurate. Because of the thermal stratification of the fluid, the walls exposed to the rocket engine exhaust gases will be hotter than an assumption of full mixing would imply. In this analysis, wall temperatures were 160 K hotter in the turbulent mixing case than in the full mixing case. The discrepancy between the full mixing and turbulent mixing analyses increased with increasing heat transfer. Both analysis methods predicted identical channel resistances at the coolant inlet, but in the stratified analysis the thermal resistance was negligible. The implications are significant. Neglect of thermal stratification could lead to underpredictions in nozzle wall temperatures. Even worse, testing at subscale conditions may be inadequate for modeling conditions that would exist in a full scale engine.

Comparison of Rocket Performance using Exhaust Diffuser and Conventional Techniques for Altitude Simulation

NASA Technical Reports Server (NTRS)

Sivo, Joseph N.; Peters, Daniel J.

1959-01-01

A rocket engine with an exhaust-nozzle area ratio of 25 was operated at a constant chamber pressure of 600 pounds per square inch absolute over a range of oxidant-fuel ratios at an altitude pressure corresponding to approximately 47,000 feet. At this condition, the nozzle flow is slightly underexpanded as it leaves the nozzle. The altitude simulation was obtained first through the use of an exhaust diffuser coupled with the rocket engine and secondly, in an altitude test chamber where separate exhauster equipment provided the altitude pressure. A comparison of performance data from these two tests has established that a diffuser used with a rocket engine operating at near-design nozzle pressure ratio can be a valid means of obtaining altitude performance data for rocket engines.

-----SPACE TRANSPORTATION

NASA Image and Video Library

2000-05-01

This photograph depicts an air-breathing rocket engine that completed an hour or 3,600 seconds of testing at the General Applied Sciences Laboratory in Ronkonkoma, New York. Referred to as ARGO by its design team, the engine is named after the mythological Greek ship that bore Jason and the Argonauts on their epic voyage of discovery. Air-breathing engines, known as rocket based, combined-cycle engines, get their initial take-off power from specially designed rockets, called air-augmented rockets, that boost performance about 15 percent over conventional rockets. When the vehicle's velocity reaches twice the speed of sound, the rockets are turned off and the engine relies totally on oxygen in the atmosphere to burn hydrogen fuel, as opposed to a rocket that must carry its own oxygen, thus reducing weight and flight costs. Once the vehicle has accelerated to about 10 times the speed of sound, the engine converts to a conventional rocket-powered system to propel the craft into orbit or sustain it to suborbital flight speed. NASA's Advanced SpaceTransportation Program at Marshall Space Flight Center, along with several industry partners and collegiate forces, is developing this technology to make space transportation affordable for everyone from business travelers to tourists. The goal is to reduce launch costs from today's price tag of $10,000 per pound to only hundreds of dollars per pound. NASA's series of hypersonic flight demonstrators currently include three air-breathing vehicles: the X-43A, X-43B and X-43C.

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

Liquid-propellant rocket engines health-monitoring—a survey

NASA Astrophysics Data System (ADS)

Wu, Jianjun

2005-02-01

This paper is intended to give a summary on the health-monitoring technology, which is one of the key technologies both for improving and enhancing the reliability and safety of current rocket engines and for developing new-generation high reliable reusable rocket engines. The implication of health-monitoring and the fundamental principle obeyed by the fault detection and diagnostics are elucidated. The main aspects of health-monitoring such as system frameworks, failure modes analysis, algorithms of fault detection and diagnosis, control means and advanced sensor techniques are illustrated in some detail. At last, the evolution trend of health-monitoring techniques of liquid-propellant rocket engines is set out.

Performance and Stability Analyses of Rocket Thrust Chambers with Oxygen/Methane Propellants

NASA Technical Reports Server (NTRS)

Hulka, James R.; Jones, Gregg W.

2010-01-01

Liquid rocket engines using oxygen and methane propellants are being considered by the National Aeronautics and Space Administration (NASA) for future in-space vehicles. This propellant combination has not been previously used in flight-qualified engine systems developed by NASA, so limited test data and analysis results are available at this stage of early development. As part of activities for the Propulsion and Cryogenic Advanced Development (PCAD) project funded under the Exploration Technology Development Program, the NASA Marshall Space Flight Center (MSFC) has been evaluating capability to model combustion performance and stability for oxygen and methane propellants. This activity has been proceeding for about two years and this paper is a summary of results to date. Hot-fire test results of oxygen/methane propellant rocket engine combustion devices for the modeling investigations have come from several sources, including multi-element injector tests with gaseous methane from the 1980s, single element tests with gaseous methane funded through the Constellation University Institutes Program, and multi-element injector tests with both gaseous and liquid methane conducted at the NASA MSFC funded by PCAD. For the latter, test results of both impinging and coaxial element injectors using liquid oxygen and liquid methane propellants are included. Configurations were modeled with two one-dimensional liquid rocket combustion analysis codes, the Rocket Combustor Interactive Design and Analysis code and the Coaxial Injector Combustion Model. Special effort was focused on 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, improved or developed in the future. Low frequency combustion instability (chug) occurred, with frequencies ranging from 150 to 250 Hz, with several multi-element injectors with liquid/liquid propellants, and was modeled using techniques from Wenzel and Szuch. High-frequency combustion instability also occurred at the first tangential (1T) mode, at about 4500 Hz, with several multi-element injectors with liquid/liquid propellants. Analyses of the transverse mode instability were conducted by evaluating injector resonances and empirical methods developed by Hewitt.

The development of a post-test diagnostic system for rocket engines

NASA Technical Reports Server (NTRS)

Zakrajsek, June F.

1991-01-01

An effort was undertaken by NASA to develop an automated post-test, post-flight diagnostic system for rocket engines. The automated system is designed to be generic and to automate the rocket engine data review process. A modular, distributed architecture with a generic software core was chosen to meet the design requirements. The diagnostic system is initially being applied to the Space Shuttle Main Engine data review process. The system modules currently under development are the session/message manager, and portions of the applications section, the component analysis section, and the intelligent knowledge server. An overview is presented of a rocket engine data review process, the design requirements and guidelines, the architecture and modules, and the projected benefits of the automated diagnostic system.

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.

Results of an investigation of jet plume effects on an 0.010-scale model (75-OTS) of the space shuttle integrated vehicle in the 9 x 7-foot leg of the NASA/Ames unitary wind tunnel (IA82B), volume 1. [an exhaust flow simulation

NASA Technical Reports Server (NTRS)

Hawthorne, P. J.

1976-01-01

The base pressure environment was investigated for the first and second stage mated vehicle in a supersonic flow field from Mach 1.55 through 2.20 with simulated rocket engine exhaust plumes. The pressure environment was investigated for the orbiter at various vent port locations at these same freestream conditions. The Mach number environment around the base of the model with rocket plumes simulated was examined. Data were obtained at angles of attack from -4 deg through +4 deg at zero yaw, and at yaw angles from -4 deg through +4 deg at zero angle of attack, with rocket plume sizes varying from smaller than nominal to much greater than nominal. Failed orbiter engine data were also obtained. Elevon hinge moments and wing panel load data were obtained during all runs. Photographs of the tested configurations are shown.

Effect of Swirl on an Unstable Single-Element Gas-Gas Rocket Engine

DTIC Science & Technology

2014-06-01

at 300 K, and the combustor is filled with a mixture of water and carbon dioxide at 1500 K. The warmer temperature in the combustor enables the auto...a variety of configurations including gas turbines and rocket engines.4–13 The single-element engine chosen for this study is the continuously...combustion systems including gas turbines , rocket engines, and industrial furnaces. Swirl can have dramatic effects on the flowfield; these include jet growth

High temperature reformation of aluminum and chlorine compounds behind the Mach disk of a solid-fuel rocket exhaust

NASA Technical Reports Server (NTRS)

Park, C.

1976-01-01

Chemical reactions expected to occur among the constituents of solid-fuel rocket engine effluents in the hot region behind a Mach disk are analyzed theoretically. With the use of a rocket plume model that assumes the flow to be separated in the base region, and a chemical reaction scheme that includes evaporation of alumina and the associated reactions of 17 gas species, the reformation of the effluent is calculated. It is shown that AlClO and AlOH are produced in exchange for a corresponding reduction in the amounts of HCl and Al2O3. For the case of the space shuttle booster engines, up to 2% of the original mass of the rocket fuel can possibly be converted to these two new species and deposited in the atmosphere between the altitudes of 10 and 40 km. No adverse effects on the atmospheric environment are anticipated with the addition of these two new species.

Comparison of theoretical and experimental thrust performance of a 1030:1 area ratio rocket nozzle at a chamber pressure of 2413 kN/sq m (350 psia)

NASA Technical Reports Server (NTRS)

Smith, Tamara A.; Pavli, Albert J.; Kacynski, Kenneth J.

1987-01-01

The Joint Army, Navy, NASA, Air Force (JANNAF) rocket-engine performance-prediction procedure is based on the use of various reference computer programs. One of the reference programs for nozzle analysis is the Two-Dimensional Kinetics (TDK) Program. The purpose of this report is to calibrate the JANNAF procedure that has been incorporated into the December 1984 version of the TDK program for the high-area-ratio rocket-engine regime. The calibration was accomplished by modeling the performance of a 1030:1 rocket nozzle tested at NASA Lewis. A detailed description of the test conditions and TDK input parameters is given. The reuslts indicate that the computer code predicts delivered vacuum specific impulse to within 0.12 to 1.9 percent of the experimental data. Vacuum thrust coefficient predictions were within + or - 1.3 percent of experimental results. Predictions of wall static pressure were within approximately + or - 5 percent of the measured values.

A History of Welding on the Space Shuttle Main Engine (1975 to 2010)

NASA Technical Reports Server (NTRS)

Zimmerman, Frank R.; Russell, Carolyn K.

2010-01-01

The Space Shuttle Main Engine (SSME) is a high performance, throttleable, liquid hydrogen fueled rocket engine. High thrust and specific impulse (Isp) are achieved through a staged combustion engine cycle, combined with high combustion pressure (approx.3000psi) generated by the two-stage pump and combustion process. The SSME is continuously throttleable from 67% to 109% of design thrust level. The design criteria for this engine maximize performance and weight, resulting in a 7,800 pound rocket engine that produces over a half million pounds of thrust in vacuum with a specific impulse of 452/sec. It is the most reliable rocket engine in the world, accumulating over one million seconds of hot-fire time and achieving 100% flight success in the Space Shuttle program. A rocket engine with the unique combination of high reliability, performance, and reusability comes at the expense of manufacturing simplicity. Several innovative design features and fabrication techniques are unique to this engine. This is as true for welding as any other manufacturing process. For many of the weld joints it seemed mean cheating physics and metallurgy to meet the requirements. This paper will present a history of the welding used to produce the world s highest performance throttleable rocket engine.

Rocket-Based Combined Cycle Flowpath Testing for Modes 1 and 4

NASA Technical Reports Server (NTRS)

Rice, Tharen

2002-01-01

Under sponsorship of the NASA Glenn Research Center (NASA GRC), the Johns Hopkins University Applied Physics Laboratory (JHU/APL) designed and built a five-inch diameter, Rocket-Based Combined Cycle (RBCC) engine to investigate mode 1 and mode 4 engine performance as well as Mach 4 inlet performance. This engine was designed so that engine area and length ratios were similar to the NASA GRC GTX engine is shown. Unlike the GTX semi-circular engine design, the APL engine is completely axisymmetric. For this design, a traditional rocket thruster was installed inside of the scramjet flowpath, along the engine centerline. A three part test series was conducted to determine Mode I and Mode 4 engine performance. In part one, testing of the rocket thruster alone was accomplished and its performance determined (average Isp efficiency = 90%). In part two, Mode 1 (air-augmented rocket) testing was conducted at a nominal chamber pressure-to-ambient pressure ratio of 100 with the engine inlet fully open. Results showed that there was neither a thrust increment nor decrement over rocket-only thrust during Mode 1 operation. In part three, Mode 4 testing was conducted with chamber pressure-to-ambient pressure ratios lower than desired (80 instead of 600) with the inlet fully closed. Results for this testing showed a performance decrease of 20% as compared to the rocket-only testing. It is felt that these results are directly related to the low pressure ratio tested and not the engine design. During this program, Mach 4 inlet testing was also conducted. For these tests, a moveable centerbody was tested to determine the maximum contraction ratio for the engine design. The experimental results agreed with CFD results conducted by NASA GRC, showing a maximum geometric contraction ratio of approximately 10.5. This report details the hardware design, test setup, experimental results and data analysis associated with the aforementioned tests.

Collaborative Sounding Rocket launch in Alaska and Development of Hybrid Rockets

NASA Astrophysics Data System (ADS)

Ono, Tomohisa; Tsutsumi, Akimasa; Ito, Toshiyuki; Kan, Yuji; Tohyama, Fumio; Nakashino, Kyouichi; Hawkins, Joseph

Tokai University student rocket project (TSRP) was established in 1995 for a purpose of the space science and engineering hands-on education, consisting of two space programs; the one is sounding rocket experiment collaboration with University of Alaska Fairbanks and the other is development and launch of small hybrid rockets. In January of 2000 and March 2002, two collaborative sounding rockets were successfully launched at Poker Flat Research Range in Alaska. In 2001, the first Tokai hybrid rocket was successfully launched at Alaska. After that, 11 hybrid rockets were launched to the level of 180-1,000 m high at Hokkaido and Akita in Japan. Currently, Tokai students design and build all parts of the rockets. In addition, they are running the organization and development of the project under the tight budget control. This program has proven to be very effective in providing students with practical, real-engineering design experience and this program also allows students to participate in all phases of a sounding rocket mission. Also students learn scientific, engineering subjects, public affairs and system management through experiences of cooperative teamwork. In this report, we summarize the TSRP's hybrid rocket program and discuss the effectiveness of the program in terms of educational aspects.

Ceramic composites for rocket engine turbines

NASA Technical Reports Server (NTRS)

Herbell, Thomas P.; Eckel, Andrew J.

1991-01-01

The use of ceramic materials in the hot section of the fuel turbopump of advanced reusable rocket engines promises increased performance and payload capability, improved component life and economics, and greater design flexibility. Severe thermal transients present during operation of the Space Shuttle Main Engine (SSME), push metallic components to the limit of their capabilities. Future engine requirements might be even more severe. In phase one of this two-phase program, performance benefits were quantified and continuous fiber reinforced ceramic matrix composite components demonstrated a potential to survive the hostile environment of an advanced rocket engine turbopump.

Ceramic composites for rocket engine turbines

NASA Technical Reports Server (NTRS)

Herbell, Thomas P.; Eckel, Andrew J.

1991-01-01

The use of ceramic materials in the hot section of the fuel turbopump of advanced reusable rocket engines promises increased performance and payload capability, improved component life and economics, and greater design flexibility. Severe thermal transients present during operation of the Space Shuttle Main Engine (SSME), push metallic components to the limit of their capabilities. Future engine requirements might be even more severe. In phase one of this two-phase program, performance benefits were quantified and continuous fiber reinforced ceramic matrix composite components demonstrated a potential to survive the hostile environment of an advaced rocket engine turbopump.

Done in 60 seconds- See a Massive Rocket Fuel Tank Built in A Minute

NASA Image and Video Library

2016-08-18

The 7.5-minute test conducted at NASA’s Stennis Space Center is part of a series of tests designed to put the upgraded former space shuttle engines through the rigorous temperature and pressure conditions they will experience during a launch. The tests also support the development of a new controller, or “brain,” for the engine, which monitors engine status and communicates between the rocket and the engine, relaying commands to the engine and transmitting data back to the rocket.

KSC-2013-4342

NASA Image and Video Library

2013-12-11

CAPE CANAVERAL, Fla. -- At NASA's Kennedy Space Center in Florida, from the left, Leandro James, rocket avionics lead, Gary Dahlke, high powered rocket subject matter expert, and Julio Najarro of Mechanical Systems make final adjustments to a small rocket prior to launch as part of Rocket University. The launch will test systems designed by the student engineers. As part of Rocket University, the engineers are given an opportunity to work a fast-track project to develop skills in developing spacecraft systems of the future. As NASA plans for future spaceflight programs to low-Earth orbit and beyond, teams of engineers at Kennedy are gaining experience in designing and flying launch vehicle systems on a small scale. Four teams of five to eight members from Kennedy are designing rockets complete with avionics and recovery systems. Launch operations require coordination with federal agencies, just as they would with rockets launched in support of a NASA mission. Photo credit: NASA/Jim Grossmann

Celebrating 50 Years of Testing

NASA Image and Video Library

2016-04-19

What better way to mark 50 years of rocket engine testing than with a rocket engine test? Stennis Space Center employees enjoyed a chance to view an RS-68 engine test at the B-1 Test Stand on April 19, almost 50 years to the day that the first test was conducted at the south Mississippi site in 1966. The test viewing was part of a weeklong celebration of the 50th year of rocket engine testing at Stennis. The first test at the site occurred April 23, 1966, with a 15-second firing of a Saturn V second stage prototype (S-II-C) on the A-2 Test Stand. The center subsequently tested Apollo rocket stages that carried humans to the moon and every main engine used to power 135 space shuttle missions. It currently tests engines for NASA’s new Space Launch System vehicle.

Space vehicle engine and heat shield environment review. Volume 1: Engineering analysis

NASA Technical Reports Server (NTRS)

Mcanelly, W. B.; Young, C. T. K.

1973-01-01

Methods for predicting the base heating characteristics of a multiple rocket engine installation are discussed. The environmental data is applied to the design of adequate protection system for the engine components. The methods for predicting the base region thermal environment are categorized as: (1) scale model testing, (2) extrapolation of previous and related flight test results, and (3) semiempirical analytical techniques.

Saturn systems holddown acoustic efficiency and normalized acoustic power spectrum.

NASA Technical Reports Server (NTRS)

Gilbert, D. W.

1972-01-01

Saturn systems field acoustic data are used to derive mid- and far-field prediction parameters for rocket engine noise. The data were obtained during Saturn vehicle launches at the Kennedy Space Center. The data base is a sorted set of acoustic data measured during the period 1961 through 1971 for Saturn system launches SA-1 through AS-509. The model assumes hemispherical radiation from a simple source located at the intersection of the longitudinal axis of each booster and the engine exit plane. The model parameters are evaluated only during vehicle holddown. The acoustic normalized power spectrum and efficiency for each system are isolated as a composite from the data using linear numerical methods. The specific definitions of each allows separation. The resulting power spectra are nondimensionalized as a function of rocket engine parameters. The nondimensional Saturn system acoustic spectrum and efficiencies are compared as a function of Strouhal number with power spectra from other systems.

Easier Analysis With Rocket Science

NASA Technical Reports Server (NTRS)

2003-01-01

Analyzing rocket engines is one of Marshall Space Flight Center's specialties. When Marshall engineers lacked a software program flexible enough to meet their needs for analyzing rocket engine fluid flow, they overcame the challenge by inventing the Generalized Fluid System Simulation Program (GFSSP), which was named the co-winner of the NASA Software of the Year award in 2001. This paper describes the GFSSP in a wide variety of applications

Orbital transfer rocket engine technology program: Soft wear ring seal technology

NASA Technical Reports Server (NTRS)

Lariviere, Brian W.

1992-01-01

Liquid oxygen (LOX) compatibility tests, including autogenous ignition, promoted ignition, LOX impact tests, and friction and wear tests on different PV products were conducted for several polymer materials as verification for the implementation of soft wear ring seals in advanced rocket engine turbopumps. Thermoplastics, polyimide based materials, and polyimide-imide base materials were compared for oxygen compatibility, specific wear coefficient, wear debris production, and heat dissipation mechanisms. A thermal model was generated that simulated the frictional heating input and calculated the surface temperature and temperature distribution within the seal. The predictions were compared against measured values. Heat loads in the model were varied to better match the test data and determine the difference between the measured and the calculated coefficients of friction.

AJ26 rocket engine testing news briefing

NASA Technical Reports Server (NTRS)

2010-01-01

Operators at NASA's John C. Stennis Space Center are completing modifications to the E-1 Test Stand to begin testing Aerojet AJ26 rocket engines in early summer of 2010. Modifications include construction of a 27-foot-deep flame deflector trench. The AJ26 rocket engines will be used to power Orbital Sciences Corp.'s Taurus II space vehicles to provide commercial cargo transportation missions to the International Space Station for NASA. Stennis has partnered with Orbital to test all engines for the transport missions.

Iridium/Rhenium Parts For Rocket Engines

NASA Technical Reports Server (NTRS)

Schneider, Steven J.; Harding, John T.; Wooten, John R.

1991-01-01

Oxidation/corrosion of metals at high temperatures primary life-limiting mechanism of parts in rocket engines. Combination of metals greatly increases operating temperature and longevity of these parts. Consists of two transition-element metals - iridium and rhenium - that melt at extremely high temperatures. Maximum operating temperature increased to 2,200 degrees C from 1,400 degrees C. Increases operating lifetimes of small rocket engines by more than factor of 10. Possible to make hotter-operating, longer-lasting components for turbines and other heat engines.

Space Launch System Base Heating Test: Sub-Scale Rocket Engine/Motor Design, Development and Performance Analysis

NASA Technical Reports Server (NTRS)

Mehta, Manish; Seaford, Mark; Kovarik, Brian; Dufrene, Aaron; Solly, Nathan; Kirchner, Robert; Engel, Carl D.

2014-01-01

The Space Launch System (SLS) base heating test is broken down into two test programs: (1) Pathfinder and (2) Main Test. The Pathfinder Test Program focuses on the design, development, hot-fire test and performance analyses of the 2% sub-scale SLS core-stage and booster element propulsion systems. The core-stage propulsion system is composed of four gaseous oxygen/hydrogen RS-25D model engines and the booster element is composed of two aluminum-based model solid rocket motors (SRMs). The first section of the paper discusses the motivation and test facility specifications for the test program. The second section briefly investigates the internal flow path of the design. The third section briefly shows the performance of the model RS-25D engines and SRMs for the conducted short duration hot-fire tests. Good agreement is observed based on design prediction analysis and test data. This program is a challenging research and development effort that has not been attempted in 40+ years for a NASA vehicle.

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.

NASA’s Space Launch System Engine Testing Heats Up

NASA Image and Video Library

2017-05-23

NASA engineers successfully conducted the second in a series of RS-25 flight controller tests on May 23, 2017, for the world’s most-powerful rocket. The 500-second test on the A-1 Test Stand at NASA’s Stennis Space Center in Mississippi marked another milestone toward launch of NASA’s new Space Launch System (SLS) rocket on its inaugural flight, the Exploration Mission-1 (EM-1). The SLS rocket, powered by four RS-25 engines, will provide 2 million pounds of thrust and work in conjunction with two solid rocket boosters. These are former space shuttle main engines, modified to perform at a higher level and with a new controller.

Program For Optimization Of Nuclear Rocket Engines

NASA Technical Reports Server (NTRS)

Plebuch, R. K.; Mcdougall, J. K.; Ridolphi, F.; Walton, James T.

1994-01-01

NOP is versatile digital-computer program devoloped for parametric analysis of beryllium-reflected, graphite-moderated nuclear rocket engines. Facilitates analysis of performance of engine with respect to such considerations as specific impulse, engine power, type of engine cycle, and engine-design constraints arising from complications of fuel loading and internal gradients of temperature. Predicts minimum weight for specified performance.

An Historical Perspective of the NERVA Nuclear Rocket Engine Technology Program

NASA Technical Reports Server (NTRS)

Robbins, W. H.; Finger, H. B.

1991-01-01

Nuclear rocket research and development was initiated in the United States in 1955 and is still being pursued to a limited extent. The major technology emphasis occurred in the decade of the 1960s and was primarily associated with the Rover/NERVA Program where the technology for a nuclear rocket engine system for space application was developed and demonstrated. The NERVA (Nuclear Engine for Rocket Vehicle Application) technology developed twenty years ago provides a comprehensive and viable propulsion technology base that can be applied and will prove to be valuable for application to the NASA Space Exploration Initiative (SEI). This paper, which is historical in scope, provides an overview of the conduct of the NERVA Engine Program, its organization and management, development philosophy, the engine configuration, and significant accomplishments.

Design of Force Sensor Leg for a Rocket Thrust Detector

NASA Astrophysics Data System (ADS)

Woten, Douglas; McGehee, Tripp; Wright, Anne

2005-03-01

A hybrid rocket is composed of a solid fuel and a separate liquid or gaseous oxidizer. These rockets may be throttled like liquid rockets, are safer than solid rockets, and are much less complex than liquid rockets. However, hybrid rockets produce thrust oscillations that are not practical for large scale use. A lab scale hybrid rocket at the University of Arkansas at Little Rock (UALR) Hybrid Rocket Facility is used to develop sensors to measure physical properties of hybrid rockets. Research is currently being conducted to design a six degree of freedom force sensor to measure the thrust and torque in all three spacial dimensions. The detector design uses six force sensor legs. Each leg utilizes strain gauges and a Wheatstone bridge to produce a voltage propotional to the force on the leg. The leg was designed using the CAD software ProEngineer and ProMechanica. Computer models of the strains on the single leg will be presented. A prototype leg was built and was tested in an INSTRON and results will be presented.

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.

J-2X, The Engine of the Future

NASA Technical Reports Server (NTRS)

Smith, Gail

2009-01-01

My project was two-fold, with both parts involving the J-2X Upper Stage engine (which will be used on both the Ares I and V). Mainly, I am responsible for using a program called Iris to create visual represen tations of the rocket engine's telemetry data. Also, my project includes the application of my newly acquired Pro Engineer skills in develo ping a 3D model of the engine's nozzle.

Dynamics of Supercritical Flows

DTIC Science & Technology

2012-08-26

to Supercritical Environment of Relevance to Rocket, Gas turbine , and Diesel Engines,” 37th AIAA Aerospace Science Meeting and Exhibit, AIAA...Visual Characteristics of a Round Jet into a Sub- to Supercritical Environment of Relevance to Rocket, Gas turbine , and Diesel Engines,” 37th AIAA...Relevance to Rocket, Gas turbine , and Diesel Engines,” 37th AIAA Aerospace Science Meeting and Exhibit, AIAA, Washington, DC, 11-14 Jan. 1999. 26Chehroudi

Calculating Nozzle Side Loads using Acceleration Measurements of Test-Based Models

NASA Technical Reports Server (NTRS)

Brown, Andrew M.; Ruf, Joe

2007-01-01

As part of a NASA/MSFC research program to evaluate the effect of different nozzle contours on the well-known but poorly characterized "side load" phenomena, we attempt to back out the net force on a sub-scale nozzle during cold-flow testing using acceleration measurements. Because modeling the test facility dynamics is problematic, new techniques for creating a "pseudo-model" of the facility and nozzle directly from modal test results are applied. Extensive verification procedures were undertaken, resulting in a loading scale factor necessary for agreement between test and model based frequency response functions. Side loads are then obtained by applying a wide-band random load onto the system model, obtaining nozzle response PSD's, and iterating both the amplitude and frequency of the input until a good comparison of the response with the measured response PSD for a specific time point is obtained. The final calculated loading can be used to compare different nozzle profiles for assessment during rocket engine nozzle development and as a basis for accurate design of the nozzle and engine structure to withstand these loads. The techniques applied within this procedure have extensive applicability to timely and accurate characterization of all test fixtures used for modal test.A viewgraph presentation on a model-test based pseudo-model used to calculate side loads on rocket engine nozzles is included. The topics include: 1) Side Loads in Rocket Nozzles; 2) Present Side Loads Research at NASA/MSFC; 3) Structural Dynamic Model Generation; 4) Pseudo-Model Generation; 5) Implementation; 6) Calibration of Pseudo-Model Response; 7) Pseudo-Model Response Verification; 8) Inverse Force Determination; 9) Results; and 10) Recent Work.

Delta II Mars Pathfinder

NASA Technical Reports Server (NTRS)

1998-01-01

Final preparations for lift off of the DELTA II Mars Pathfinder Rocket are shown. Activities include loading the liquid oxygen, completing the construction of the Rover, and placing the Rover into the Lander. After the countdown, important visual events include the launch of the Delta Rocket, burnout and separation of the three Solid Rocket Boosters, and the main engine cutoff. The cutoff of the main engine marks the beginning of the second stage engine. After the completion of the second stage, the third stage engine ignites and then cuts off. Once the third stage engine cuts off spacecraft separation occurs.

XCALIBUR: a Vertical Takeoff TSTO RLV Concept with a HEDM Upperstage and a Scram-Rocket Booster

NASA Astrophysics Data System (ADS)

Bradford, J.

2002-01-01

A new 3rd generation, two-stage-to-orbit (TSTO) reusable launch vehicle (RLV) has been designed. The Xcalibur concept represents a novel approach due to its integration method for the upperstage element of the system. The vertical-takeoff booster, which is powered by rocket-based combined-cycle (RBCC) engines, carries the upperstage internally in the aft section of the airframe to a Mach 15 staging condition. The upperstage is released from the booster and carries the 6,820 kg of payload to low earth orbit (LEO) using its high energy density matter (HEDM) propulsion system. The booster element is capable of returning to the original launch site in a ramjet-cruise propulsion mode. Both the booster and the upperstage utilize advanced technologies including: graphite-epoxy tanks, metal-matrix composites, UHTC TPS materials, electro- mechanical actuators (EMAs), and lightweight subsystems (avionics, power distribution, etc.). The booster system is enabled main propulsion system which utilizes four RBCC engines. These engines operate in four distinct modes: air- augmented rocket (AAR), ramjet, scram-rocket, and all-rocket. The booster operates in AAR mode from takeoff to Mach 3, with ramjet mode operation from Mach 3 to Mach 6. The rocket re-ignition for scram-rocket mode occurs at Mach 6, with all-rocket mode from Mach 14 to the staging condition. The extended utilization of the scram-rocket mode greatly improves vehicle performance by providing superior vehicle acceleration when compared to the scramjet mode performance over the same flight region. Results indicate that the specific impulse penalty due to the scram-rocket mode operation is outweighed by the reduced flight time, smaller vehicle size due to increased mixture ratio, and lower allowable maximum dynamic pressure. A complete vehicle system life-cycle analysis was performed in an automated, multi-disciplinary design environment. Automated disciplinary performance analysis tools include: trajectory (POST), propulsion (SCCREAM), aeroheating (TCAT II), and an Excel spreadsheet for component weight estimation. These tools were automated using `file wrappers' in Phoenix Integration's ModelCenter collaborative design environment. Performance tools utilized for the analysis, but not requiring automation included IDEAS for solid modeling and APAS for the aerodynamic analysis. The paper describes the vehicle concept and operation, discussing the types of technologies used and the nominal flight scenario. A brief discussion explaining the decision-making process for the vehicle configuration is included. For cost predictions, NAFCOM-derived cost estimating relationships were used. Economic predictions were developed using a number of codes, including CABAM (financials), AATe (operations), and GTSafetyII (safety and reliability).

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.

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.

CFD Simulation of Liquid Rocket Engine Injectors

NASA Technical Reports Server (NTRS)

Farmer, Richard; Cheng, Gary; Chen, Yen-Sen; Garcia, Roberto (Technical Monitor)

2001-01-01

Detailed design issues associated with liquid rocket engine injectors and combustion chamber operation require CFD methodology which simulates highly three-dimensional, turbulent, vaporizing, and combusting flows. The primary utility of such simulations involves predicting multi-dimensional effects caused by specific injector configurations. SECA, Inc. and Engineering Sciences, Inc. have been developing appropriate computational methodology for NASA/MSFC for the past decade. CFD tools and computers have improved dramatically during this time period; however, the physical submodels used in these analyses must still remain relatively simple in order to produce useful results. Simulations of clustered coaxial and impinger injector elements for hydrogen and hydrocarbon fuels, which account for real fluid properties, is the immediate goal of this research. The spray combustion codes are based on the FDNS CFD code' and are structured to represent homogeneous and heterogeneous spray combustion. The homogeneous spray model treats the flow as a continuum of multi-phase, multicomponent fluids which move without thermal or velocity lags between the phases. Two heterogeneous models were developed: (1) a volume-of-fluid (VOF) model which represents the liquid core of coaxial or impinger jets and their atomization and vaporization, and (2) a Blob model which represents the injected streams as a cloud of droplets the size of the injector orifice which subsequently exhibit particle interaction, vaporization, and combustion. All of these spray models are computationally intensive, but this is unavoidable to accurately account for the complex physics and combustion which is to be predicted, Work is currently in progress to parallelize these codes to improve their computational efficiency. These spray combustion codes were used to simulate the three test cases which are the subject of the 2nd International Workshop on-Rocket Combustion Modeling. Such test cases are considered by these investigators to be very valuable for code validation because combustion kinetics, turbulence models and atomization models based on low pressure experiments of hydrogen air combustion do not adequately verify analytical or CFD submodels which are necessary to simulate rocket engine combustion. We wish to emphasize that the simulations which we prepared for this meeting are meant to test the accuracy of the approximations used in our general purpose spray combustion models, rather than represent a definitive analysis of each of the experiments which were conducted. Our goal is to accurately predict local temperatures and mixture ratios in rocket engines; hence predicting individual experiments is used only for code validation. To replace the conventional JANNAF standard axisymmetric finite-rate (TDK) computer code 2 for performance prediction with CFD cases, such codes must posses two features. Firstly, they must be as easy to use and of comparable run times for conventional performance predictions. Secondly, they must provide more detailed predictions of the flowfields near the injector face. Specifically, they must accurately predict the convective mixing of injected liquid propellants in terms of the injector element configurations.

The Viking Orbiter 1975 beryllium INTEREGEN rocket engine assembly.

NASA Technical Reports Server (NTRS)

Martinez, R. S.; Mcfarland, B. L.; Fischler, S.

1972-01-01

Description of the conversion of the Mariner 9 rocket engine for Viking Orbiter use. Engine conversion consists of replacing the 40:1 expansion area ratio nozzle with a 60:1 nozzle of the internal regeneratively (INTEREGEN) cooled rocket engine. Five converted engines using nitrogen tetroxide and monomethylhydrazine demonstrated thermal stability during the nominal 2730-sec burn, but experienced difficulty at operating extremes. The thermal stability characteristic was treated in two ways. The first treatment consisted of mapping the operating regime of the engine to determine its safest operating boundaries as regards thermal equilibrium. Six engines were used for this purpose. Two of the six engines were then modified to effect the second approach - i.e., extend the operating regime. The engines were modified by permitting fuel injection into the acoustic cavity.

Reliability evaluation methodology for NASA applications

NASA Technical Reports Server (NTRS)

Taneja, Vidya S.

1992-01-01

Liquid rocket engine technology has been characterized by the development of complex systems containing large number of subsystems, components, and parts. The trend to even larger and more complex system is continuing. The liquid rocket engineers have been focusing mainly on performance driven designs to increase payload delivery of a launch vehicle for a given mission. In otherwords, although the failure of a single inexpensive part or component may cause the failure of the system, reliability in general has not been considered as one of the system parameters like cost or performance. Up till now, quantification of reliability has not been a consideration during system design and development in the liquid rocket industry. Engineers and managers have long been aware of the fact that the reliability of the system increases during development, but no serious attempts have been made to quantify reliability. As a result, a method to quantify reliability during design and development is needed. This includes application of probabilistic models which utilize both engineering analysis and test data. Classical methods require the use of operating data for reliability demonstration. In contrast, the method described in this paper is based on similarity, analysis, and testing combined with Bayesian statistical analysis.

Monomethylhydrazine versus hydrazine fuels - Test results using a 100 pound thrust bipropellant rocket engine

NASA Technical Reports Server (NTRS)

Smith, J. A.; Stechman, R. C.

1981-01-01

A test program was performed to evaluate hydrazine (N2H4) as a fuel for a 445 Newton (100 lbf) thrust bipropellant rocket engine. Results of testing with an identical thruster utilizing monomethylhydrazine (MMH) are included for comparison. Engine performance with hydrazine fuel was essentially identical to that experienced with monomethylhydrazine although higher combustor wall temperatures (approximately 400 F) were obtained with hydrazine. Results are presented which indicate that hydrazine as a fuel is compatible with Marquardt bipropellant rocket engines which use monomethylhydrazine as a baseline fuel.

Video File - NASA on a Roll Testing Space Launch System Flight Engines

NASA Image and Video Library

2017-08-09

Just two weeks after conducting another in a series of tests on new RS-25 rocket engine flight controllers for NASA’s Space Launch System (SLS) rocket, engineers at NASA’s Stennis Space Center in Mississippi completed one more hot-fire test of a flight controller on August 9, 2017. With the hot fire, NASA has moved a step closer in completing testing on the four RS-25 engines which will power the first integrated flight of the SLS rocket and Orion capsule known as Exploration Mission 1.

Cryogenic gear technology for an orbital transfer vehicle engine and tester design

NASA Technical Reports Server (NTRS)

Calandra, M.; Duncan, G.

1986-01-01

Technology available for gears used in advanced Orbital Transfer Vehicle rocket engines and the design of a cryogenic adapted tester used for evaluating advanced gears are presented. The only high-speed, unlubricated gears currently in cryogenic service are used in the RL10 rocket engine turbomachinery. Advanced rocket engine gear systems experience operational load conditions and rotational speed that are beyond current experience levels. The work under this task consisted of a technology assessment and requirements definition followed by design of a self-contained portable cryogenic adapted gear test rig system.

Uncertainty Evaluation of Computational Model Used to Support the Integrated Powerhead Demonstration Project

NASA Technical Reports Server (NTRS)

Steele, W. G.; Molder, K. J.; Hudson, S. T.; Vadasy, K. V.; Rieder, P. T.; Giel, T.

2005-01-01

NASA and the U.S. Air Force are working on a joint project to develop a new hydrogen-fueled, full-flow, staged combustion rocket engine. The initial testing and modeling work for the Integrated Powerhead Demonstrator (IPD) project is being performed by NASA Marshall and Stennis Space Centers. A key factor in the testing of this engine is the ability to predict and measure the transient fluid flow during engine start and shutdown phases of operation. A model built by NASA Marshall in the ROCket Engine Transient Simulation (ROCETS) program is used to predict transient engine fluid flows. The model is initially calibrated to data from previous tests on the Stennis E1 test stand. The model is then used to predict the next run. Data from this run can then be used to recalibrate the model providing a tool to guide the test program in incremental steps to reduce the risk to the prototype engine. In this paper, they define this type of model as a calibrated model. This paper proposes a method to estimate the uncertainty of a model calibrated to a set of experimental test data. The method is similar to that used in the calibration of experiment instrumentation. For the IPD example used in this paper, the model uncertainty is determined for both LOX and LH flow rates using previous data. The successful use of this model is then demonstrated to predict another similar test run within the uncertainty bounds. The paper summarizes the uncertainty methodology when a model is continually recalibrated with new test data. The methodology is general and can be applied to other calibrated models.

THRSTER: A THRee-STream Ejector Ramjet Analysis and Design Tool

NASA Technical Reports Server (NTRS)

Chue, R. S.; Sabean, J.; Tyll, J.; Bakos, R. J.

2000-01-01

An engineering tool for analyzing ejectors in rocket based combined cycle (RBCC) engines has been developed. A key technology for multi-cycle RBCC propulsion systems is the ejector which functions as the compression stage of the ejector ramjet cycle. The THRee STream Ejector Ramjet analysis tool was developed to analyze the complex aerothermodynamic and combustion processes that occur in this device. The formulated model consists of three quasi-one-dimensional streams, one each for the ejector primary flow, the secondary flow, and the mixed region. The model space marches through the mixer, combustor, and nozzle to evaluate the solution along the engine. In its present form, the model is intended for an analysis mode in which the diffusion rates of the primary and secondary into the mixed stream are stipulated. The model offers the ability to analyze the highly two-dimensional ejector flowfield while still benefits from the simplicity and speed of an engineering tool. To validate the developed code, wall static pressure measurements from the Penn-State and NASA-ART RBCC experiments were used to compare with the results generated by the code. The calculated solutions were generally found to have satisfactory agreement with the pressure measurements along the engines, although further modeling effort may be required when a strong shock train is formed at the rocket exhaust. The range of parameters in which the code would generate valid results are presented and discussed.

THRSTER: A Three-Stream Ejector Ramjet Analysis and Design Tool

NASA Technical Reports Server (NTRS)

Chue, R. S.; Sabean, J.; Tyll, J.; Bakos, R. J.; Komar, D. R. (Technical Monitor)

2000-01-01

An engineering tool for analyzing ejectors in rocket based combined cycle (RBCC) engines has been developed. A key technology for multi-cycle RBCC propulsion systems is the ejector which functions as the compression stage of the ejector ramjet cycle. The THRee STream Ejector Ramjet analysis tool was developed to analyze the complex aerothermodynamic and combustion processes that occur in this device. The formulated model consists of three quasi-one-dimensional streams, one each for the ejector primary flow, the secondary flow, and the mixed region. The model space marches through the mixer, combustor, and nozzle to evaluate the solution along the engine. In its present form, the model is intended for an analysis mode in which the diffusion rates of the primary and secondary into the mixed stream are stipulated. The model offers the ability to analyze the highly two-dimensional ejector flowfield while still benefits from the simplicity and speed of an engineering tool. To validate the developed code, wall static pressure measurements from the Penn-State and NASA-ART RBCC experiments were used to compare with the results generated by the code. The calculated solutions were generally found to have satisfactory agreement with the pressure measurements along the engines, although further modeling effort may be required when a strong shock train is formed at the rocket exhaust. The range of parameters in which the code would generate valid results are presented and discussed.

Studies of an extensively axisymmetric rocket based combined cycle (RBCC) engine powered single-stage-to-orbit (SSTO) vehicle

NASA Technical Reports Server (NTRS)

Foster, Richard W.; Escher, William J. D.; Robinson, John W.

1989-01-01

The present comparative performance study has established that rocket-based combined cycle (RBCC) propulsion systems, when incorporated by essentially axisymmetric SSTO launch vehicle configurations whose conical forebody maximizes both capture-area ratio and total capture area, are capable of furnishing payload-delivery capabilities superior to those of most multistage, all-rocket launchers. Airbreathing thrust augmentation in the rocket-ejector mode of an RBCC powerplant is noted to make a major contribution to final payload capability, by comparison to nonair-augmented rocket engine propulsion systems.

Use of Soft Computing Technologies For Rocket Engine Control

NASA Technical Reports Server (NTRS)

Trevino, Luis C.; Olcmen, Semih; Polites, Michael

2003-01-01

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

Flow Separation Side Loads Excitation of Rocket Nozzle FEM

NASA Technical Reports Server (NTRS)

Smalley, Kurt B.; Brown, Andrew; Ruf, Joseph; Gilbert, John

2007-01-01

Modern rocket nozzles are designed to operate over a wide range of altitudes, and are also built with large aspect ratios to enable high efficiencies. Nozzles designed to operate over specific regions of a trajectory are being replaced in modern launch vehicles by those that are designed to operate from earth to orbit. This is happening in parallel with modern manufacturing and wall cooling techniques allowing for larger aspect ratio nozzles to be produced. Such nozzles, though operating over a large range of altitudes and ambient pressures, are typically designed for one specific altitude. Above that altitude the nozzle flow is 'underexpanded' and below that altitude, the nozzle flow is 'overexpanded'. In both conditions the nozzle produces less than the maximum possible thrust at that altitude. Usually the nozzle design altitude is well above sea level, leaving the nozzle flow in an overexpanded state for its start up as well as for its ground testing where, if it is a reusable nozzle such as the Space Shuttle Main Engine (SSME), the nozzle will operate for the majority of its life. Overexpansion in a rocket nozzle presents the critical, and sometimes design driving, problem of flow separation induced side loads. To increase their understanding of nozzle side loads, engineers at MSFC began an investigation in 2000 into the phenomenon through a task entitled "Characterization and Accurate Modeling of Rocket Engine Nozzle Side Loads", led by A. Brown. The stated objective of this study was to develop a methodology to accurately predict the character and magnitude of nozzle side loads. The study included further hot-fire testing of the MC-l engine, cold flow testing of subscale nozzles, CFD analyses of both hot-fire and cold flow nozzle testing, and finite element (fe.) analysis of the MC-1 engine and cold flow tested nozzles. A follow on task included an effort to formulate a simplified methodology for modeling a side load during a two nodal diameter fluid/structure interaction for a single moment in time.

Turbulent Mixing of Primary and Secondary Flow Streams in a Rocket-Based Combined Cycle Engine

NASA Technical Reports Server (NTRS)

Cramer, J. M.; Greene, M. U.; Pal, S.; Santoro, R. J.; Turner, Jim (Technical Monitor)

2002-01-01

This viewgraph presentation gives an overview of the turbulent mixing of primary and secondary flow streams in a rocket-based combined cycle (RBCC) engine. A significant RBCC ejector mode database has been generated, detailing single and twin thruster configurations and global and local measurements. On-going analysis and correlation efforts include Marshall Space Flight Center computational fluid dynamics modeling and turbulent shear layer analysis. Potential follow-on activities include detailed measurements of air flow static pressure and velocity profiles, investigations into other thruster spacing configurations, performing a fundamental shear layer mixing study, and demonstrating single-shot Raman measurements.

Damage-mitigating control of space propulsion systems for high performance and extended life

NASA Technical Reports Server (NTRS)

Ray, Asok; Wu, Min-Kuang; Dai, Xiaowen; Carpino, Marc; Lorenzo, Carl F.

1993-01-01

Calculations are presented showing that a substantial improvement in service life of a reusable rocket engine can be achieved by an insignificant reduction in the system dynamic performance. The paper introduces the concept of damage mitigation and formulates a continuous-time model of fatigue damage dynamics. For control of complex mechanical systems, damage prediction and damage mitigation are carried out based on the available sensory and operational information such that the plant can be inexpensively maintained and safely and efficiently steered under diverse operating conditions. The results of simulation experiments are presented for transient operations of a reusable rocket engine.

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.

KSC-2013-4343

NASA Image and Video Library

2013-12-11

CAPE CANAVERAL, Fla. -- At NASA's Kennedy Space Center in Florida, from the left, Leandro James, rocket avionics lead, and Julio Najarro of Mechanical Systems make final adjustments to a small rocket prior to launch as part of Rocket University. The launch will test systems designed by the student engineers. As part of Rocket University, the engineers are given an opportunity to work a fast-track project to develop skills in developing spacecraft systems of the future. As NASA plans for future spaceflight programs to low-Earth orbit and beyond, teams of engineers at Kennedy are gaining experience in designing and flying launch vehicle systems on a small scale. Four teams of five to eight members from Kennedy are designing rockets complete with avionics and recovery systems. Launch operations require coordination with federal agencies, just as they would with rockets launched in support of a NASA mission. Photo credit: NASA/Jim Grossmann

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.

A computer model for liquid jet atomization in rocket thrust chambers

NASA Astrophysics Data System (ADS)

Giridharan, M. G.; Lee, J. G.; Krishnan, A.; Yang, H. Q.; Ibrahim, E.; Chuech, S.; Przekwas, A. J.

1991-12-01

The process of atomization has been used as an efficient means of burning liquid fuels in rocket engines, gas turbine engines, internal combustion engines, and industrial furnaces. Despite its widespread application, this complex hydrodynamic phenomenon has not been well understood, and predictive models for this process are still in their infancy. The difficulty in simulating the atomization process arises from the relatively large number of parameters that influence it, including the details of the injector geometry, liquid and gas turbulence, and the operating conditions. In this study, numerical models are developed from first principles, to quantify factors influencing atomization. For example, the surface wave dynamics theory is used for modeling the primary atomization and the droplet energy conservation principle is applied for modeling the secondary atomization. The use of empirical correlations has been minimized by shifting the analyses to fundamental levels. During applications of these models, parametric studies are performed to understand and correlate the influence of relevant parameters on the atomization process. The predictions of these models are compared with existing experimental data. The main tasks of this study were the following: development of a primary atomization model; development of a secondary atomization model; development of a model for impinging jets; development of a model for swirling jets; and coupling of the primary atomization model with a CFD code.

What is the Relationship Between Altitude and Weight in a Model Rocket?

NASA Technical Reports Server (NTRS)

Betz, Jonathan

1993-01-01

This experiment was designed to find a function of payload weight for altitude. The same rocket was launched a repeated number of times with the same engine and varying amounts of weight. After performing experimentation, it was calculated that the altitude in meters could be predicted with the equation A = (2.8(W exp 2)) - (70.6W + 310.3), with weight expressed in the unit ounces.

50 Years of Testing

NASA Image and Video Library

2016-04-23

A 15-second test of a Saturn V rocket stage on the A-2 Test Stand at Stennis Space Center ushered in the Space Age for south Mississippi. Fifty years later, Stennis has grown into the nation’s largest rocket engine test site, continuing to test rocket engines and stages that power the nation’s space program.

Rockets -- Part II.

ERIC Educational Resources Information Center

Leitner, Alfred

1982-01-01

If two rockets are identical except that one engine burns in one-tenth the time of the other (total impulse and initial fuel mass of the two engines being the same), which rocket will rise higher? Why? The answer to this question (part 1 response in v20 n6, p410, Sep 1982) is provided. (Author/JN)

Computer Design Technology of the Small Thrust Rocket Engines Using CAE / CAD Systems

NASA Astrophysics Data System (ADS)

Ryzhkov, V.; Lapshin, E.

2018-01-01

The paper presents an algorithm for designing liquid small thrust rocket engine, the process of which consists of five aggregated stages with feedback. Three stages of the algorithm provide engineering support for design, and two stages - the actual engine design. A distinctive feature of the proposed approach is a deep study of the main technical solutions at the stage of engineering analysis and interaction with the created knowledge (data) base, which accelerates the process and provides enhanced design quality. The using multifunctional graphic package Siemens NX allows to obtain the final product -rocket engine and a set of design documentation in a fairly short time; the engine design does not require a long experimental development.

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.

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.

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.

Unsteady response of flow system around balance piston in a rocket pump

NASA Astrophysics Data System (ADS)

Kawasaki, S.; Shimura, T.; Uchiumi, M.; Hayashi, M.; Matsui, J.

2013-03-01

In the rocket engine turbopump, a self-balancing type of axial thrust balancing system using a balance piston is often applied. In this study, the balancing system in liquid-hydrogen (LH2) rocket pump was modeled combining the mechanical structure and the flow system, and the unsteady response of the balance piston was investigated. The axial vibration characteristics of the balance piston with a large amplitude were determined, sweeping the frequency of the pressure fluctuation on the inlet of the balance piston. This vibration was significantly affected by the compressibility of LH2.

Base pressure and heat transfer tests of the 0.0225-scale space shuttle plume simulation model (19-OTS) in yawed flight conditions in the NASA-Lewis 10x10-foot supersonic wind tunnel (test IH83)

NASA Technical Reports Server (NTRS)

Foust, J. W.

1979-01-01

Wind tunnel tests were performed to determine pressures, heat transfer rates, and gas recovery temperatures in the base region of a rocket firing model of the space shuttle integrated vehicle during simulated yawed flight conditions. First and second stage flight of the space shuttle were simulated by firing the main engines in conjunction with the SRB rocket motors or only the SSME's into the continuous tunnel airstream. For the correct rocket plume environment, the simulated altitude pressures were halved to maintain the rocket chamber/altitude pressure ratio. Tunnel freestream Mach numbers from 2.2 to 3.5 were simulated over an altitude range of 60 to 130 thousand feet with varying angle of attack, yaw angle, nozzle gimbal angle and SRB chamber pressure. Gas recovery temperature data derived from nine gas temperature probe runs are presented. The model configuration, instrumentation, test procedures, and data reduction are described.

Parametric Study Conducted of Rocket- Based, Combined-Cycle Nozzles

NASA Technical Reports Server (NTRS)

Steffen, Christopher J., Jr.; Smith, Timothy D.

1998-01-01

Having reached the end of the 20th century, our society is quite familiar with the many benefits of recycling and reusing the products of civilization. The high-technology world of aerospace vehicle design is no exception. Because of the many potential economic benefits of reusable launch vehicles, NASA is aggressively pursuing this technology on several fronts. One of the most promising technologies receiving renewed attention is Rocket-Based, Combined-Cycle (RBCC) propulsion. This propulsion method combines many of the efficiencies of high-performance jet aircraft with the power and high-altitude capability of rocket engines. The goal of the present work at the NASA Lewis Research Center is to further understand the complex fluid physics within RBCC engines that govern system performance. This work is being performed in support of NASA's Advanced Reusable Technologies program. A robust RBCC engine design optimization demands further investigation of the subsystem performance of the engine's complex propulsion cycles. The RBCC propulsion system under consideration at Lewis is defined by four modes of operation in a singlestage- to-orbit configuration. In the first mode, the engine functions as a rocket-driven ejector. When the rocket engine is switched off, subsonic combustion (mode 2) is present in the ramjet mode. As the vehicle continues to accelerate, supersonic combustion (mode 3) occurs in the ramjet mode. Finally, as the edge of the atmosphere is approached and the engine inlet is closed off, the rocket is reignited and the final accent to orbit is undertaken in an all-rocket mode (mode 4). The performance of this fourth and final mode is the subject of this present study. Performance is being monitored in terms of the amount of thrust generated from a given amount of propellant.

Current Development of Nuclear Thermal Propulsion technologies at the Center for Space Nuclear Research

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

Robert C. O'Brien; Steven K. Cook; Nathan D. Jerred

Nuclear power and propulsion has been considered for space applications since the 1950s. Between 1955 and 1972 the US built and tested over twenty nuclear reactors / rocket engines in the Rover/NERVA programs1. The Aerojet Corporation was the prime contractor for the NERVA program. Modern changes in environmental laws present challenges for the redevelopment of the nuclear rocket. Recent advances in fuel fabrication and testing options indicate that a nuclear rocket with a fuel composition that is significantly different from those of the NERVA project can be engineered; this may be needed to ensure public support and compliance with safetymore » requirements. The Center for Space Nuclear Research (CSNR) is pursuing a number of technologies, modeling and testing processes to further the development of safe, practical and affordable nuclear thermal propulsion systems.« less

The use of programmable logic controllers (PLC) for rocket engine component testing

NASA Technical Reports Server (NTRS)

Nail, William; Scheuermann, Patrick; Witcher, Kern

1991-01-01

Application of PLCs to the rocket engine component testing at a new Stennis Space Center Component Test Facility is suggested as an alternative to dedicated specialized computers. The PLC systems are characterized by rugged design, intuitive software, fault tolerance, flexibility, multiple end device options, networking capability, and built-in diagnostics. A distributed PLC-based system is projected to be used for testing LH2/LOx turbopumps required for the ALS/NLS rocket engines.

Navier-Stokes analysis of a liquid rocket engine disk cavity

NASA Technical Reports Server (NTRS)

Benjamin, Theodore G.; Mcconnaughey, Paul K.

1991-01-01

This paper presents a Navier-Stokes analysis of hydrodynamic phenomena occurring in the aft disk cavity of a liquid rocket engine turbine. The cavity analyzed in the Space Shuttle Main Engine Alternate Turbopump currently being developed by NASA and Pratt and Whitney. Comparison of results obtained from the Navier-Stokes code for two rotating disk datasets available in the literature are presented as benchmark validations. The benchmark results obtained using the code show good agreement relative to experimental data, and the turbine disk cavity was analyzed with comparable grid resolution, dissipation levels, and turbulence models. Predicted temperatures in the cavity show that little mixing of hot and cold fluid occurs in the cavity and the flow is dominated by swirl and pumping up the rotating disk.

Linear quadratic servo control of a reusable rocket engine

NASA Technical Reports Server (NTRS)

Musgrave, Jeffrey L.

1991-01-01

The paper deals with the development of a design method for a servo component in the frequency domain using singular values and its application to a reusable rocket engine. A general methodology used to design a class of linear multivariable controllers for intelligent control systems is presented. Focus is placed on performance and robustness characteristics, and an estimator design performed in the framework of the Kalman-filter formalism with emphasis on using a sensor set different from the commanded values is discussed. It is noted that loop transfer recovery modifies the nominal plant noise intensities in order to obtain the desired degree of robustness to uncertainty reflected at the plant input. Simulation results demonstrating the performance of the linear design on a nonlinear engine model over all power levels during mainstage operation are discussed.

Evaluation of an Ejector Ramjet Based Propulsion System for Air-Breathing Hypersonic Flight

NASA Technical Reports Server (NTRS)

Thomas, Scott R.; Perkins, H. Douglas; Trefny, Charles J.

1997-01-01

A Rocket Based Combined Cycle (RBCC) engine system is designed to combine the high thrust to weight ratio of a rocket along with the high specific impulse of a ramjet in a single, integrated propulsion system. This integrated, combined cycle propulsion system is designed to provide higher vehicle performance than that achievable with a separate rocket and ramjet. The RBCC engine system studied in the current program is the Aerojet strutjet engine concept, which is being developed jointly by a government-industry team as part of the Air Force HyTech program pre-PRDA activity. The strutjet is an ejector-ramjet engine in which small rocket chambers are embedded into the trailing edges of the inlet compression struts. The engine operates as an ejector-ramjet from takeoff to slightly above Mach 3. Above Mach 3 the engine operates as a ramjet and transitions to a scramjet at high Mach numbers. For space launch applications the rockets would be re-ignited at a Mach number or altitude beyond which air-breathing propulsion alone becomes impractical. The focus of the present study is to develop and demonstrate a strutjet flowpath using hydrocarbon fuel at up to Mach 7 conditions.

Fiber-reinforced ceramic composites for Earth-to-orbit rocket engine turbines

NASA Technical Reports Server (NTRS)

Brockmeyer, Jerry W.; Schnittgrund, Gary D.

1990-01-01

Fiber reinforced ceramic matrix composites (FRCMC) are emerging materials systems that offer potential for use in liquid rocket engines. Advantages of these materials in rocket engine turbomachinery include performance gain due to higher turbine inlet temperature, reduced launch costs, reduced maintenance with associated cost benefits, and reduced weight. This program was initiated to assess the state of FRCMC development and to propose a plan for their implementation into liquid rocket engine turbomachinery. A complete range of FRCMC materials was investigated relative to their development status and feasibility for use in the hot gas path of earth-to-orbit rocket engine turbomachinery. Of the candidate systems, carbon fiber-reinforced silicon carbide (C/SiC) offers the greatest near-term potential. Critical hot gas path components were identified, and the first stage inlet nozzle and turbine rotor of the fuel turbopump for the liquid oxygen/hydrogen Space Transportation Main Engine (STME) were selected for conceptual design and analysis. The critical issues associated with the use of FRCMC were identified. Turbine blades were designed, analyzed and fabricated. The Technology Development Plan, completed as Task 5 of this program, provides a course of action for resolution of these issues.

Theoretical Acoustic Absorber Design Approach for LOX/LCH4 Pintle Injector Rocket Engines

NASA Astrophysics Data System (ADS)

Candelaria, Jonathan

Liquid rocket engines, or LREs, have served a key role in space exploration efforts. One current effort involves the utilization of liquid oxygen (LOX) and liquid methane (LCH4) LREs to explore Mars with in-situ resource utilization for propellant production. This on-site production of propellant will allow for greater payload allocation instead of fuel to travel to the Mars surface, and refueling of propellants to travel back to Earth. More useable mass yields a greater benefit to cost ratio. The University of Texas at El Paso's (UTEP) Center for Space Exploration and Technology Research Center (cSETR) aims to further advance these methane propulsion systems with the development of two liquid methane - liquid oxygen propellant combination rocket engines. The design of rocket engines, specifically liquid rocket engines, is complex in that many variables are present that must be taken into consideration in the design. A problem that occurs in almost every rocket engine development program is combustion instability, or oscillatory combustion. It can result in the destruction of the rocket, subsequent destruction of the vehicle and compromise the mission. These combustion oscillations can vary in frequency from 100 to 20,000 Hz or more, with varying effects, and occur from different coupling phenomena. It is important to understand the effects of combustion instability, its physical manifestations, how to identify the instabilities, and how to mitigate or dampen them. Linear theory methods have been developed to provide a mathematical understanding of the low- to mid-range instabilities. Nonlinear theory is more complex and difficult to analyze mathematically, therefore no general analytical method that yields a solution exists. With limited resources, time, and the advice of our NASA mentors, a data driven experimental approach utilizing quarter wave acoustic dampener cavities was designed. This thesis outlines the methodology behind the design of an acoustic dampening system for a 500 lbf and a 2000 lbf throttleable liquid oxygen liquid methane pintle injector rocket engine.

Overview of Engineering Design and Analysis at the NASA John C. Stennis Space Center

NASA Technical Reports Server (NTRS)

Ryan, Harry; Congiardo, Jared; Junell, Justin; Kirkpatrick, Richard

2007-01-01

A wide range of rocket propulsion test work occurs at the NASA John C. Stennis Space Center (SSC) including full-scale engine test activities at test facilities A-1, A-2, B-1 and B-2 as well as combustion device research and development activities at the E-Complex (E-1, E-2, E-3 and E-4) test facilities. The propulsion test engineer at NASA SSC faces many challenges associated with designing and operating a test facility due to the extreme operating conditions (e.g., cryogenic temperatures, high pressures) of the various system components and the uniqueness of many of the components and systems. The purpose of this paper is to briefly describe the NASA SSC Engineering Science Directorate s design and analysis processes, experience, and modeling techniques that are used to design and support the operation of unique rocket propulsion test facilities.

Apollo Contour Rocket Nozzle in the Propulsion Systems Laboratory

NASA Image and Video Library

1964-07-21

Bill Harrison and Bud Meilander check the setup of an Apollo Contour rocket nozzle in the Propulsion Systems Laboratory at the National Aeronautics and Space Administration (NASA) Lewis Research Center. The Propulsion Systems Laboratory contained two 14-foot diameter test chambers that could simulate conditions found at very high altitudes. The facility was used in the 1960s to study complex rocket engines such as the Pratt and Whitney RL-10 and rocket components such as the Apollo Contour nozzle, seen here. Meilander oversaw the facility’s mechanics and the installation of test articles into the chambers. Harrison was head of the Supersonic Tunnels Branch in the Test Installations Division. Researchers sought to determine the impulse value of the storable propellant mix, classify and improve the internal engine performance, and compare the results with analytical tools. A special setup was installed in the chamber that included a device to measure the thrust load and a calibration stand. Both cylindrical and conical combustion chambers were examined with the conical large area ratio nozzles. In addition, two contour nozzles were tested, one based on the Apollo Service Propulsion System and the other on the Air Force’s Titan transtage engine. Three types of injectors were investigated, including a Lewis-designed model that produced 98-percent efficiency. It was determined that combustion instability did not affect the nozzle performance. Although much valuable information was obtained during the tests, attempts to improve the engine performance were not successful.

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.

FNAS/summer faculty fellowship research continuation program. Task 6: Integrated model development for liquid fueled rocket propulsion systems. Task 9: Aspects of model-based rocket engine condition monitoring and control

NASA Technical Reports Server (NTRS)

Santi, L. Michael; Helmicki, Arthur J.

1993-01-01

The objective of Phase I of this research effort was to develop an advanced mathematical-empirical model of SSME steady-state performance. Task 6 of Phase I is to develop component specific modification strategy for baseline case influence coefficient matrices. This report describes the background of SSME performance characteristics and provides a description of the control variable basis of three different gains models. The procedure used to establish influence coefficients for each of these three models is also described. Gains model analysis results are compared to Rocketdyne's power balance model (PBM).

Engineers demonstrate the pocket rocket

NASA Technical Reports Server (NTRS)

1996-01-01

Part of Stennis Space Center's mission with its traveling exhibits is to educate the younger generation on how propulsion systems work. A popular tool is the 'pocket rocket,' which demonstrates how a hybrid rocket works. A hybrid rocket is a cross breed between a solid fuel rocket and a liquid fuel rocket.

Rocket engine numerical simulator

NASA Technical Reports Server (NTRS)

Davidian, Ken

1993-01-01

The topics are presented in viewgraph form and include the following: a rocket engine numerical simulator (RENS) definition; objectives; justification; approach; potential applications; potential users; RENS work flowchart; RENS prototype; and conclusion.

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.

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.

Rocket Based Combined Cycle (RBCC) engine inlet

NASA Technical Reports Server (NTRS)

2004-01-01

Pictured is a component of the Rocket Based Combined Cycle (RBCC) engine. This engine was designed to ultimately serve as the near term basis for Two Stage to Orbit (TSTO) air breathing propulsion systems and ultimately a Single Stage to Orbit (SSTO) air breathing propulsion system.

AJ26 engine testing moves forward

NASA Image and Video Library

2010-07-19

Stennis employees at the E-1 Test Stand position an Aerojet AJ26 rocket engine in preparation for a series of early tests. Stennis has partnered with Orbital Sciences Corporation to test the rocket engine for the company's commercial cargo flights to the International Space Station.

Rainbows and Rocket Engine

NASA Image and Video Library

2017-02-22

Rainbows and rocket engines – doesn’t get much better than that! Check out these gorgeous aerial views from today’s Space Launch System RS-25 engine test @NASA’s Stennis Space Center. PAO Name:Kim Henry Phone Number:256-544-1899 Email Address: kimberly.m.henry@nasa.gov

Options for flight testing rocket-based combined-cycle (RBCC) engines

NASA Technical Reports Server (NTRS)

Olds, John

1996-01-01

While NASA's current next-generation launch vehicle research has largely focused on advanced all-rocket single-stage-to-orbit vehicles (i.e. the X-33 and it's RLV operational follow-on), some attention is being given to advanced propulsion concepts suitable for 'next-generation-and-a-half' vehicles. Rocket-based combined-cycle (RBCC) engines combining rocket and airbreathing elements are one candidate concept. Preliminary RBCC engine development was undertaken by the United States in the 1960's. However, additional ground and flight research is required to bring the engine to technological maturity. This paper presents two options for flight testing early versions of the RBCC ejector scramjet engine. The first option mounts a single RBCC engine module to the X-34 air-launched technology testbed for test flights up to about Mach 6.4. The second option links RBCC engine testing to the simultaneous development of a small-payload (220 lb.) two-stage-to-orbit operational vehicle in the Bantam payload class. This launcher/testbed concept has been dubbed the W vehicle. The W vehicle can also serve as an early ejector ramjet RBCC launcher (albeit at a lower payload). To complement current RBCC ground testing efforts, both flight test engines will use earth-storable propellants for their RBCC rocket primaries and hydrocarbon fuel for their airbreathing modes. Performance and vehicle sizing results are presented for both options.

CLOSEUP VIEW OF THE FIRST STAGE OF THE SATURN I ...

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

CLOSE-UP VIEW OF THE FIRST STAGE OF THE SATURN I ROCKET, SHOWING A DETAIL VIEW OF THE ENGINE CLUSTER. THE SATURN I ROCKET WAS THE FIRST UNITED STATES ROCKET TO HAVE MULTIPLE ENGINES ON A SINGLE STAGE. - Marshall Space Flight Center, Saturn Propulsion & Structural Test Facility, East Test Area, Huntsville, Madison County, AL

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.

Smoke and fire Rocket-engine ablaze on This Week @NASA – August 14, 2015

NASA Image and Video Library

2015-08-14

On Aug. 13, NASA conducted a test firing of the RS-25 rocket engine at Stennis Space Center. The 535 second test was the sixth in the current series of seven developmental tests of the former space shuttle main engine. Four RS-25 engines will power the core stage of the new Space Launch System (SLS) rocket, which will carry humans deeper into space than ever before, including to an asteroid and Mars. Also, Veggies in space, Russian spacewalk, Supply ship undocks from ISS, Smallest giant black hole, 10th anniversary of MRO launch and more!

Space shuttle hypergolic bipropellant RCS engine design study, Bell model 8701

NASA Technical Reports Server (NTRS)

1974-01-01

A research program was conducted to define the level of the current technology base for reaction control system rocket engines suitable for space shuttle applications. The project consisted of engine analyses, design, fabrication, and tests. The specific objectives are: (1) extrapolating current engine design experience to design of an RCS engine with required safety, reliability, performance, and operational capability, (2) demonstration of multiple reuse capability, and (3) identification of current design and technology deficiencies and critical areas for future effort.

Focused Experimental and Analytical Studies of the RBCC Rocket-Ejector

NASA Technical Reports Server (NTRS)

Lehman, M.; Pal, S.; Schwes, D.; Chen, J. D.; Santoro, R. J.

1999-01-01

The rocket-ejector mode of a Rocket Based Combined Cycle Engine (RBCC) was studied through a joint experimental/analytical approach. A two-dimensional variable geometry rocket-ejector system with enhanced optical access was designed and fabricated for experimentation. The rocket-ejector system utilizes a single two-dimensional gaseous oxygen/gaseous hydrogen rocket as the ejector. 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 configurations for a range of rocket operating conditions Overall system performance was obtained through Global measurements of wall static pressure profiles, heat flux profiles and engine thrust, whereas detailed mixing and combustion information was obtained through Raman spectroscopy measurements of major species (gaseous oxygen, hydrogen. nitrogen and water vapor). These experimental efforts were complemented by Computational Fluid Dynamic (CFD) flowfield analyses.

A reusable rocket engine intelligen control

NASA Technical Reports Server (NTRS)

Merrill, Walter C.; Lorenzo, Carl F.

1988-01-01

An intelligent control system for reusable space propulsion systems for future launch vehicles is described. The system description includes a framework for the design. The framework consists of an execution level with high-speed control and diagnostics, and a coordination level which marries expert system concepts with traditional control. A comparison is made between air breathing and rocket engine control concepts to assess the relative levels of development and to determine the applicability of air breathing control concepts to future reusable rocket engine systems.

A reusable rocket engine intelligent control

NASA Technical Reports Server (NTRS)

Merrill, Walter C.; Lorenzo, Carl F.

1988-01-01

An intelligent control system for reusable space propulsion systems for future launch vehicles is described. The system description includes a framework for the design. The framework consists of an execution level with high-speed control and diagnostics, and a coordination level which marries expert system concepts with traditional control. A comparison is made between air breathing and rocket engine control concepts to assess the relative levels of development and to determine the applicability of air breathing control concepts ot future reusable rocket engine systems.

Fiber-Reinforced Superalloys For Rocket Engines

NASA Technical Reports Server (NTRS)

Lewis, Jack R.; Yuen, Jim L.; Petrasek, Donald W.; Stephens, Joseph R.

1990-01-01

Report discusses experimental studies of fiber-reinforced superalloy (FRS) composite materials for use in turbine blades in rocket engines. Intended to withstand extreme conditions of high temperature, thermal shock, atmospheres containing hydrogen, high cycle fatigue loading, and thermal fatigue, which tax capabilities of even most-advanced current blade material - directionally-solidified, hafnium-modified MAR M-246 {MAR M-246 (Hf) (DS)}. FRS composites attractive combination of properties for use in turbopump blades of advanced rocket engines at temperatures from 870 to 1,100 degrees C.

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.

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.

Andy Hardin with 3-D printed engine part

NASA Image and Video Library

2015-06-22

ANDY HARDIN, A PROPULSION ENGINEER AT NASA'S MARSHALL SPACE FLIGHT CENTER IN HUNTSVILLE, ALABAMA, SHOWS A 3-D PRINTED ROCKET PART MADE WITH A SELECTIVE LASER MELTING MACHINE. PARTS FOR THE SPACE LAUNCH SYSTEM'S RS-25 ROCKET ENGINE ARE BEING MADE WITH THE MACHINE IN THE BACKGROUND

Space Storable Rocket Technology (SSRT) basic program

NASA Technical Reports Server (NTRS)

Chazen, M. L.; Mueller, T.; Casillas, A. R.; Huang, D.

1992-01-01

The Space Storable Rocket Technology Program (SSRT) was conducted to establish a technology for a new class of high performance and long life bipropellant engines using space storable propellants. The results are described. Task 1 evaluated several characteristics for a number of fuels to determine the best space storable fuel for use with LO2. The results indicated that LO2-N2H4 is the best propellant combination and provides the maximum mission/system capability maximum payload into GEO of satellites. Task 2 developed two models, performance and thermal. The performance model indicated the performance goal of specific impulse greater than or = 340 seconds (sigma = 204) could be achieved. The thermal model was developed and anchored to hot fire test data. Task 3 consisted of design, fabrication, and testing of a 200 lbf thrust test engine operating at a chamber pressure of 200 psia using LO2-N2H4. A total of 76 hot fire tests were conducted demonstrating performance greater than 340 (sigma = 204) which is a 25 second specific impulse improvement over the existing highest performance flight apogee type engines.

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.

Hybrid Rocket Performance Prediction with Coupling Method of CFD and Thermal Conduction Calculation

NASA Astrophysics Data System (ADS)

Funami, Yuki; Shimada, Toru

The final purpose of this study is to develop a design tool for hybrid rocket engines. This tool is a computer code which will be used in order to investigate rocket performance characteristics and unsteady phenomena lasting through the burning time, such as fuel regression or combustion oscillation. When phenomena inside a combustion chamber, namely boundary layer combustion, are described, it is difficult to use rigorous models for this target. It is because calculation cost may be too expensive. Therefore simple models are required for this calculation. In this study, quasi-one-dimensional compressible Euler equations for flowfields inside a chamber and the equation for thermal conduction inside a solid fuel are numerically solved. The energy balance equation at the solid fuel surface is solved to estimate fuel regression rate. Heat feedback model is Karabeyoglu's model dependent on total mass flux. Combustion model is global single step reaction model for 4 chemical species or chemical equilibrium model for 9 chemical species. As a first step, steady-state solutions are reported.

Two-Dimensional Motions of Rockets

ERIC Educational Resources Information Center

Kang, Yoonhwan; Bae, Saebyok

2007-01-01

We analyse the two-dimensional motions of the rockets for various types of rocket thrusts, the air friction and the gravitation by using a suitable representation of the rocket equation and the numerical calculation. The slope shapes of the rocket trajectories are discussed for the three types of rocket engines. Unlike the projectile motions, the…

Rocket engine numerical simulation

NASA Astrophysics Data System (ADS)

Davidian, Ken

1993-12-01

The topics are presented in view graph form and include the following: a definition of the rocket engine numerical simulator (RENS); objectives; justification; approach; potential applications; potential users; RENS work flowchart; RENS prototype; and conclusions.

Rocket engine numerical simulation

NASA Technical Reports Server (NTRS)

Davidian, Ken

1993-01-01

The topics are presented in view graph form and include the following: a definition of the rocket engine numerical simulator (RENS); objectives; justification; approach; potential applications; potential users; RENS work flowchart; RENS prototype; and conclusions.

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.

NASA Marches on with Test of RS-25 Engine for New Space Launch System

NASA Image and Video Library

2016-07-29

NASA engineers conducted a successful developmental test of RS-25 rocket engine No. 0528 July 29, 2016, to collect critical performance data for the most powerful rocket in the world – the Space Launch System (SLS). The engine roared to life for a full 650-second test on the A-1 Test Stand at NASA’s Stennis Space Center, near Bay St. Louis, Mississippi, marking another step forward in development of the SLS, which will launch humans deeper into space than ever before, including on the journey to Mars. Four RS-25 engines, joined with a pair of solid rocket boosters, will power the SLS core stage at launch. The RS-25 engines used on the first four SLS flights are former space shuttle main engines, modified to operate at a higher performance level and with a new engine controller, which allows communication between the vehicle and engine.

Electrodynamic actuators for rocket engine valves

NASA Technical Reports Server (NTRS)

Fiet, O.; Doshi, D.

1972-01-01

Actuators, employed in acoustic loudspeakers, operate liquid rocket engine valves by replacing light paper cones with flexible metal diaphragms. Comparative analysis indicates better response time than solenoid actuators, and improved service life and reliability.

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.

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.

Measurements of Unexpected Ozone Loss in a Nighttime Space Shuttle Exhaust Plume: Implications for Geo-Engineering Projects

NASA Astrophysics Data System (ADS)

Avallone, L. M.; Kalnajs, L. E.; Toohey, D. W.; Ross, M. N.

2008-12-01

Measurements of ozone, carbon dioxide and particulate water were made in the nighttime exhaust plume of the Space Shuttle (STS-116) on 9 December 2006 as part of the PUMA/WAVE campaign (Plume Ultrafast Measurements Acquisition/WB-57F Ascent Video Experiment). The launch took place from Kennedy Space Center at 8:47 pm (local time) on a moonless night and the WB-57F aircraft penetrated the shuttle plume approximately 25 minutes after launch in the lowermost stratosphere. Ozone loss is not predicted to occur in a nighttime Space Shuttle plume since it has long been assumed that the main ozone loss mechanism associated with rocket emissions requires solar photolysis to drive several chlorine-based catalytic cycles. However, the nighttime in situ observations show an unexpected loss of ozone of approximately 250 ppb in the evolving exhaust plume, inconsistent with model predictions. We will present the observations of the shuttle exhaust plume composition and the results of photochemical models of the Space Shuttle plume. We will show that models constrained by known rocket emission kinetics, including afterburning, and reasonable plume dispersion rates, based on the CO2 observations, cannot explain the observed ozone loss. We will propose potential explanations for the lack of agreement between models and the observations, and will discuss the implications of these explanations for our understanding of the composition of rocket emissions. We will describe the potential consequences of the observed ozone loss for long-term damage to the stratospheric ozone layer should geo-engineering projects based on rocket launches be employed.

Technicians Manufacture a Nozzle for the Kiwi B-1-B Engine

NASA Image and Video Library

1964-05-21

Technicians manufacture a nozzle for the Kiwi B-1-B nuclear rocket engine in the Fabrication Shop’s vacuum oven at the National Aeronautics and Space Administration (NASA) Lewis Research Center. 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 basic nuclear rocket principles in a non-flying nuclear engine. The next phase, NERVA, would create an entire flyable engine. The final phase of the program, called Reactor-In-Flight-Test, would be an actual launch test. The AEC was responsible for designing the nuclear reactor and overall engine. NASA Lewis was responsible for developing the liquid-hydrogen fuel system. The turbopump, which pumped the fuels from the storage tanks to the engine, was the primary tool for restarting the engine. The NERVA had to be able to restart in space on its own using a safe preprogrammed startup system. Lewis researchers endeavored to design and test this system. This non-nuclear Kiwi engine, seen here, was being prepared for tests at Lewis’ High Energy Rocket Engine Research Facility (B-1) located at Plum Brook Station. The tests were designed to start an unfueled Kiwi B-1-B reactor and its Aerojet Mark IX turbopump without any external power.

Analysis of Flame Deflector Spray Nozzles in Rocket Engine Test Stands

NASA Technical Reports Server (NTRS)

Sachdev, Jai S.; Ahuja, Vineet; Hosangadi, Ashvin; Allgood, Daniel C.

2010-01-01

The development of a unified tightly coupled multi-phase computational framework is described for the analysis and design of cooling spray nozzle configurations on the flame deflector in rocket engine test stands. An Eulerian formulation is used to model the disperse phase and is coupled to the gas-phase equations through momentum and heat transfer as well as phase change. The phase change formulation is modeled according to a modified form of the Hertz-Knudsen equation. Various simple test cases are presented to verify the validity of the numerical framework. The ability of the methodology to accurately predict the temperature load on the flame deflector is demonstrated though application to an actual sub-scale test facility. The CFD simulation was able to reproduce the result of the test-firing, showing that the spray nozzle configuration provided insufficient amount of cooling.

Evaluation of Foam Coolants.

DTIC Science & Technology

HYPERGOLIC ROCKET PROPELLANTS, * FOAM , FILM COOLING, FILM COOLING, LIQUID COOLING, LIQUID ROCKET FUELS, ADDITIVES, HEAT TRANSFER, COOLANTS, LIQUID PROPELLANT ROCKET ENGINES, LIQUID COOLING, CAPTIVE TESTS, FEASIBILITY STUDIES.

Experiment/Analytical Characterization of the RBCC Rocket-Ejector Mode

NASA Technical Reports Server (NTRS)

Ruf, J. H.; Lehman, M.; Pal, S.; Santoro, R. J.; West, J.; Turner, James E. (Technical Monitor)

2000-01-01

Experimental and complementary CFD results from the study of the rocket-ejector mode of a Rocket Based Combined Cycle (RBCC) engine are presented and discussed. The experiments involved systematic flowfield measurements in a two-dimensional, variable geometry rocket-ejector system. The rocket-ejector system utilizes a single two-dimensional, gaseous oxygen/gaseous hydrogen rocket as the ejector. To gain a thorough understanding of the rocket-ejector's internal fluid mechanic/combustion phenomena, experiments were conducted with both direct-connect and sea-level static configurations for a range of rocket operating conditions. Overall system performance was obtained through global measurements of wall static pressure profiles, heat flux profiles and engine thrust, whereas detailed mixing and combustion information was obtained through Raman spectroscopy measurements of major species (oxygen, hydrogen, nitrogen and water vapor). The experimental results for both the direct-connect and sea-level static configurations are compared with CFD predictions of the flowfield.

A Comparison of Propulsion Concepts for SSTO Reusable Launchers

NASA Astrophysics Data System (ADS)

Varvill, R.; Bond, A.

This paper discusses the relevant selection criteria for a single stage to orbit (SSTO) propulsion system and then reviews the characteristics of the typical engine types proposed for this role against these criteria. The engine types considered include Hydrogen/Oxygen (H2/O2) rockets, Scramjets, Turbojets, Turborockets and Liquid Air Cycle Engines. In the authors opinion none of the above engines are able to meet all the necessary criteria for an SSTO propulsion system simultaneously. However by selecting appropriate features from each it is possible to synthesise a new class of engines which are specifically optimised for the SSTO role. The resulting engines employ precooling of the airstream and a high internal pressure ratio to enable a relatively conventional high pressure rocket combustion chamber to be utilised in both airbreathing and rocket modes. This results in a significant mass saving with installation advantages which by careful design of the cycle thermodynamics enables the full potential of airbreathing to be realised. The SABRE engine which powers the SKYLON launch vehicle is an example of one of these so called `Precooled hybrid airbreathing rocket engines' and the concep- tual reasoning which leads to its main design parameters are described in the paper.

Rocket Engine Nozzle Side Load Transient Analysis Methodology: A Practical Approach

NASA Technical Reports Server (NTRS)

Shi, John J.

2005-01-01

At the sea level, a phenomenon common with all rocket engines, especially for a highly over-expanded nozzle, during ignition and shutdown is that of flow separation as the plume fills and empties the nozzle, Since the flow will be separated randomly. it will generate side loads, i.e. non-axial forces. Since rocket engines are designed to produce axial thrust to power the vehicles, it is not desirable to be excited by non-axial input forcing functions, In the past, several engine failures were attributed to side loads. During the development stage, in order to design/size the rocket engine components and to reduce the risks, the local dynamic environments as well as dynamic interface loads have to be defined. The methodology developed here is the way to determine the peak loads and shock environments for new engine components. In the past it is not feasible to predict the shock environments, e.g. shock response spectra, from one engine to the other, because it is not scaleable. Therefore, the problem has been resolved and the shock environments can be defined in the early stage of new engine development. Additional information is included in the original extended abstract.

Mathematical and computational model for the analysis of micro hybrid rocket motor

NASA Astrophysics Data System (ADS)

Stoia-Djeska, Marius; Mingireanu, Florin

2012-11-01

The hybrid rockets use a two-phase propellant system. In the present work we first develop a simplified model of the coupling of the hybrid combustion process with the complete unsteady flow, starting from the combustion port and ending with the nozzle. The physical and mathematical model are adapted to the simulations of micro hybrid rocket motors. The flow model is based on the one-dimensional Euler equations with source terms. The flow equations and the fuel regression rate law are solved in a coupled manner. The platform of the numerical simulations is an implicit fourth-order Runge-Kutta second order cell-centred finite volume method. The numerical results obtained with this model show a good agreement with published experimental and numerical results. The computational model developed in this work is simple, computationally efficient and offers the advantage of taking into account a large number of functional and constructive parameters that are used by the engineers.

Parametric Model of an Aerospike Rocket Engine

NASA Technical Reports Server (NTRS)

Korte, J. J.

2000-01-01

A suite of computer codes was assembled to simulate the performance of an aerospike engine and to generate the engine input for the Program to Optimize Simulated Trajectories. First an engine simulator module was developed that predicts the aerospike engine performance for a given mixture ratio, power level, thrust vectoring level, and altitude. This module was then used to rapidly generate the aerospike engine performance tables for axial thrust, normal thrust, pitching moment, and specific thrust. Parametric engine geometry was defined for use with the engine simulator module. The parametric model was also integrated into the iSIGHTI multidisciplinary framework so that alternate designs could be determined. The computer codes were used to support in-house conceptual studies of reusable launch vehicle designs.

Parametric Model of an Aerospike Rocket Engine

NASA Technical Reports Server (NTRS)

Korte, J. J.

2000-01-01

A suite of computer codes was assembled to simulate the performance of an aerospike engine and to generate the engine input for the Program to Optimize Simulated Trajectories. First an engine simulator module was developed that predicts the aerospike engine performance for a given mixture ratio, power level, thrust vectoring level, and altitude. This module was then used to rapidly generate the aerospike engine performance tables for axial thrust, normal thrust, pitching moment, and specific thrust. Parametric engine geometry was defined for use with the engine simulator module. The parametric model was also integrated into the iSIGHT multidisciplinary framework so that alternate designs could be determined. The computer codes were used to support in-house conceptual studies of reusable launch vehicle designs.

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.

Cryostatless high temperature supercurrent bearings for rocket engine turbopumps

NASA Technical Reports Server (NTRS)

Rao, Dantam K.; Dill, James F.

1989-01-01

The rocket engine systems examined include SSME, ALS, and CTV systems. The liquid hydrogen turbopumps in the SSME and ALS vehicle systems are identified as potentially attractive candidates for development of Supercurrent Bearings since the temperatures around the bearings is about 30 K, which is considerably lower than the 95 K transition temperatures of HTS materials. At these temperatures, the current HTS materials are shown to be capable of developing significantly higher current densities. This higher current density capability makes the development of supercurrent bearings for rocket engines an attractive proposition. These supercurrent bearings are also shown to offer significant advantages over conventional bearings used in rocket engines. They can increase the life and reliability over rolling element bearings because of noncontact operation. They offer lower power loss over conventional fluid film bearings. Compared to conventional magnetic bearings, they can reduce the weight of controllers significantly, and require lower power because of the use of persistent currents. In addition, four technology areas that require further attention have been identified. These are: Supercurrent Bearing Conceptual Design Verification; HTS Magnet Fabrication and Testing; Cryosensors and Controller Development; and Rocket Engine Environmental Compatibility Testing.

Hot rocket plume experiment - Survey and conceptual design. [of rhenium-iridium bipropellants

NASA Technical Reports Server (NTRS)

Millard, Jerry M.; Luan, Taylor W.; Dowdy, Mack W.

1992-01-01

Attention is given to a space-borne engine plume experiment study to fly an experiment which will both verify and quantify the reduced contamination from advanced rhenium-iridium earth-storable bipropellant rockets (hot rockets) and provide a correlation between high-fidelity, in-space measurements and theoretical plume and surface contamination models. The experiment conceptual design is based on survey results from plume and contamination technologists throughout the U.S. With respect to shuttle use, cursory investigations validate Hitchhiker availability and adaptability, adequate remote manipulator system (RMS) articulation and dynamic capability, acceptable RMS attachment capability, adequate power and telemetry capability, and adequate flight altitude and attitude/orbital capability.

Technology for Transient Simulation of Vibration during Combustion Process in Rocket Thruster

NASA Astrophysics Data System (ADS)

Zubanov, V. M.; Stepanov, D. V.; Shabliy, L. S.

2018-01-01

The article describes the technology for simulation of transient combustion processes in the rocket thruster for determination of vibration frequency occurs during combustion. The engine operates on gaseous propellant: oxygen and hydrogen. Combustion simulation was performed using the ANSYS CFX software. Three reaction mechanisms for the stationary mode were considered and described in detail. The way for obtaining quick CFD-results with intermediate combustion components using an EDM model was found. The way to generate the Flamelet library with CFX-RIF was described. A technique for modeling transient combustion processes in the rocket thruster was proposed based on the Flamelet library. A cyclic irregularity of the temperature field like vortex core precession was detected in the chamber. Frequency of flame precession was obtained with the proposed simulation technique.

Composite Material Application to Liquid Rocket Engines

NASA Technical Reports Server (NTRS)

Judd, D. C.

1982-01-01

The substitution of reinforced plastic composite (RPC) materials for metal was studied. The major objectives were to: (1) determine the extent to which composite materials can be beneficially used in liquid rocket engines; (2) identify additional technology requirements; and (3) determine those areas which have the greatest potential for return. Weight savings, fabrication costs, performance, life, and maintainability factors were considered. Two baseline designs, representative of Earth to orbit and orbit to orbit engine systems, were selected. Weight savings are found to be possible for selected components with the substitution of materials for metal. Various technology needs are identified before RPC material can be used in rocket engine applications.

Reusable rocket engine intelligent control system framework design, phase 2

NASA Technical Reports Server (NTRS)

Nemeth, ED; Anderson, Ron; Ols, Joe; Olsasky, Mark

1991-01-01

Elements of an advanced functional framework for reusable rocket engine propulsion system control are presented for the Space Shuttle Main Engine (SSME) demonstration case. Functional elements of the baseline functional framework are defined in detail. The SSME failure modes are evaluated and specific failure modes identified for inclusion in the advanced functional framework diagnostic system. Active control of the SSME start transient is investigated, leading to the identification of a promising approach to mitigating start transient excursions. Key elements of the functional framework are simulated and demonstration cases are provided. Finally, the advanced function framework for control of reusable rocket engines is presented.

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.

29. SATURN ROCKET ENGINE LOCATED ON NORTH SIDE OF STATIC ...

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

29. SATURN ROCKET ENGINE LOCATED ON NORTH SIDE OF STATIC TEST STAND - DETAILS OF THE EXPANSION NOZZLE. - Marshall Space Flight Center, Saturn Propulsion & Structural Test Facility, East Test Area, Huntsville, Madison County, AL

B-1 and B-3 Test Stands at NASA’s Plum Brook Station

NASA Image and Video Library

1966-09-21

Operation of the High Energy Rocket Engine Research Facility (B-1), left, and Nuclear Rocket Dynamics and Control Facility (B-3) at the National Aeronautics and Space Administration’s (NASA) Plum Brook Station in Sandusky, Ohio. The test stands were constructed in the early 1960s to test full-scale liquid hydrogen fuel systems in simulated altitude conditions. Over the next decade each stand was used for two major series of liquid hydrogen rocket tests: the Nuclear Engine for Rocket Vehicle Application (NERVA) and the Centaur second-stage rocket program. The different components of these rocket engines could be studied under flight conditions and adjusted without having to fire the engine. Once the preliminary studies were complete, the entire engine could be fired in larger facilities. The test stands were vertical towers with cryogenic fuel and steam ejector systems. B-1 was 135 feet tall, and B-3 was 210 feet tall. Each test stand had several levels, a test section, and ground floor shop areas. The test stands relied on an array of support buildings to conduct their tests, including a control building, steam exhaust system, and fuel storage and pumping facilities. A large steam-powered altitude exhaust system reduced the pressure at the exhaust nozzle exit of each test stand. This allowed B-1 and B-3 to test turbopump performance in conditions that matched the altitudes of space.

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.

Engine System Loads Development for the Fastrac 60K Flight Engine

NASA Technical Reports Server (NTRS)

Frady, Greg; Christensen, Eric R.; Mims, Katherine; Harris, Don; Parks, Russell; Brunty, Joseph

2000-01-01

Early implementation of structural dynamics finite element analyses for calculation of design loads is considered common design practice for high volume manufacturing industries such as automotive and aeronautical industries. However, with the rarity of rocket engine development programs starts, these tools are relatively new to the design of rocket engines. In the new Fastrac engine program, the focus has been to reduce the cost to weight ratio; current structural dynamics analysis practices were tailored in order to meet both production and structural design goals. Perturbation of rocket engine design parameters resulted in a number of Fastrac load cycles necessary to characterize the impact due to mass and stiffness changes. Evolution of loads and load extraction methodologies, parametric considerations and a discussion of load path sensitivities are discussed.

Design considerations in clustering nuclear rocket engines

NASA Technical Reports Server (NTRS)

Sager, Paul H.

1992-01-01

An initial investigation of the design considerations in clustering nuclear rocket engines for space transfer vehicles has been made. The clustering of both propulsion modules (which include start tanks) and nuclear rocket engines installed directly to a vehicle core tank appears to be feasible. Special provisions to shield opposite run tanks and the opposite side of a core tank - in the case of the boost pump concept - are required; the installation of a circumferential reactor side shield sector appears to provide an effective solution to this problem. While the time response to an engine-out event does not appear to be critical, the gimbal displacement required appears to be important. Since an installation of three engines offers a substantial reduction in gimbal requirements for engine-out and it may be possible to further enhance mission reliability with the greater number of engines, it is recommended that a cluster of four engines be considered.

Design considerations in clustering nuclear rocket engines

NASA Astrophysics Data System (ADS)

Sager, Paul H.

1992-07-01

An initial investigation of the design considerations in clustering nuclear rocket engines for space transfer vehicles has been made. The clustering of both propulsion modules (which include start tanks) and nuclear rocket engines installed directly to a vehicle core tank appears to be feasible. Special provisions to shield opposite run tanks and the opposite side of a core tank - in the case of the boost pump concept - are required; the installation of a circumferential reactor side shield sector appears to provide an effective solution to this problem. While the time response to an engine-out event does not appear to be critical, the gimbal displacement required appears to be important. Since an installation of three engines offers a substantial reduction in gimbal requirements for engine-out and it may be possible to further enhance mission reliability with the greater number of engines, it is recommended that a cluster of four engines be considered.

Investigation of the Rocket Induced Flow Field in a Rectangular Duct

NASA Technical Reports Server (NTRS)

Landrum, D. Brian; Thames, Mignon; Parkinson, Doug; Gautney, Serena; Hawk, Clark

1999-01-01

Several tests were performed on a one-sixth scale Rocket Based Combined Cycle (RBCC) engine model at the University of Alabama in Huntsville. The UAH RBCC facility consists of a rectangular duct with a vertical strut mounted in the center. The scaled strut consists of two supersonic rocket nozzles with an embedded vertical turbine between the rocket nozzles. The tests included mass flow, flow visualization and horizontal pressure traverses. The mass flow test indicated a c:hoked condition when the rocket chamber pressure is between 200 psi and 300 psi. The flow visualization tests narrowed the rocket chamber pressure range from, 250 psi to 300 psi. Also, from this t.est, an assumption of a minimum

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.

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.

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.

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.

Sea level side loads in high-area-ratio rocket engines

NASA Technical Reports Server (NTRS)

Nave, L. H.; Coffey, G. A.

1973-01-01

An empirical separation and side load model to obtain applied aerodynamic loads has been developed based on data obtained from full-scale J-2S (265K-pound-thrust engine with an area ratio of 40:1) engine and model testing. Experimental data include visual observations of the separation patterns that show the dynamic nature of the separation phenomenon. Comparisons between measured and applied side loads are made. Correlations relating the separation location to the applied side loads and the methods used to determine the separation location are given.

Chemical propulsion - The old and the new challenges

NASA Technical Reports Server (NTRS)

Mccarty, J. P.; Lombardo, J. A.

1973-01-01

The historical background concerning the application of liquid propellant rockets is considered. Progress to date in chemical liquid propellant rocket engines can be summarized as an increase in performance through the use of more energetic propellant combinations and increased combustion pressure. New advances regarding liquid propellant rocket engines are related to the requirement for reusability in connection with the development of the Space Shuttle.

Dual-fuel, dual-mode rocket engine

NASA Technical Reports Server (NTRS)

Martin, James A. (Inventor)

1989-01-01

The invention relates to a dual fuel, dual mode rocket engine designed to improve the performance of earth-to-orbit vehicles. For any vehicle that operates from the earth's surface to earth orbit, it is advantageous to use two different fuels during its ascent. A high density impulse fuel, such as kerosene, is most efficient during the first half of the trajectory. A high specific impulse fuel, such as hydrogen, is most efficient during the second half of the trajectory. The invention allows both fuels to be used with a single rocket engine. It does so by adding a minimum number of state-of-the-art components to baseline single made rocket engines, and is therefore relatively easy to develop for near term applications. The novelty of this invention resides in the mixing of fuels before exhaust nozzle cooling. This allows all of the engine fuel to cool the exhaust nozzle, and allows the ratio of fuels used throughout the flight depend solely on performance requirements, not cooling requirements.

Nonlinear Control of a Reusable Rocket Engine for Life Extension

NASA Technical Reports Server (NTRS)

Lorenzo, Carl F.; Holmes, Michael S.; Ray, Asok

1998-01-01

This paper presents the conceptual development of a life-extending control system where the objective is to achieve high performance and structural durability of the plant. A life-extending controller is designed for a reusable rocket engine via damage mitigation in both the fuel (H2) and oxidizer (O2) turbines while achieving high performance for transient responses of the combustion chamber pressure and the O2/H2 mixture ratio. The design procedure makes use of a combination of linear and nonlinear controller synthesis techniques and also allows adaptation of the life-extending controller module to augment a conventional performance controller of the rocket engine. The nonlinear aspect of the design is achieved using non-linear parameter optimization of a prescribed control structure. Fatigue damage in fuel and oxidizer turbine blades is primarily caused by stress cycling during start-up, shutdown, and transient operations of a rocket engine. Fatigue damage in the turbine blades is one of the most serious causes for engine failure.

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.

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.

Metal Matrix Composites for Rocket Engine Applications

NASA Technical Reports Server (NTRS)

McDonald, Kathleen R.; Wooten, John R.

2000-01-01

This document is from a presentation about the applications of Metal Matrix Composites (MMC) in rocket engines. Both NASA and the Air Force have goals which would reduce the costs and the weight of launching spacecraft. Charts show the engine weight distribution for both reuseable and expendable engine components. The presentation reviews the operating requirements for several components of the rocket engines. The next slide reviews the potential benefits of MMCs in general and in use as materials for Advanced Pressure Casting. The next slide reviews the drawbacks of MMCs. The reusable turbopump housing is selected to review for potential MMC application. The presentation reviews solutions for reusable turbopump materials, pointing out some of the issues. It also reviews the development of some of the materials.

36. Historic photo of Building 202 interior, shows shop area ...

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

36. Historic photo of Building 202 interior, shows shop area with engineers assembling twenty-thousand-pound-thrust rocket engine, December 15, 1958. On file at NASA Plumbrook Research Center, Sandusky, Ohio. NASA photo number C-49343. - Rocket Engine Testing Facility, GRC Building No. 202, NASA Glenn Research Center, Cleveland, Cuyahoga County, OH

32. Historic view of Building 202 test stand A with ...

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

32. Historic view of Building 202 test stand A with rocket engine, close-up detail of engine, November 19, 1957. On file at NASA Plumbrook Research Center, Sandusky, Ohio. NASA photo number C-46492. - Rocket Engine Testing Facility, GRC Building No. 202, NASA Glenn Research Center, Cleveland, Cuyahoga County, OH

40. Historic photo of Building 202 test cell interior, with ...

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

40. Historic photo of Building 202 test cell interior, with engineers working on rocket engine mounted on test stand A, June 26, 1959. On file at NASA Plumbrook Research Center, Sandusky, Ohio. NASA photo number C-51026. - Rocket Engine Testing Facility, GRC Building No. 202, NASA Glenn Research Center, Cleveland, Cuyahoga County, OH

A Versatile Rocket Engine Hot Gas Facility

NASA Technical Reports Server (NTRS)

Green, James M.

1993-01-01

The capabilities of a versatile rocket engine facility, located in the Rocket Laboratory at the NASA Lewis Research Center, are presented. The gaseous hydrogen/oxygen facility can be used for thermal shock and hot gas testing of materials and structures as well as rocket propulsion testing. Testing over a wide range of operating conditions in both fuel and oxygen rich regimes can be conducted, with cooled or uncooled test specimens. The size and location of the test cell provide the ability to conduct large amounts of testing in short time periods with rapid turnaround between programs.

NASA Conducts Final RS-25 Rocket Engine Test of 2017

NASA Image and Video Library

2017-12-13

NASA engineers at Stennis Space Center capped a year of Space Launch System testing with a final RS-25 rocket engine hot fire on Dec. 13. The 470-second test on the A-1 Test Stand was a “green run” test of an RS-25 flight controller. The engine tested also included a large 3-D-printed part, a pogo accumulator assembly, scheduled for use on future RS-25 flight engines.

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.

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.

Saturn Apollo Program

NASA Image and Video Library

1963-01-01

This drawing clearly shows the comparative sizes of the rocket engines used to launch the Saturn vehicles. The RL-10 and the H-1 engines were used to launch the Saturn I rockets. The J-2 engine was used on the second stage of Saturn IB and the second and third stages of Saturn V. The F-1 engine was used on the first stage of the Saturn V.

Enhanced Large Solid Rocket Motor Understanding Through Performance Margin Testing: RSRM Five-Segment Engineering Test Motor (ETM-3)

NASA Technical Reports Server (NTRS)

Huppi, Hal; Tobias, Mark; Seiler, James

2003-01-01

The Five-Segment Engineering Test Motor (ETM-3) is an extended length reusable solid rocket motor (RSRM) intended to increase motor performance and internal environments above the current four-segment RSRM flight motor. The principal purpose of ETM-3 is to provide a test article for RSRM component margin testing. As the RSRM and Space Shuttle in general continue to age, replacing obsolete materials becomes an ever-increasing issue. Having a five-segment motor that provides environments in excess of normal opera- tion allows a mechanism to subject replacement materials to a more severe environment than experienced in flight. Additionally, ETM-3 offers a second design data point from which to develop and/or validate analytical models that currently have some level of empiricism associated with them. These enhanced models have the potential to further the understanding of RSRM motor performance and solid rocket motor (SRM) propulsion in general. Furthermore, these data could be leveraged to support a five-segment booster (FSB) development program should the Space Shuttle program choose to pursue this option for abort mode enhancements during the ascent phase. A tertiary goal of ETM-3 is to challenge both the ATK Thiokol Propulsion and NASA MSFC technical personnel through the design and analysis of a large solid rocket motor without the benefit of a well-established performance database such as the RSRM. The end result of this undertaking will be a more competent and experienced workforce for both organizations. Of particular interest are the motor design characteristics and the systems engineering approach used to conduct a complex yet successful large motor static test. These aspects of ETM-3 and more will be summarized.

Entry, Descent, and Landing With Propulsive Deceleration: Supersonic Retropropulsion Wind Tunnel Testing

NASA Technical Reports Server (NTRS)

Palaszewski, Bryan

2012-01-01

The future exploration of the Solar System will require innovations in transportation and the use of entry, descent, and landing (EDL) systems at many planetary landing sites. The cost of space missions has always been prohibitive, and using the natural planetary and planet s moons atmosphere for entry, descent, and landing can reduce the cost, mass, and complexity of these missions. This paper will describe some of the EDL ideas for planetary entry and survey the overall technologies for EDL that may be attractive for future Solar System missions. Future EDL systems may include an inflatable decelerator for the initial atmospheric entry and an additional supersonic retro-propulsion (SRP) rocket system for the final soft landing. As part of those efforts, NASA began to conduct experiments to gather the experimental data to make informed decisions on the "best" EDL options. A model of a three engine retro-propulsion configuration with a 2.5 in. diameter sphere-cone aeroshell model was tested in the NASA Glenn 1- by 1-Foot Supersonic Wind Tunnel (SWT). The testing was conducted to identify potential blockage issues in the tunnel, and visualize the rocket flow and shock interactions during supersonic and hypersonic entry conditions. Earlier experimental testing of a 70 Viking-like (sphere-cone) aeroshell was conducted as a baseline for testing of a supersonic retro-propulsion system. This baseline testing defined the flow field around the aeroshell and from this comparative baseline data, retro-propulsion options will be assessed. Images and analyses from the SWT testing with 300- and 500-psia rocket engine chamber pressures are presented here. The rocket engine flow was simulated with a non-combusting flow of air.

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.

A Roadmap for Aircraft Engine Life Extending Control

NASA Technical Reports Server (NTRS)

Guo, Ten-Huei

2001-01-01

The concept of Aircraft Engine Life Extending Control is introduced. A brief description of the tradeoffs between performance and engine life are first explained. The overall goal of the life extending controller is to reduce the engine operating cost by extending the on-wing engine life while improving operational safety. The research results for NASA's Rocket Engine life extending control program are also briefly described. Major building blocks of the Engine Life Extending Control architecture are examined. These blocks include: life prediction models, engine operation models, stress and thermal analysis tools, control schemes, and intelligent control systems. The technology areas that would likely impact the successful implementation of an aircraft engine life extending control are also briefly described. Near, intermediate, and long term goals of NASA's activities are also presented.

Thrust Vector Control for Nuclear Thermal Rockets

NASA Technical Reports Server (NTRS)

Ensworth, Clinton B. F.

2013-01-01

Future space missions may use Nuclear Thermal Rocket (NTR) stages for human and cargo missions to Mars and other destinations. The vehicles are likely to require engine thrust vector control (TVC) to maintain desired flight trajectories. This paper explores requirements and concepts for TVC systems for representative NTR missions. Requirements for TVC systems were derived using 6 degree-of-freedom models of NTR vehicles. Various flight scenarios were evaluated to determine vehicle attitude control needs and to determine the applicability of TVC. Outputs from the models yielded key characteristics including engine gimbal angles, gimbal rates and gimbal actuator power. Additional factors such as engine thrust variability and engine thrust alignment errors were examined for impacts to gimbal requirements. Various technologies are surveyed for TVC systems for the NTR applications. A key factor in technology selection is the unique radiation environment present in NTR stages. Other considerations including mission duration and thermal environments influence the selection of optimal TVC technologies. Candidate technologies are compared to see which technologies, or combinations of technologies best fit the requirements for selected NTR missions. Representative TVC systems are proposed and key properties such as mass and power requirements are defined. The outputs from this effort can be used to refine NTR system sizing models, providing higher fidelity definition for TVC systems for future studies.

Advanced rocket propulsion

NASA Technical Reports Server (NTRS)

Obrien, Charles J.

1993-01-01

Existing NASA research contracts are supporting development of advanced reinforced polymer and metal matrix composites for use in liquid rocket engines of the future. Advanced rocket propulsion concepts, such as modular platelet engines, dual-fuel dual-expander engines, and variable mixture ratio engines, require advanced materials and structures to reduce overall vehicle weight as well as address specific propulsion system problems related to elevated operating temperatures, new engine components, and unique operating processes. High performance propulsion systems with improved manufacturability and maintainability are needed for single stage to orbit vehicles and other high performance mission applications. One way to satisfy these needs is to develop a small engine which can be clustered in modules to provide required levels of total thrust. This approach should reduce development schedule and cost requirements by lowering hardware lead times and permitting the use of existing test facilities. Modular engines should also reduce operational costs associated with maintenance and parts inventories.

Exposed by Rocket Engine Blasts

NASA Image and Video Library

2012-08-12

This color image from NASA Curiosity rover shows an area excavated by the blast of the Mars Science Laboratory descent stage rocket engines. This is part of a larger, high-resolution color mosaic made from images obtained by Curiosity Mast Camera.

Computational Analyses in Support of Sub-scale Diffuser Testing for the A-3 Facility. Part 3; Aero-Acoustic Analyses and Experimental Validation

NASA Technical Reports Server (NTRS)

Allgood, Daniel C.; Graham, Jason S.; McVay, Greg P.; Langford, Lester L.

2008-01-01

A unique assessment of acoustic similarity scaling laws and acoustic analogy methodologies in predicting the far-field acoustic signature from a sub-scale altitude rocket test facility at the NASA Stennis Space Center was performed. A directional, point-source similarity analysis was implemented for predicting the acoustic far-field. In this approach, experimental acoustic data obtained from "similar" rocket engine tests were appropriately scaled using key geometric and dynamic parameters. The accuracy of this engineering-level method is discussed by comparing the predictions with acoustic far-field measurements obtained. In addition, a CFD solver was coupled with a Lilley's acoustic analogy formulation to determine the improvement of using a physics-based methodology over an experimental correlation approach. In the current work, steady-state Reynolds-averaged Navier-Stokes calculations were used to model the internal flow of the rocket engine and altitude diffuser. These internal flow simulations provided the necessary realistic input conditions for external plume simulations. The CFD plume simulations were then used to provide the spatial turbulent noise source distributions in the acoustic analogy calculations. Preliminary findings of these studies will be discussed.

USM3D Simulations of Saturn V Plume Induced Flow Separation

NASA Technical Reports Server (NTRS)

Deere, Karen; Elmlilgui, Alaa; Abdol-Hamid, K. S.

2011-01-01

The NASA Constellation Program included the Ares V heavy lift cargo vehicle. During the design stage, engineers questioned if the Plume Induced Flow Separation (PIFS) that occurred along Saturn V rocket during moon missions at some flight conditions, would also plague the newly proposed rocket. Computational fluid dynamics (CFD) was offered as a tool for initiating the investigation of PIFS along the Ares V rocket. However, CFD best practice guidelines were not available for such an investigation. In an effort to establish a CFD process and define guidelines for Ares V powered simulations, the Saturn V vehicle was used because PIFS flight data existed. The ideal gas, computational flow solver USM3D was evaluated for its viability in computing PIFS along the Saturn V vehicle with F-1 engines firing. Solutions were computed at supersonic freestream conditions, zero degree angle of attack, zero degree sideslip, and at flight Reynolds numbers. The effects of solution sensitivity to grid refinement, turbulence models, and the engine boundary conditions on the predicted PIFS distance along the Saturn V were discussed and compared to flight data from the Apollo 11 mission AS-506.

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

Blood Pump Development Using Rocket Engine Flow Simulation Technology

NASA Technical Reports Server (NTRS)

Kiris, Cetin C.; Kwak, Dochan

2002-01-01

This viewgraph presentation provides information on the transfer of rocket engine flow simulation technology to work involving the development of blood pumps. Details are offered regarding the design and requirements of mechanical heart assist devices, or VADs (ventricular assist device). There are various computational fluid dynamics issues involved in the visualization of flow in such devices, and these are highlighted and compared to those of rocket turbopumps.

Modeling Potential Carbon Monoxide Exposure Due to Operation of a Major Rocket Engine Altitude Test Facility Using Computational Fluid Dynamics

NASA Technical Reports Server (NTRS)

Blotzer, Michael J.; Woods, Jody L.

2009-01-01

This viewgraph presentation reviews computational fluid dynamics as a tool for modelling the dispersion of carbon monoxide at the Stennis Space Center's A3 Test Stand. The contents include: 1) Constellation Program; 2) Constellation Launch Vehicles; 3) J2X Engine; 4) A-3 Test Stand; 5) Chemical Steam Generators; 6) Emission Estimates; 7) Located in Existing Test Complex; 8) Computational Fluid Dynamics; 9) Computational Tools; 10) CO Modeling; 11) CO Model results; and 12) Next steps.

FDNS CFD Code Benchmark for RBCC Ejector Mode Operation: Continuing Toward Dual Rocket Effects

NASA Technical Reports Server (NTRS)

West, Jeff; Ruf, Joseph H.; Turner, James E. (Technical Monitor)

2000-01-01

Computational Fluid Dynamics (CFD) analysis results are compared with benchmark quality test data from the Propulsion Engineering Research Center's (PERC) Rocket Based Combined Cycle (RBCC) experiments to verify fluid dynamic code and application procedures. RBCC engine flowpath development will rely on CFD applications to capture the multi -dimensional fluid dynamic interactions and to quantify their effect on the RBCC system performance. Therefore, the accuracy of these CFD codes must be determined through detailed comparisons with test data. The PERC experiments build upon the well-known 1968 rocket-ejector experiments of Odegaard and Stroup by employing advanced optical and laser based diagnostics to evaluate mixing and secondary combustion. The Finite Difference Navier Stokes (FDNS) code [2] was used to model the fluid dynamics of the PERC RBCC ejector mode configuration. Analyses were performed for the Diffusion and Afterburning (DAB) test conditions at the 200-psia thruster operation point, Results with and without downstream fuel injection are presented.

A Review of ETM-03 (A Five Segment Shuttle RSRM Configuration) Ballistic Performance

NASA Technical Reports Server (NTRS)

McMillin, J. E.; Furfaro, J. A.

2004-01-01

Marshall Space Flight Center and ATK Thiokol Propulsion worked together on the engineering design of a five-segment Engineering Test Motor (ETM-03), the world's largest segmented solid rocket motor. The data from ETM-03's static test have helped to provide a better understanding of the Reusable Solid Rocket Motor's (RSRM's) margins and the techniques and models used to simulate solid rocket motor performance. The enhanced performance of ETM-03 was achieved primarily by the addition of a RSRM center segment. Added motor performance was also achieved with a nozzle throat diameter increase and the incorporation of an Extended Aft Exit Cone (EAEC). Performance parameters such as web time, action time, head-end pressure, web time average pressure, maximum thrust, mass flow rate, centerline Mach number, pressure and thrust integrals were all increased over RSRM. In some cases, the performance increases were substantial. Overall, the measured data were exceptionally close to the pretest predictions.

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.

Upper-stage space shuttle propulsion by means of separate scramjet and rocket engines

NASA Technical Reports Server (NTRS)

Franciscus, L. C.; Allen, J. L.

1972-01-01

A preliminary mission study of a reusable vehicle from staging to orbit indicates payload advantages for a dual-propulsion system consisting of separate scramjet and rocket engines. In the analysis the scramjet operated continuously and the initiation of rocket operation was varied. For a stage weight of 500,000 lb the payload was 10.4 percent of stage weight or 70 percent greater than that of a comparable all-rocket-powered stage. When compared with a reusable two-state rocket vehicle having 50,000 lb payload, the use of the dual propulsion system for the second stage resulted in significant decreases in lift-off weight and empty weight, indicating possible lower hardware costs.

Main lines of scientific and technical research at the Soviet Jet Propulsion Research Institute (RNII), 1933 - 1942

NASA Technical Reports Server (NTRS)

Shchetinkov, Y. S.

1977-01-01

The rapid development of rocketry in the U.S.S.R. during the post-war years was due largely to pre-war activity; in particular, to investigations conducted in the Jet Propulsion Research Institute (RNII). The history of RNII commenced in 1933, resulting from the merger of two rocket research organizations. Previous research was continued in areas of solid-propellant rockets, jet-assisted take-off of aircraft, liquid propellant engines (generally with nitric acid as the oxidizer), liquid-propellant rockets (generally with oxgen as the oxidizer), ram jet engines, rockets with and without wings, and rocket planes. RNII research is described and summarized for the years 1933-1942.

AJ26 rocket engine testing news briefing

NASA Technical Reports Server (NTRS)

2010-01-01

NASA's John C. Stennis Space Center Director Gene Goldman (center) stands in front of a 'pathfinder' rocket engine with Orbital Sciences Corp. President and Chief Operating Officer J.R. Thompson (left) and Aerojet President Scott Seymour during a Feb. 24 news briefing at the south Mississippi facility. The leaders appeared together to announce a partnership for testing Aerojet AJ26 rocket engines at Stennis. The engines will be used to power Orbital's Taurus II space vehicles to provide commercial cargo transportation missions to the International Space Station for NASA. During the event, the Stennis partnership with Orbital was cited as an example of the new direction of NASA to work with commercial interests for space travel and transport.

Orbit Transfer Rocket Engine Technology Program, Advanced Engine Study Task D.6

DTIC Science & Technology

1992-02-28

l!J~iliiJl 1. Report No. 2. Government Accession No. 3 . Recipient’s Catalog No. NASA 187215 4. Title and Subtitle 5. Report Date ORBIT TRANSFER ROCKET...Engine Study, three primary subtasks were accomplished: 1) Design and Parametric Data, 2) Engine Requirement Variation Studies, and 3 ) Vehicle Study...Mixture Ratio Parametrics 18 3 . Thrust Parametrics Off-Design Mixture Ratio Scans 22 4. Expansion Area Ratio Parametrics 24 5. OTV 20 klbf Engine Off

Video File - NASA Conducts Final RS-25 Rocket Engine Test of 2017

NASA Image and Video Library

2017-12-13

NASA engineers at Stennis Space Center capped a year of Space Launch System testing with a final RS-25 rocket engine hot fire on Dec. 13. The 470-second test on the A-1 Test Stand was a “green run” test of an RS-25 flight controller. The engine tested also included a large 3-D-printed part, a pogo accumulator assembly, scheduled for use on future RS-25 flight engines.

RS 25 Hot Fire test

NASA Image and Video Library

2016-08-18

The 7.5-minute test conducted at NASA’s Stennis Space Center is part of a series of tests designed to put the upgraded former space shuttle engines through the rigorous temperature and pressure conditions they will experience during a launch. The tests also support the development of a new controller, or “brain,” for the engine, which monitors engine status and communicates between the rocket and the engine, relaying commands to the engine and transmitting data back to the rocket.

RS-25 Hot Fire test

NASA Image and Video Library

2016-08-18

The 7.5-minute test conducted at NASA’s Stennis Space Center is part of a series of tests designed to put the upgraded former space shuttle engines through the rigorous temperature and pressure conditions they will experience during a launch. The tests also support the development of a new controller, or “brain,” for the engine, which monitors engine status and communicates between the rocket and the engine, relaying commands to the engine and transmitting data back to the rocket.

Daniel Sokolowski in the Rocket Operations Building

NASA Image and Video Library

1966-06-21

Dan Sokolowski worked as an engineering coop student at the National Aeronautics and Space Administration (NASA) Lewis Research Center from 1962 to 1966 while earning his Mechanical Engineering degree from Purdue. At the time of this photograph Sokolowski had just been hired as a permanent NASA employee in the Chemical Rocket Evaluation Branch of the Chemical Rocket Division. He had also just won a regional American Institute of Aeronautics and Astronautics competition for his paper on high and low-frequency combustion instability. The resolution of the low-frequency combustion instability, or chugging, in liquid hydrogen rocket systems was one of Lewis’ more significant feats of the early 1960s. In most rocket engine combustion chambers, the pressure, temperature, and flows are in constant flux. The engine is considered to be operating normally if the fluctuations remain random and within certain limits. Lewis researchers used high-speed photography to study and define Pratt and Whitney’s RL-10’s combustion instability by throttling the engine under the simulated flight conditions. They found that the injection of a small stream of helium gas into the liquid-oxygen tank immediately stabilized the system. Sokolowski’s later work focused on combustion in airbreathing engines. In 1983 was named Manager of a multidisciplinary program aimed at improving durability of combustor and turbine components. After 39 years Sokolowski retired from NASA in September 2002.

38. Historic photo of Building 202 test cell interior, showing ...

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

38. Historic photo of Building 202 test cell interior, showing damage to test stand A and rocket engine after failure and explosion of engine, December 12, 1958. On file at NASA Plumbrook Research Center, Sandusky, Ohio. NASA photo number C-49376. - Rocket Engine Testing Facility, GRC Building No. 202, NASA Glenn Research Center, Cleveland, Cuyahoga County, OH

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.

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.

Tripropellant combustion process

NASA Technical Reports Server (NTRS)

Kmiec, T. D.; Carroll, R. G.

1988-01-01

The addition of small amounts of hydrogen to the combustion of LOX/hydrocarbon propellants in large rocket booster engines has the potential to enhance the system stability. Programs being conducted to evaluate the effects of hydrogen on the combustion of LOX/hydrocarbon propellants at supercritical pressures are described. Combustion instability has been a problem during the development of large hydrocarbon fueled rocket engines. At the higher combustion chamber pressures expected for the next generation of booster engines, the effect of unstable combustion could be even more destructive. The tripropellant engine cycle takes advantage of the superior cooling characteristics of hydrogen to cool the combustion chamber and a small amount of the hydrogen coolant can be used in the combustion process to enhance the system stability. Three aspects of work that will be accomplished to evaluate tripropellant combustion are described. The first is laboratory demonstration of the benefits through the evaluation of drop size, ignition delay and burning rate. The second is analytical modeling of the combustion process using the empirical relationship determined in the laboratory. The third is a subscale demonstration in which the system stability will be evaluated. The approach for each aspect is described and the analytical models that will be used are presented.

Ignition of Hydrogen Balloons by Model-Rocket-Engine Igniters.

ERIC Educational Resources Information Center

Hartman, Nicholas T.

2003-01-01

Describes an alternative method for exploding hydrogen balloons as a classroom demonstration. Uses the method of igniting the balloons via an electronic match. Includes necessary materials to conduct the demonstration and discusses potential hazards. (SOE)

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.

FDNS CFD Code Benchmark for RBCC Ejector Mode Operation

NASA Technical Reports Server (NTRS)

Holt, James B.; Ruf, Joe

1999-01-01

Computational Fluid Dynamics (CFD) analysis results are compared with benchmark quality test data from the Propulsion Engineering Research Center's (PERC) Rocket Based Combined Cycle (RBCC) experiments to verify fluid dynamic code and application procedures. RBCC engine flowpath development will rely on CFD applications to capture the multi-dimensional fluid dynamic interactions and to quantify their effect on the RBCC system performance. Therefore, the accuracy of these CFD codes must be determined through detailed comparisons with test data. The PERC experiments build upon the well-known 1968 rocket-ejector experiments of Odegaard and Stroup by employing advanced optical and laser based diagnostics to evaluate mixing and secondary combustion. The Finite Difference Navier Stokes (FDNS) code was used to model the fluid dynamics of the PERC RBCC ejector mode configuration. Analyses were performed for both Diffusion and Afterburning (DAB) and Simultaneous Mixing and Combustion (SMC) test conditions. Results from both the 2D and the 3D models are presented.

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.

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.

F-1 Gas Generator test

NASA Image and Video Library

2015-09-03

THE GAS GENERATOR TO AN F-1 ENGINE, THE MOST POWERFUL ROCKET ENGINE EVER BUILT, IS TEST-FIRED AT NASA'S MARSHALL SPACE FLIGHT CENTER IN HUNTSVILLE, ALABAMA, ON SEPT. 3. ALTHOUGH THE ENGINE WAS ORIGINALLY BUILT TO POWER THE SATURN V ROCKETS DURING AMERICA'S MISSIONS TO THE MOON, THIS TEST ARTICLE HAD NEW PARTS CREATED USING ADDITIVE MANUFACTURING, OR 3-D PRINTING, TO TEST THE VIABILITY OF THE TECHNOLOGY FOR BUILDING NEW ENGINE DESIGNS.

The Development of the Wall Momentum Erosive Burning Scaling Law and Macro Scale Erosive Burning Model

DTIC Science & Technology

2010-05-01

burn rate in excess of 2 in/sec at 1000 psi, and Mach numbers that reach 1.0 at the aft end at ignition . Typically, motors with high burning rate...37 VI I. INTRODUCTION Interior ballistics of solid propellant rocket motors continues to be an engineering discipline that is...and one open source paper published between 2005 and 2009 [2, 3, 13]. II. BACKGROUND Erosive burning is a term used in the solid rocket motor

Robot Rocket Rally

NASA Image and Video Library

2014-03-14

CAPE CANAVERAL, Fla. – A torso model of Robonaut 2, identical to R2 already on the International Space Station, is introduced to a crowd of onlookers by Ron Diftler of NASA's Johnson Space Center in Houston. The demonstration was one of several provided during the Robot Rocket Rally. The three-day event at Florida's Kennedy Space Center Visitor Complex is highlighted by exhibits, games and demonstrations of a variety of robots, with exhibitors ranging from school robotics clubs to veteran NASA scientists and engineers. Photo credit: NASA/Kim Shiflett

Calculating Mass Diffusion in High-Pressure Binary Fluids

NASA Technical Reports Server (NTRS)

Bellan, Josette; Harstad, Kenneth

2004-01-01

A comprehensive mathematical model of mass diffusion has been developed for binary fluids at high pressures, including critical and supercritical pressures. Heretofore, diverse expressions, valid for limited parameter ranges, have been used to correlate high-pressure binary mass-diffusion-coefficient data. This model will likely be especially useful in the computational simulation and analysis of combustion phenomena in diesel engines, gas turbines, and liquid rocket engines, wherein mass diffusion at high pressure plays a major role.

Two-phase flow in the cooling circuit of a cryogenic rocket engine

NASA Astrophysics Data System (ADS)

Preclik, D.

1992-07-01

Transient two-phase flow was investigated for the hydrogen cooling circuit of the HM7 rocket engine. The nuclear reactor code ATHLET/THESEUS was adapted to cryogenics and applied to both principal and prototype experiments for validation and simulation purposes. The cooling circuit two-phase flow simulation focused on the hydrogen prechilling and pump transient phase prior to ignition. Both a single- and a multichannel model were designed and employed for a valve leakage flow, a nominal prechilling flow, and a prechilling with a subsequent pump-transient flow. The latter case was performed in order to evaluate the difference between a nominal and a delayed turbo-pump start-up. It was found that an extension of the nominal prechilling sequence in the order of 1 second is sufficient to finally provide for liquid injection conditions of hydrogen which, as commonly known, is undesirable for smooth ignition and engine starting transients.

Parallel Unsteady Turbopump Simulations for Liquid Rocket Engines

NASA Technical Reports Server (NTRS)

Kiris, Cetin C.; Kwak, Dochan; Chan, William

2000-01-01

This paper reports the progress being made towards complete turbo-pump simulation capability for liquid rocket engines. Space Shuttle Main Engine (SSME) turbo-pump impeller is used as a test case for the performance evaluation of the MPI and hybrid MPI/Open-MP versions of the INS3D code. Then, a computational model of a turbo-pump has been developed for the shuttle upgrade program. Relative motion of the grid system for rotor-stator interaction was obtained by employing overset grid techniques. Time-accuracy of the scheme has been evaluated by using simple test cases. Unsteady computations for SSME turbo-pump, which contains 136 zones with 35 Million grid points, are currently underway on Origin 2000 systems at NASA Ames Research Center. Results from time-accurate simulations with moving boundary capability, and the performance of the parallel versions of the code will be presented in the final paper.

Aerodynamic characteristics of MSFC model 454 of the 142 inch solid rocket booster tested in the LeRC 10 foot SWT at Mach numbers of 2.0 and 2.7 (SA6F)

NASA Technical Reports Server (NTRS)

Johnson, J. D.; Burstadt, P. L.; Radford, W. D.

1975-01-01

A 2.112 percent scale Space Shuttle Solid Rocket Booster (SRB) was tested in a ten foot, supersonic wind tunnel. The test Mach numbers were 2.0 and 2.7. Test angles of attack were from minus 5 degrees to plus 185 degrees. The Reynolds numbers ranged from 0.514 to 2.81 million per foot. Test roll angles were 0, 22.5, 45, 90, and 135 degrees. The following configurations were tested: (1) SRB without external protuberances, (2) SRB with an electrical tunnel and a thrust attachment structure, (3) SRB with two engine shroud strakes, (4) SRB with eight engine shroud strakes, and (5) SRB with an electrical tunnel, thrust attachment structure, eight engine shroud strakes, and separation motors.

Focused RBCC Experiments: Two-Rocket Configuration Experiments and Hydrocarbon/Oxygen Rocket Ejector Experiments

NASA Technical Reports Server (NTRS)

Santoro, Robert J.; Pal, Sibtosh

2003-01-01

This addendum report documents the results of two additional efforts for the Rocket Based Combined Cycle (RBCC) rocket-ejector mode research work carried out at the Penn State Propulsion Engineering Research Center in support of NASA s technology development efforts for enabling 3 d generation Reusable Launch Vehicles (RLV). The tasks reported here build on an earlier NASA MSFC funded research program on rocket ejector investigations. The first task investigated the improvements of a gaseous hydrogen/oxygen twin thruster RBCC rocket ejector system over a single rocket system. The second task investigated 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. Detailed mixing and combustion information was obtained through Raman spectroscopy measurements of major species (gaseous oxygen, hydrogen, nitrogen and water vapor) for the gaseous hydrogen/oxygen rocket ejector experiments.

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.

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.

Rocket Science: The Shuttle's Main Engines, though Old, Are not Forgotten in the New Exploration Initiative

NASA Technical Reports Server (NTRS)

Covault, Craig

2005-01-01

The Space Shuttle Main Engine (SSME), developed 30 years ago, remains a strong candidate for use in the new Exploration Initiative as part of a shuttle-derived heavy-lift expendable booster. This is because the Boeing-Rocket- dyne man-rated SSME remains the most highly efficient liquid rocket engine ever developed. There are only enough parts for 12-15 existing SSMEs, however, so one NASA option is to reinitiate SSME production to use it as a throw-away, as opposed to a reusable, powerplant for NASA s new heavy-lift booster.

Effects of high combustion chamber pressure on rocket noise environment

NASA Technical Reports Server (NTRS)

Pao, S. P.

1972-01-01

The acoustical environment for a high combustion chamber pressure engine was examined in detail, using both conventional and advanced theoretical analysis. The influence of elevated chamber pressure on the rocket noise environment was established, based on increase in exit velocity and flame temperature, and changes in basic engine dimensions. Compared to large rocket engines, the overall sound power level is found to be 1.5 dB higher, if the thrust is the same. The peak Strouhal number shifted about one octave lower to a value near 0.01. Data on apparent sound source location and directivity patterns are also presented.

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.

Demonstrated survivability of a high temperature optical fiber cable on a 1500 pound thrust rocket chamber

NASA Technical Reports Server (NTRS)

Sovie, Amy L.

1992-01-01

A demonstration of the ability of an existing optical fiber cable to survive the harsh environment of a rocket engine was performed at the NASA Lewis Research Center. The intent of this demonstration was to prove the feasibility of applying fiber optic technology to rocket engine instrumentation systems. Extreme thermal transient tests were achieved by wrapping a high temperature optical fiber, which was cablized for mechanical robustness, around the combustion chamber outside wall of a 1500 lb Hydrogen-Oxygen rocket engine. Additionally, the fiber was wrapped around coolant inlet pipes which were subject to near liquid hydrogen temperatures. Light from an LED was sent through the multimode fiber, and output power was monitored as a function of time while the engine was fired. The fiber showed no mechanical damage after 419 firings during which it was subject to transients from 30 K to 350 K, and total exposure time to near liquid hydrogen temperatures in excess of 990 seconds. These extreme temperatures did cause attenuation greater than 3 dB, but the signal was fully recovered at room temperature. This experiment demonstrates that commercially available optical fiber cables can survive the environment seen by a typical rocket engine instrumentation system, and disclose a temperature-dependent attenuation observed during exposure to near liquid hydrogen temperatures.

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.

Space shuttle SRM plume expansion sensitivity analysis. [flow characteristics of exhaust gases from solid propellant rocket engines

NASA Technical Reports Server (NTRS)

Smith, S. D.; Tevepaugh, J. A.; Penny, M. M.

1975-01-01

The exhaust plumes of the space shuttle solid rocket motors can have a significant effect on the base pressure and base drag of the shuttle vehicle. A parametric analysis was conducted to assess the sensitivity of the initial plume expansion angle of analytical solid rocket motor flow fields to various analytical input parameters and operating conditions. The results of the analysis are presented and conclusions reached regarding the sensitivity of the initial plume expansion angle to each parameter investigated. Operating conditions parametrically varied were chamber pressure, nozzle inlet angle, nozzle throat radius of curvature ratio and propellant particle loading. Empirical particle parameters investigated were mean size, local drag coefficient and local heat transfer coefficient. Sensitivity of the initial plume expansion angle to gas thermochemistry model and local drag coefficient model assumptions were determined.

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.

Optimization of the rocket mode trajectory in a rocket based combined cycle (RBCC) engine powered SSTO vehicle

NASA Astrophysics Data System (ADS)

Foster, Richard W.

1989-07-01

The application of rocket-based combined cycle (RBCC) engines to booster-stage propulsion, in combination with all-rocket second stages in orbital-ascent missions, has been studied since the mid-1960s; attention is presently given to the case of the 'ejector scramjet' RBCC configuration's application to SSTO vehicles. While total mass delivered to initial orbit is optimized at Mach 20, payload delivery capability to initial orbit optimizes at Mach 17, primarily due to the reduction of hydrogen fuel tankage structure, insulation, and thermal protection system weights.

Preliminary 2-D shell analysis of the space shuttle solid rocket boosters

NASA Technical Reports Server (NTRS)

Knight, Norman F., Jr.; Gillian, Ronnie E.; Nemeth, Michael P.

1987-01-01

A two-dimensional shell model of an entire solid rocket booster (SRB) has been developed using the STAGSC-1 computer code and executed on the Ames CRAY computer. The purpose of these analyses is to calculate the overall deflection and stress distributions for the SRB when subjected to mechanical loads corresponding to critical times during the launch sequence. The mechanical loading conditions for the full SRB arise from the external tank (ET) attachment points, the solid rocket motor (SRM) pressure load, and the SRB hold down posts. The ET strut loads vary with time after the Space Shuttle main engine (SSME) ignition. The SRM internal pressure varies axially by approximately 100 psi. Static analyses of the full SRB are performed using a snapshot picture of the loads. The field and factory joints are modeled by using equivalent stiffness joints instead of detailed models of the joint. As such, local joint behavior cannot be obtained from this global model.

6. "EXPERIMENTAL ROCKET ENGINE TEST STATION AT AFFTC." A low ...

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

6. "EXPERIMENTAL ROCKET ENGINE TEST STATION AT AFFTC." A low oblique aerial view of Test Area 1-115, looking south, showing Test Stand 1-3 at left, Instrumentation and Control building 8668 at center, and Test Stand 15 at right. The test area is under construction; no evidence of railroad line in photo. - Edwards Air Force Base, Air Force Rocket Propulsion Laboratory, Leuhman Ridge near Highways 58 & 395, Boron, Kern County, CA

48. HISTORIC CLOSEUP VIEW OF THE REDSTONE ROCKET IN THE ...

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

48. HISTORIC CLOSE-UP VIEW OF THE REDSTONE ROCKET IN THE TEST STAND, WITH THE TAIL SECTION REMOVED, REVEALING THE ROCKET ENGINE WITH SOME OF THE TESTING SENSORS ATTACHED. - Marshall Space Flight Center, Redstone Rocket (Missile) Test Stand, Dodd Road, Huntsville, Madison County, AL

NASA Researchers Examine a Pratt and Whitney RL-10 Rocket Engine

NASA Image and Video Library

1962-04-21

Lead Test Engineer John Kobak (right) and a technician use an oscilloscope to test the installation of a Pratt and Whitney RL-10 engine in the Propulsion Systems Laboratory at the National Aeronautics and Space Administration (NASA) Lewis Research Center. In 1955 the military asked Pratt and Whitney to develop hydrogen engines specifically for aircraft. The program was canceled in 1958, but Pratt and Whitney decided to use the experience to develop a liquid-hydrogen rocket engine, the RL-10. Two of the 15,000-pound-thrust RL-10 engines were used to power the new Centaur second-stage rocket. Centaur was designed to carry the Surveyor spacecraft on its mission to soft-land on the Moon. Pratt and Whitney ran into problems while testing the RL-10 at their facilities. NASA Headquarters assigned Lewis the responsibility for investigating the RL-10 problems because of the center’s long history of liquid-hydrogen development. Lewis’ Chemical Rocket Division began a series of tests to study the RL-10 at its Propulsion Systems Laboratory in March 1960. The facility contained two test chambers that could study powerful engines in simulated altitude conditions. The first series of RL-10 tests in early 1961 involved gimballing the engine as it fired. Lewis researchers were able to yaw and pitch the engine to simulate its behavior during a real flight.

Liquid Rocket Engine Testing

NASA Technical Reports Server (NTRS)

Rahman, Shamim

2005-01-01

Comprehensive Liquid Rocket Engine testing is essential to risk reduction for Space Flight. Test capability represents significant national investments in expertise and infrastructure. Historical experience underpins current test capabilities. Test facilities continually seek proactive alignment with national space development goals and objectives including government and commercial sectors.

Measuring System Value in the Ares 1 Rocket Using an Uncertainty-Based Coupling Analysis Approach

NASA Astrophysics Data System (ADS)

Wenger, Christopher

Coupling of physics in large-scale complex engineering systems must be correctly accounted for during the systems engineering process to ensure no unanticipated behaviors or unintended consequences arise in the system during operation. Structural vibration of large segmented solid rocket motors, known as thrust oscillation, is a well-documented problem that can affect the health and safety of any crew onboard. Within the Ares 1 rocket, larger than anticipated vibrations were recorded during late stage flight that propagated from the engine chamber to the Orion crew module. Upon investigation engineers found the root cause to be the structure of the rockets feedback onto fluid flow within the engine. The goal of this paper is to showcase a coupling strength analysis from the field of Multidisciplinary Design Optimization to identify the major impacts that caused the Thrust Oscillation event in the Ares 1. Once identified an uncertainty analysis of the coupled system using an uncertainty based optimization technique is used to identify the likelihood of occurrence for these strong or weak interactions to take place.

AJ26 engine test

NASA Image and Video Library

2011-03-19

A team of engineers from NASA's John C. Stennis Space Center, Orbital Sciences Corporation and Aerojet conduct a successful test of an Aerojet AJ26 rocket engine on March 19. Stennis is testing AJ26 engines for Orbital Sciences to power commercial cargo missions to the International Space Station. Orbital has partnered with NASA through the Commercial Orbital Transportation Services initiative to carry out eight cargo missions to the space station by 2015, using Taurus II rockets.

Progress toward an advanced condition monitoring system for reusable rocket engines

NASA Technical Reports Server (NTRS)

Maram, J.; Barkhoudarian, S.

1987-01-01

A new generation of advanced sensor technologies will allow the direct measurement of critical/degradable rocket engine components' health and the detection of degraded conditions before component deterioration affects engine performance, leading to substantial improvements in reusable engines' operation and maintenance. When combined with a computer-based engine condition-monitoring system, these sensors can furnish a continuously updated data base for the prediction of engine availability and advanced warning of emergent maintenance requirements. Attention is given to the case of a practical turbopump and combustion device diagnostic/prognostic health-monitoring system.

Embedded expert system for space shuttle main engine maintenance

NASA Technical Reports Server (NTRS)

Pooley, J.; Thompson, W.; Homsley, T.; Teoh, W.; Jones, J.; Lewallen, P.

1987-01-01

The SPARTA Embedded Expert System (SEES) is an intelligent health monitoring system that directs analysis by placing confidence factors on possible engine status and then recommends a course of action to an engineer or engine controller. The technique can prevent catastropic failures or costly rocket engine down time because of false alarms. Further, the SEES has potential as an on-board flight monitor for reusable rocket engine systems. The SEES methodology synergistically integrates vibration analysis, pattern recognition and communications theory techniques with an artificial intelligence technique - the Embedded Expert System (EES).

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.

Calculated concentrations of any radionuclide deposited on the ground by release from underground nuclear detonations, tests of nuclear rockets, and tests of nuclear ramjet engines

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

Hicks, H.G.

1981-11-01

This report presents calculated gamma radiation exposure rates and ground deposition of related radionuclides resulting from three types of event that deposited detectable radioactivity outside the Nevada Test Site complex, namely, underground nuclear detonations, tests of nuclear rocket engines and tests of nuclear ramjet engines.

Creative Technology for Schoolchildren.

ERIC Educational Resources Information Center

Stolyarov, Yuri

1981-01-01

Describes creative technology programs for elementary and secondary school children in the Soviet Union. Elementary school projects include aircraft, ship, and rocket models, amateur radio, electrical engineering, and electronics. Senior high school students design and build small-capacity vehicles, agricultural equipment, and electronic…

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.

Pumping Performance or RBCC Engine under Sea Level Static Condition

NASA Astrophysics Data System (ADS)

Kouchi, Toshinori; Tomioka, Sadatake; Kanda, Takeshi

Numerical simulations were conducted to predict the ejector pumping performance of a rocket-ramjet combined-cycle engine under a take-off condition. The numerical simulations revealed that the suction airflow was chocked at the exit of the engine throat when the ejector rocket was driven by cold N2 gas at the chamber pressure of 3MPa. When the ejector-driving gas was changed from cold N2 gas to hot combustion gas, the suction performance decreased remarkably. Mach contours in the engine revealed that the rocket plume constricted when the driving gas was the hot combustion gas. The change of the area of the stream tube area seemed to induce the pressure rise in the duct and decreasing in the pumping performance.

KSC-2010-5768

NASA Image and Video Library

2010-12-03

CAPE CANAVERAL, Fla. -- The SpaceX Falcon 9 rocket awaits a static fire test on Space Launch Complex-40 at Cape Canaveral Air Force Station, in which all nine Merlin engines will fire at once. The engines use rocket-grade kerosene and liquid oxygen to produce 1 million pounds of thrust. After the test, SpaceX will conduct a thorough review of all data as engineers make final preparations for the first launch of the Commercial Orbital Transportation Services (COTS) Dragon spacecraft to low Earth orbit atop the Falcon 9. This first stage firing is part of a full launch dress rehearsal, which will end after the engines fire at full power for two seconds, with only the hold-down system restraining the rocket from flight. Photo credit: NASA/Rusty Backer

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.

Real-Time Rocket/Vehicle System Integrated Health Management Laboratory For Development and Testing of Health Monitoring/Management Systems

NASA Technical Reports Server (NTRS)

Aguilar, R.

2006-01-01

Pratt & Whitney Rocketdyne has developed a real-time engine/vehicle system integrated health management laboratory, or testbed, for developing and testing health management system concepts. This laboratory simulates components of an integrated system such as the rocket engine, rocket engine controller, vehicle or test controller, as well as a health management computer on separate general purpose computers. These general purpose computers can be replaced with more realistic components such as actual electronic controllers and valve actuators for hardware-in-the-loop simulation. Various engine configurations and propellant combinations are available. Fault or failure insertion capability on-the-fly using direct memory insertion from a user console is used to test system detection and response. The laboratory is currently capable of simulating the flow-path of a single rocket engine but work is underway to include structural and multiengine simulation capability as well as a dedicated data acquisition system. The ultimate goal is to simulate as accurately and realistically as possible the environment in which the health management system will operate including noise, dynamic response of the engine/engine controller, sensor time delays, and asynchronous operation of the various components. The rationale for the laboratory is also discussed including limited alternatives for demonstrating the effectiveness and safety of a flight system.

Liquid rocket engine turbines

NASA Technical Reports Server (NTRS)

1974-01-01

Criteria for the design and development of turbines for rocket engines to meet specific performance, and installation requirements are summarized. The total design problem, and design elements are identified, and the current technology pertaining to these elements is described. Recommended practices for achieving a successful design are included.

Centrifugal pumps for rocket engines

NASA Technical Reports Server (NTRS)

Campbell, W. E.; Farquhar, J.

1974-01-01

The use of centrifugal pumps for rocket engines is described in terms of general requirements of operational and planned systems. Hydrodynamic and mechanical design considerations and techniques and test procedures are summarized. Some of the pump development experiences, in terms of both problems and solutions, are highlighted.

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.

A survey of instabilities within centrifugal pumps and concepts for improving the flow range of pumps in rocket engines

NASA Technical Reports Server (NTRS)

Veres, Joseph P.

1992-01-01

Design features and concepts that have primary influence on the stable operating flow range of propellant-feed centrifugal turbopumps in a rocket engine are discussed. One of the throttling limitations of a pump-fed rocket engine is the stable operating range of the pump. Several varieties of pump hydraulic instabilities are mentioned. Some pump design criteria are summarized and a qualitative correlation of key parameters to pump stall and surge are referenced. Some of the design criteria were taken from the literature on high pressure ratio centrifugal compressors. Therefore, these have yet to be validated for extending the stable operating flow range of high-head pumps. Casing treatment devices, dynamic fluid-damping plenums, backflow-stabilizing vanes and flow-reinjection techniques are summarized. A planned program was undertaken at LeRC to validate these concepts. Technologies developed by this program will be available for the design of turbopumps for advanced space rocket engines for use by NASA in future space missions where throttling is essential.

20. Building 202, detail of stand A, rocket test stand ...

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

20. Building 202, detail of stand A, rocket test stand in test cell. View looking southeast. - Rocket Engine Testing Facility, GRC Building No. 202, NASA Glenn Research Center, Cleveland, Cuyahoga County, OH

Multiphase Modeling of Water Injection on Flame Deflector

NASA Technical Reports Server (NTRS)

Vu, Bruce T.; Bachchan, Nili; Peroomian, Oshin; Akdag, Vedat

2013-01-01

This paper describes the use of an Eulerian Dispersed Phase (EDP) model to simulate the water injected from the flame deflector and its interaction with supersonic rocket exhaust from a proposed Space Launch System (SLS) vehicle. The Eulerian formulation, as part of the multi-phase framework, is described. The simulations show that water cooling is only effective over the region under the liquid engines. Likewise, the water injection provides only minor effects over the surface area under the solid engines.

Modeling Liquid Rocket Engine Atomization and Swirl/Coaxial Injectors

DTIC Science & Technology

2008-02-27

47-61, 2004. 2. Yoon, S . S ., and Heister, S . D., "A Fully Nonlinear Model for Atomization of High - Speed Jets," Engineering Analysis with... Power , V20, pp 468-479, 2004. 5. Yoon, S . S ., and Heister, S . D., "Analytic Solutions for Computing Velocities Induced from Potential Vortex Ring...Heister, S . D., "Three Dimensional Flow Simulations in Recessed Region of a Coaxial Injector," J. Propulsion and Power , V21, No.4, pp. 728-742

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.

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.

Five Pioneers with Scale Models of Their Missiles

NASA Technical Reports Server (NTRS)

1950-01-01

Five pioneers pose with scale models of their missiles they created in the 1950s. From left to right: Dr. Ernst Stuhlinger, a member of the original German rocket team who directed the Research Projects Office, Army Ballistic Missile Agency (ABMA); Major General Holger Toftoy, who consolidated U.S. missile and rocketry development; Professor Herman Oberth, a rocket pioneer and Dr. von Braun's mentor; Dr. Wernher von Braun, Director, Development Operation Division, ABMA; and Dr. Robert Lusser, who served as assistant director for Reliability Engineering for ABMA. This photographis was taken February 1, 1956 by Hank Walker and appeared in February 27, 1956 issue of Life magazine.

EDIN design study alternate space shuttle booster replacement concepts. Volume 1: Engineering analysis

NASA Technical Reports Server (NTRS)

Demakes, P. T.; Hirsch, G. N.; Stewart, W. A.; Glatt, C. R.

1976-01-01

The use of a recoverable liquid rocket booster (LRB) system to replace the existing solid rocket booster (SRB) system for the shuttle was studied. Historical weight estimating relationships were developed for the LRB using Saturn technology and modified as required. Mission performance was computed using February 1975 shuttle configuration groundrules to allow reasonable comparison of the existing shuttle with the study designs. The launch trajectory was constrained to pass through both the RTLS/AOA and main engine cut off points of the shuttle reference mission 1. Performance analysis is based on a point design trajectory model which optimizes initial tilt rate and exoatmospheric pitch profile. A gravity turn was employed during the boost phase in place of the shuttle angle of attack profile. Engine throttling add/or shutdown was used to constrain dynamic pressure and/or longitudinal acceleration where necessary. Four basic configurations were investigated: a parallel burn vehicle with an F-1 engine powered LRB; a parallel burn vehicle with a high pressure engine powered LRB; a series burn vehicle with a high pressure engine powered LRB. The relative sizes of the LRB and the ET are optimized to minimize GLOW in most cases.

Analysis of a New Rocket-Based Combined-Cycle Engine Concept at Low Speed

NASA Technical Reports Server (NTRS)

Yungster, S.; Trefny, C. J.

1999-01-01

An analysis of the Independent Ramjet Stream (IRS) cycle is presented. The IRS cycle is a variation of the conventional ejector-Ramjet, and is used at low speed in a rocket-based combined-cycle (RBCC) propulsion system. In this new cycle, complete mixing between the rocket and ramjet streams is not required, and a single rocket chamber can be used without a long mixing duct. Furthermore, this concept allows flexibility in controlling the thermal choke process. The resulting propulsion system is intended to be simpler, more robust, and lighter than an ejector-ramjet. The performance characteristics of the IRS cycle are analyzed for a new single-stage-to-orbit (SSTO) launch vehicle concept, known as "Trailblazer." The study is based on a quasi-one-dimensional model of the rocket and air streams at speeds ranging from lift-off to Mach 3. The numerical formulation is described in detail. A performance comparison between the IRS and ejector-ramjet cycles is also presented.

Development of Modeling Approaches for Nuclear Thermal Propulsion Test Facilities

NASA Technical Reports Server (NTRS)

Jones, Daniel R.; Allgood, Daniel C.; Nguyen, Ke

2014-01-01

High efficiency of rocket propul-sion systems is essential for humanity to venture be-yond the moon. Nuclear Thermal Propulsion (NTP) is a promising alternative to conventional chemical rock-ets with relatively high thrust and twice the efficiency of the Space Shuttle Main Engine. NASA is in the pro-cess of developing a new NTP engine, and is evaluat-ing ground test facility concepts that allow for the thor-ough testing of NTP devices. NTP engine exhaust, hot gaseous hydrogen, is nominally expected to be free of radioactive byproducts from the nuclear reactor; how-ever, it has the potential to be contaminated due to off-nominal engine reactor performance. Several options are being investigated to mitigate this hazard potential with one option in particular that completely contains the engine exhaust during engine test operations. The exhaust products are subsequently disposed of between engine tests. For this concept (see Figure 1), oxygen is injected into the high-temperature hydrogen exhaust that reacts to produce steam, excess oxygen and any trace amounts of radioactive noble gases released by off-nominal NTP engine reactor performance. Water is injected to condense the potentially contaminated steam into water. This water and the gaseous oxygen (GO2) are subsequently passed to a containment area where the water and GO2 are separated into separate containment tanks.

Large liquid rocket engine transient performance simulation system

NASA Technical Reports Server (NTRS)

Mason, J. R.; Southwick, R. D.

1989-01-01

Phase 1 of the Rocket Engine Transient Simulation (ROCETS) program consists of seven technical tasks: architecture; system requirements; component and submodel requirements; submodel implementation; component implementation; submodel testing and verification; and subsystem testing and verification. These tasks were completed. Phase 2 of ROCETS consists of two technical tasks: Technology Test Bed Engine (TTBE) model data generation; and system testing verification. During this period specific coding of the system processors was begun and the engineering representations of Phase 1 were expanded to produce a simple model of the TTBE. As the code was completed, some minor modifications to the system architecture centering on the global variable common, GLOBVAR, were necessary to increase processor efficiency. The engineering modules completed during Phase 2 are listed: INJTOO - main injector; MCHBOO - main chamber; NOZLOO - nozzle thrust calculations; PBRNOO - preburner; PIPE02 - compressible flow without inertia; PUMPOO - polytropic pump; ROTROO - rotor torque balance/speed derivative; and TURBOO - turbine. Detailed documentation of these modules is in the Appendix. In addition to the engineering modules, several submodules were also completed. These submodules include combustion properties, component performance characteristics (maps), and specific utilities. Specific coding was begun on the system configuration processor. All functions necessary for multiple module operation were completed but the SOLVER implementation is still under development. This system, the Verification Checkout Facility (VCF) allows interactive comparison of module results to store data as well as provides an intermediate checkout of the processor code. After validation using the VCF, the engineering modules and submodules were used to build a simple TTBE.

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.

Nuclear Physics Made Very, Very Easy

NASA Technical Reports Server (NTRS)

Hanlen, D. F.; Morse, W. J.

1968-01-01

The fundamental approach to nuclear physics was prepared to introduce basic reactor principles to various groups of non-nuclear technical personnel associated with NERVA Test Operations. NERVA Test Operations functions as the field test group for the Nuclear Rocket Engine Program. Nuclear Engine for Rocket Vehicle Application (NERVA) program is the combined efforts of Aerojet-General Corporation as prime contractor, and Westinghouse Astronuclear Laboratory as the major subcontractor, for the assembly and testing of nuclear rocket engines. Development of the NERVA Program is under the direction of the Space Nuclear Propulsion Office, a joint agency of the U.S. Atomic Energy Commission and the National Aeronautics and Space Administration.

Design and Evaluation of Dual-Expander Aerospike Nozzle Upper Stage Engine

DTIC Science & Technology

2014-09-18

Nozzle , taken from Martin [2] . . . . . 19 2.3 Typical Liquid Rocket Engine Cycles from Huzel and Huang[3], credit J. Hall[4] 21 2.4 Liquid Rocket Engine...giving the maximum thrust. For steady, supersonic flow (no separation from the nozzle ) the exit pressure is constant for a given engine plus nozzle ...performance independent of a rocket’s nozzle . Assuming one-dimensional, steady, and isentropic flow of a perfect gas gives the definition for characteristic

High-Temperature Rocket Engine

NASA Technical Reports Server (NTRS)

Schneider, Steven J.; Rosenberg, Sanders D.; Chazen, Melvin L.

1994-01-01

Two rocket engines that operate at temperature of 2,500 K designed to provide thrust for station-keeping adjustments of geosynchronous satellites, for raising and lowering orbits, and for changing orbital planes. Also useful as final propulsion stages of launch vehicles delivering small satellites to low orbits around Earth. With further development, engines used on planetary exploration missions for orbital maneuvers. High-temperature technology of engines adaptable to gas-turbine combustors, ramjets, scramjets, and hot components of many energy-conversion systems.

KSC-2012-6222

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 and Rocket University project leads Kyle Dixon, left, and Evelyn Orozco-Smith check 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

NASA on a Strong Roll in Preparing Space Launch System Flight Engines

NASA Image and Video Library

2017-08-09

NASA is on a roll when it comes to testing engines for its new Space Launch System (SLS) rocket that will send astronauts to deep-space destinations, including Mars. Just two weeks after the third test of a new RS-25 engine flight controller, the space agency recorded its fourth full-duration controller test Aug. 9 at Stennis Space Center near Bay St. Louis, Mississippi. Engineers conducted a 500-second test of the RS-25 engine controller on the A-1 Test Stand at Stennis. The test involved installing the controller on an RS-25 development engine and firing it in the same manner, and for the same length of time, as needed during an actual SLS launch. The test marked another milestone toward launch of the first integrated flight of the SLS rocket and Orion crew vehicle. Exploration Mission-1 will be an uncrewed mission into lunar orbit, designed to provide a final check-out test of rocket and Orion capabilities before astronauts are returned to deep space. The SLS rocket will be powered at launch by four RS-25 engines, providing a combined 2 million pounds of thrust, and with a pair of solid rocket boosters, providing more than 8 million pounds of total thrust. The RS-25 engines for the initial SLS flights are former space shuttle main engines that are now being used to launch the larger and heavier SLS rocket and with the new controller. The controller is a critical component that operates as the engine “brain” that communicates with SLS flight computers to receive operation performance commands and to provide diagnostic data on engine health and status. Engineers conducted early prototype tests at Stennis to collect data for development of the new controller by NASA, RS-25 prime contractor Aerojet Rocketdyne and subcontractor Honeywell. Testing of actual flight controllers began at Stennis in March. NASA is testing all controllers and engines designated for the EM-1 flight at Stennis. It also will test the SLS core stage for the flight at Stennis, which will involve installing the stage on the B-2 Test Stand and firing its four RS-25 engines simultaneously, as during an actual launch. RS-25 tests at Stennis are conducted by a team of NASA, Aerojet Rocketdyne and Syncom Space Services engineers and operators. Aerojet Rocketdyne is the RS-25 prime contractor. Syncom Space Services is the prime contractor for Stennis facilities and operations.

Integrated Main Propulsion System Performance Reconstruction Process/Models

NASA Technical Reports Server (NTRS)

Lopez, Eduardo; Elliott, Katie; Snell, Steven; Evans, Michael

2013-01-01

The Integrated Main Propulsion System (MPS) Performance Reconstruction process provides the MPS post-flight data files needed for postflight reporting to the project integration management and key customers to verify flight performance. This process/model was used as the baseline for the currently ongoing Space Launch System (SLS) work. The process utilizes several methodologies, including multiple software programs, to model integrated propulsion system performance through space shuttle ascent. It is used to evaluate integrated propulsion systems, including propellant tanks, feed systems, rocket engine, and pressurization systems performance throughout ascent based on flight pressure and temperature data. The latest revision incorporates new methods based on main engine power balance model updates to model higher mixture ratio operation at lower engine power levels.

Stennis Space Center goes to Washington Folklife Festival

NASA Image and Video Library

2008-07-03

Bryon Maynard (left), an aerospace technologist for Propulsion Systems & Tech in Stennis' Engineering and Science Directorate, uses a 'pocket rocket' to demonstrate the concept of rocket propulsion as part of NASA's exhibit at the Smithsonian Folklife Festival in Washington, D.C. Maynard is joined by Bradley Messer (right), chief of the Systems Engineering & Integration Division in Stennis' Engineering and Science Directorate, and a pair of exhibit visitors.

Stennis Space Center goes to Washington Folklife Festival

NASA Technical Reports Server (NTRS)

2008-01-01

Bryon Maynard (left), an aerospace technologist for Propulsion Systems & Tech in Stennis' Engineering and Science Directorate, uses a 'pocket rocket' to demonstrate the concept of rocket propulsion as part of NASA's exhibit at the Smithsonian Folklife Festival in Washington, D.C. Maynard is joined by Bradley Messer (right), chief of the Systems Engineering & Integration Division in Stennis' Engineering and Science Directorate, and a pair of exhibit visitors.

Analysis of rocket engine injection combustion processes

NASA Technical Reports Server (NTRS)

Salmon, J. W.; Saltzman, D. H.

1977-01-01

Mixing methodology improvement for the JANNAF DER and CICM injection/combustion analysis computer programs was accomplished. ZOM plane prediction model development was improved for installation into the new standardized DER computer program. An intra-element mixing model developing approach was recommended for gas/liquid coaxial injection elements for possible future incorporation into the CICM computer program.

Electrophoretic separator for purifying biologicals, part 1

NASA Technical Reports Server (NTRS)

Mccreight, L. R.

1978-01-01

A program to develop an engineering model of an electrophoretic separator for purifying biologicals is summarized. An extensive mathematical modeling study and numerous ground based tests were included. Focus was placed on developing an actual electrophoretic separator of the continuous flow type, configured and suitable for flight testing as a space processing applications rocket payload.

Turbopump systems for liquid rocket engines

NASA Technical Reports Server (NTRS)

1974-01-01

The turbopump system, from preliminary design through rocket engine testing is examined. Selection of proper system type for each application and integration of the components into a working system are dealt with. Details are also given on the design of various components including inducers, pumps, turbines, gears, and bearings.

A Combustion Research Facility for Testing Advanced Materials for Space Applications

NASA Technical Reports Server (NTRS)

Bur, Michael J.

2003-01-01

The test facility presented herein uses a groundbased rocket combustor to test the durability of new ceramic composite and metallic materials in a rocket engine thermal environment. A gaseous H2/02 rocket combustor (essentially a ground-based rocket engine) is used to generate a high temperature/high heat flux environment to which advanced ceramic and/or metallic materials are exposed. These materials can either be an integral part of the combustor (nozzle, thrust chamber etc) or can be mounted downstream of the combustor in the combustor exhaust plume. The test materials can be uncooled, water cooled or cooled with gaseous hydrogen.

High-End Computing Challenges in Aerospace Design and Engineering

NASA Technical Reports Server (NTRS)

Bailey, F. Ronald

2004-01-01

High-End Computing (HEC) has had significant impact on aerospace design and engineering and is poised to make even more in the future. In this paper we describe four aerospace design and engineering challenges: Digital Flight, Launch Simulation, Rocket Fuel System and Digital Astronaut. The paper discusses modeling capabilities needed for each challenge and presents projections of future near and far-term HEC computing requirements. NASA's HEC Project Columbia is described and programming strategies presented that are necessary to achieve high real performance.

Walter Thiel—Short life of a rocket scientist

NASA Astrophysics Data System (ADS)

Thiel, Karen; Przybilski, Olaf

2013-10-01

In 2012 we celebrate the 70th anniversary of the first successful rocket launch that reached a height of 84.5 km and had a speed of 4.824 km/h (5x sonic speed). This rocket flew 190 km to the target location. One of the masterminds of this launch was Walter Thiel, a German chemist and rocket engineer. Thiel was highly talented, during his education from primary school until diploma exams he always received a grade of A in his exams. He was called "the student with the 7 A grades". In 1934 Thiel became Dr.-Ing. (chem.), with the highest possible honor (summa cum laude), when he was only 24 years old. He started to work for the rocket development department at Humboldt University, Berlin. Walter Dornberger asked him to leave the university research department and become head of rocket propulsion development in his team in Kummersdorf, near Berlin. Thiel's groundbreaking ideas for the rocket engine would lead to a significant reduction in material, weight and work processes, as well as a shortening in the length of the engine itself. Thiel and his team also defined the fuel itself and the best ratio of mixture between ethanol and liquid oxygen for the engine. In 1940 the propulsion team moved from Kummersdorf to Peenemünde after the launch sites were completed there. Thiel became deputy of Wernher von Braun at the R&D units. One of Thiel's team members was Konrad Dannenberg, who later became famous in the development of the Saturn program. On the night from August 17 to August 18, 1943, Thiel and his family (wife and two children) were killed during a Royal Air Force bombing raid (Operation Hydra). The Moon crater "Thiel" on the far side of the Moon is named after Walter Thiel. The research results of Walter Thiel had a strong impact on the United States' rocket program as well as the Russian rocket development program.

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.

Determination of Combustion Product Radicals in a Hydrocarbon Fueled Rocket Exhaust Plume

NASA Technical Reports Server (NTRS)

Langford, Lester A.; Allgood, Daniel C.; Junell, Justin C.

2007-01-01

The identification of metallic effluent materials in a rocket engine exhaust plume indicates the health of the engine. Since 1989, emission spectroscopy of the plume of the Space Shuttle Main Engine (SSME) has been used for ground testing at NASA's Stennis Space Center (SSC). This technique allows the identification and quantification of alloys from the metallic elements observed in the plume. With the prospect of hydrocarbon-fueled rocket engines, such as Rocket Propellant 1 (RP-1) or methane (CH4) fueled engines being considered for use in future space flight systems, the contributions of intermediate or final combustion products resulting from the hydrocarbon fuels are of great interest. The effect of several diatomic molecular radicals, such as Carbon Dioxide , Carbon Monoxide, Molecular Carbon, Methylene Radical, Cyanide or Cyano Radical, and Nitric Oxide, needs to be identified and the effects of their band systems on the spectral region from 300 nm to 850 nm determined. Hydrocarbon-fueled rocket engines will play a prominent role in future space exploration programs. Although hydrogen fuel provides for higher engine performance, hydrocarbon fuels are denser, safer to handle, and less costly. For hydrocarbon-fueled engines using RP-1 or CH4 , the plume is different from a hydrogen fueled engine due to the presence of several other species, such as CO2, C2, CO, CH, CN, and NO, in the exhaust plume, in addition to the standard H2O and OH. These species occur as intermediate or final combustion products or as a result of mixing of the hot plume with the atmosphere. Exhaust plume emission spectroscopy has emerged as a comprehensive non-intrusive sensing technology which can be applied to a wide variety of engine performance conditions with a high degree of sensitivity and specificity. Stennis Space Center researchers have been in the forefront of advancing experimental techniques and developing theoretical approaches in order to bring this technology to a more mature stage.

J-2X engine assembly

NASA Image and Video Library

2011-03-03

Pratt & Whitney Rocketdyne employees Carlos Alfaro (l) and Oliver Swanier work on the main combustion element of the J-2X rocket engine at their John C. Stennis Space Center facility. Assembly of the J-2X rocket engine to be tested at the site is under way, with completion and delivery to the A-2 Test Stand set for June. The J-2X is being developed as a next-generation engine that can carry humans into deep space. Stennis Space Center is preparing a trio of stands to test the new engine.

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.

J-2X engine

NASA Image and Video Library

2012-09-14

NASA engineers continued to collect test performance data on the new J-2X rocket engine at Stennis Space Center with a 250-second test Sept. 14. The test on the A-2 Test Stand was the 19th in a series of firings to gather critical data for continued development of the engine. The J-2X is being developed by Pratt and Whitney Rocketdyne for NASA's Marshall Space Flight Center in Huntsville, Ala. It is the first liquid oxygen and liquid hydrogen rocket engine rated to carry humans into space to be developed in 40 years.

Combustion dynamics in cryogenic rocket engines: Research programme at DLR Lampoldshausen

NASA Astrophysics Data System (ADS)

Hardi, Justin S.; Traudt, Tobias; Bombardieri, Cristiano; Börner, Michael; Beinke, Scott K.; Armbruster, Wolfgang; Nicolas Blanco, P.; Tonti, Federica; Suslov, Dmitry; Dally, Bassam; Oschwald, Michael

2018-06-01

The Combustion Dynamics group in the Rocket Propulsion Department at the German Aerospace Center (DLR), Lampoldshausen, strives to advance the understanding of dynamic processes in cryogenic rocket engines. Leveraging the test facilities and experimental expertise at DLR Lampoldshausen, the group has taken a primarily experimental approach to investigating transient flows, ignition, and combustion instabilities for over one and a half decades. This article provides a summary of recent achievements, and an overview of current and planned research activities.

Experimental Flow Models for SSME Flowfield Characterization

NASA Technical Reports Server (NTRS)

Abel, L. C.; Ramsey, P. E.

1989-01-01

Full scale flow models with extensive instrumentation were designed and manufactured to provide data necessary for flow field characterization in rocket engines of the Space Shuttle Main Engine (SSME) type. These models include accurate flow path geometries from the pre-burner outlet through the throat of the main combustion chamber. The turbines are simulated with static models designed to provide the correct pressure drop and swirl for specific power levels. The correct turbopump-hot gas manifold interfaces were designed into the flow models to permit parametric/integration studies for new turbine designs. These experimental flow models provide a vehicle for understanding the fluid dynamics associated with specific engine issues and also fill the more general need for establishing a more detailed fluid dynamic base to support development and verification of advanced math models.

Rocket Engine Numerical Simulator (RENS)

NASA Technical Reports Server (NTRS)

Davidian, Kenneth O.

1997-01-01

Work is being done at three universities to help today's NASA engineers use the knowledge and experience of their Apolloera predecessors in designing liquid rocket engines. Ground-breaking work is being done in important subject areas to create a prototype of the most important functions for the Rocket Engine Numerical Simulator (RENS). The goal of RENS is to develop an interactive, realtime application that engineers can utilize for comprehensive preliminary propulsion system design functions. RENS will employ computer science and artificial intelligence research in knowledge acquisition, computer code parallelization and objectification, expert system architecture design, and object-oriented programming. In 1995, a 3year grant from the NASA Lewis Research Center was awarded to Dr. Douglas Moreman and Dr. John Dyer of Southern University at Baton Rouge, Louisiana, to begin acquiring knowledge in liquid rocket propulsion systems. Resources of the University of West Florida in Pensacola were enlisted to begin the process of enlisting knowledge from senior NASA engineers who are recognized experts in liquid rocket engine propulsion systems. Dr. John Coffey of the University of West Florida is utilizing his expertise in interviewing and concept mapping techniques to encode, classify, and integrate information obtained through personal interviews. The expertise extracted from the NASA engineers has been put into concept maps with supporting textual, audio, graphic, and video material. A fundamental concept map was delivered by the end of the first year of work and the development of maps containing increasing amounts of information is continuing. Find out more information about this work at the Southern University/University of West Florida. In 1996, the Southern University/University of West Florida team conducted a 4day group interview with a panel of five experts to discuss failures of the RL10 rocket engine in conjunction with the Centaur launch vehicle. The discussion was recorded on video and audio tape. Transcriptions of the entire proceedings and an abbreviated video presentation of the discussion highlights are under development. Also in 1996, two additional 3year grants were awarded to conduct parallel efforts that would complement the work being done by Southern University and the University of West Florida. Dr. Prem Bhalla of Jackson State University in Jackson, Mississippi, is developing the architectural framework for RENS. By employing the Rose Rational language and Booch Object Oriented Programming (OOP) technology, Dr. Bhalla is developing the basic structure of RENS by identifying and encoding propulsion system components, their individual characteristics, and cross-functionality and dependencies. Dr. Ruknet Cezzar of Hampton University, located in Hampton, Virginia, began working on the parallelization and objectification of rocket engine analysis and design codes. Dr. Cezzar will use the Turbo C++ OOP language to translate important liquid rocket engine computer codes from FORTRAN and permit their inclusion into the RENS framework being developed at Jackson State University. The Southern University/University of West Florida grant was extended by 1 year to coordinate the conclusion of all three efforts in 1999.

Materials for Liquid Propulsion Systems. Chapter 12

NASA Technical Reports Server (NTRS)

Halchak, John A.; Cannon, James L.; Brown, Corey

2016-01-01

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

NASA Engineers Test Combustion Chamber to Advance 3-D Printed Rocket Engine Design

NASA Image and Video Library

2016-12-08

A series of test firings like this one in late August brought a group of engineers at NASA's Marshall Space Flight Center in Huntsville, Alabama, a big step closer to their goal of a 100-percent 3-D printed rocket engine, said Andrew Hanks, test lead for the additively manufactured demonstration engine project. The main combustion chamber, fuel turbopump, fuel injector, valves and other components used in the tests were of the team's new design, and all major engine components except the main combustion chamber were 3-D printed. (NASA/MSFC)

Potential climate impact of black carbon emitted by rockets

NASA Astrophysics Data System (ADS)

Ross, Martin; Mills, Michael; Toohey, Darin

2010-12-01

A new type of hydrocarbon rocket engine is expected to power a fleet of suborbital rockets for commercial and scientific purposes in coming decades. A global climate model predicts that emissions from a fleet of 1000 launches per year of suborbital rockets would create a persistent layer of black carbon particles in the northern stratosphere that could cause potentially significant changes in the global atmospheric circulation and distributions of ozone and temperature. Tropical stratospheric ozone abundances are predicted to change as much as 1%, while polar ozone changes by up to 6%. Polar surface temperatures change as much as one degree K regionally with significant impacts on polar sea ice fractions. After one decade of continuous launches, globally averaged radiative forcing from the black carbon would exceed the forcing from the emitted CO2 by a factor of about 105 and would be comparable to the radiative forcing estimated from current subsonic aviation.

Graduate students Chris Hill and Ryan Anderson examine a cross section of the prototype rocket engine igniter.

NASA Image and Video Library

2017-09-08

Majid Babai along with Dr. Judy Schneider, and graduate students Chris Hill and Ryan Anderson examine a cross section of the prototype rocket engine igniter created by an innovative bi-metallic 3-D printing advanced manufacturing process under a microscope.

An intelligent control system for rocket engines - Need, vision, and issues

NASA Technical Reports Server (NTRS)

Lorenzo, Carl F.; Merrill, Walter C.

1991-01-01

Several components of intelligence are defined. Within the context of these definitions an intelligent control system for rocket engines is described. The description includes a framework for development of an intelligent control system, including diagnostics, coordination, and direct control. Some current results and issues are presented.

31. Historic view of Building 202 test stand A with ...

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

31. Historic view of Building 202 test stand A with rocket engine, November 19, 1957. On file at NASA Plumbrook Research Center, Sandusky, Ohio. NASA GRC photo number C-46491. - Rocket Engine Testing Facility, GRC Building No. 202, NASA Glenn Research Center, Cleveland, Cuyahoga County, OH

4. Historic photo of fuel and oxidant tanks in hilltop ...

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

4. Historic photo of fuel and oxidant tanks in hilltop area of rocket engine test facility. 1956. On file at NASA Plumbrook Research Center, Sandusky, Ohio. NASA GRC photo number C-1956-160D. - Rocket Engine Testing Facility, NASA Glenn Research Center, Cleveland, Cuyahoga County, OH

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.

High-temperature, high-pressure optical port for rocket engine applications

NASA Technical Reports Server (NTRS)

Delcher, Ray; Nemeth, ED; Powers, W. T.

1993-01-01

This paper discusses the design, fabrication, and test of a window assembly for instrumentation of liquid-fueled rocket engine hot gas systems. The window was designed to allow optical measurements of hot gas in the SSME fuel preburner and appears to be the first window designed for application in a rocket engine hot gas system. Such a window could allow the use of a number of remote optical measurement technologies including: Raman temperature and species concentration measurement, Raleigh temperature measurements, flame emission monitoring, flow mapping, laser-induced florescence, and hardware imaging during engine operation. The window assembly has been successfully tested to 8,000 psi at 1000 F and over 11,000 psi at room temperature. A computer stress analysis shows the window will withstand high temperature and cryogenic thermal shock.

KSC-2010-5772

NASA Image and Video Library

2010-12-03

CAPE CANAVERAL, Fla. -- The SpaceX Falcon 9 rocket awaits a static fire test on Space Launch Complex-40 at Cape Canaveral Air Force Station, in which all nine Merlin engines will fire at once. The engines use rocket-grade kerosene and liquid oxygen to produce 1 million pounds of thrust. After the test, SpaceX will conduct a thorough review of all data as engineers make final preparations for the first launch of the Commercial Orbital Transportation Services (COTS) Dragon spacecraft to low Earth orbit atop the Falcon 9. This first stage firing is part of a full launch dress rehearsal, which will end after the engines fire at full power for two seconds, with only the hold-down system restraining the rocket from flight. Photo credit: NASA/Tony Gray and Kevin O'Connell

KSC-2010-5774

NASA Image and Video Library

2010-12-04

CAPE CANAVERAL, Fla. -- During a static fire test on Space Launch Complex-40 at Cape Canaveral Air Force Station, all nine Merlin engines of the SpaceX Falcon 9 rocket fire at once. The engines use rocket-grade kerosene and liquid oxygen to produce 1 million pounds of thrust. After the test, SpaceX began to conduct a thorough review of all data as engineers make final preparations for the first launch of the Commercial Orbital Transportation Services (COTS) Dragon spacecraft to low Earth orbit atop the Falcon 9. This first stage firing is part of a full launch dress rehearsal, which ended after the engines fired at full power for two seconds, with only the hold-down system restraining the rocket from flight. Photo credit: NASA/Tony Gray and Kevin O'Connell

KSC-2010-5775

NASA Image and Video Library

2010-12-04

CAPE CANAVERAL, Fla. -- During a static fire test on Space Launch Complex-40 at Cape Canaveral Air Force Station, all nine Merlin engines of the SpaceX Falcon 9 rocket fire at once. The engines use rocket-grade kerosene and liquid oxygen to produce 1 million pounds of thrust. After the test, SpaceX began to conduct a thorough review of all data as engineers make final preparations for the first launch of the Commercial Orbital Transportation Services (COTS) Dragon spacecraft to low Earth orbit atop the Falcon 9. This first stage firing is part of a full launch dress rehearsal, which ended after the engines fired at full power for two seconds, with only the hold-down system restraining the rocket from flight. Photo credit: NASA/Tony Gray and Kevin O'Connell

KSC-2010-5776

NASA Image and Video Library

2010-12-04

CAPE CANAVERAL, Fla. -- During a static fire test on Space Launch Complex-40 at Cape Canaveral Air Force Station, all nine Merlin engines of the SpaceX Falcon 9 rocket fire at once. The engines use rocket-grade kerosene and liquid oxygen to produce 1 million pounds of thrust. After the test, SpaceX began to conduct a thorough review of all data as engineers make final preparations for the first launch of the Commercial Orbital Transportation Services (COTS) Dragon spacecraft to low Earth orbit atop the Falcon 9. This first stage firing is part of a full launch dress rehearsal, which ended after the engines fired at full power for two seconds, with only the hold-down system restraining the rocket from flight. Photo credit: NASA/Tony Gray and Kevin O'Connell

Developing Avionics Hardware and Software for Rocket Engine Testing

NASA Technical Reports Server (NTRS)

Aberg, Bryce Robert

2014-01-01

My summer was spent working as an intern at Kennedy Space Center in the Propulsion Avionics Branch of the NASA Engineering Directorate Avionics Division. The work that I was involved with was part of Rocket University's Project Neo, a small scale liquid rocket engine test bed. I began by learning about the layout of Neo in order to more fully understand what was required of me. I then developed software in LabView to gather and scale data from two flowmeters and integrated that code into the main control software. Next, I developed more LabView code to control an igniter circuit and integrated that into the main software, as well. Throughout the internship, I performed work that mechanics and technicians would do in order to maintain and assemble the engine.

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.

Experimental/Analytical Characterization of the RBCC Rocket-Ejector Mode

NASA Technical Reports Server (NTRS)

Ruf, J. H.; Lehman, M.; Pal, S.; Santoro, R. J.

2000-01-01

The experimental/analytical research work described here addresses the rocket-ejector mode (Mach 0-2 operational range) of the RBCC engine. The experimental phase of the program includes studying the mixing and combustion characteristics of the rocket-ejector system utilizing state-of-the-art diagnostic techniques. A two-dimensional variable geometry rocket-ejector system with enhanced optical access was utilized as the experimental platform. The goals of the experimental phase of the research being conducted at Penn State are to: (a) systematically increase the range of rocket-ejector understanding over a wide range of flow/geometry parameters and (b) provide a comprehensive data base for evaluating and anchoring CFD codes. Concurrent with the experimental activities, a CFD code benchmarking effort at Marshall Space Flight Center is also being used to further investigate the RBCC rocket-ejector mode. Experiments involving the single rocket based optically-accessible rocket-ejector system have been conducted for Diffusion and Afterburning (DAB) as well as Simultaneous Mixing and Combustion configurations. For the DAB configuration, air is introduced (direct-connect) or ejected (sea-level static) into a constant area mixer section with a centrally located gaseous oxygen (GO2)/gaseous hydrogen (GH2) rocket combustor. The downstream flowpath for this configuration includes a diffuser, an afterburner and a final converging nozzle. For the SMC configuration, the rocket is centrally located in a slightly divergent duct. For all tested configurations, global measurements of the axial pressure and heat transfer profiles as well as the overall engine thrust were made. Detailed measurements include major species concentration (H2 O2 N2 and H2O) profiles at various mixer locations made using Raman spectroscopy. Complementary CFD calculations of the flowfield at the experimental conditions also provide additional information on the physics of the problem. These calculations are being conducted at Marshall Space Flight Center to benchmark the FDNS code for RBCC engine operations for such configurations. The primary fluid physics of interests are the mixing and interaction of the rocket plume and secondary flow, subsequent combustion of the fuel rich rocket exhaust with the secondary flow and combustion of the injected afterburner flow. The CFD results are compared to static pressure along the RBCC duct walls, Raman Spectroscopy specie distribution data at several axial locations, net engine thrust and entrained air for the SLS cases. The CFD results compare reasonably well with the experimental results.

Rehabilitation of the Rocket Vehicle Integration Test Stand at Edwards Air Force Base

NASA Technical Reports Server (NTRS)

Jones, Daniel S.; Ray, Ronald J.; Phillips, Paul

2005-01-01

Since initial use in 1958 for the X-15 rocket-powered research airplane, the Rocket Engine Test Facility has proven essential for testing and servicing rocket-powered vehicles at Edwards Air Force Base. For almost two decades, several successful flight-test programs utilized the capability of this facility. The Department of Defense has recently demonstrated a renewed interest in propulsion technology development with the establishment of the National Aerospace Initiative. More recently, the National Aeronautics and Space Administration is undergoing a transformation to realign the organization, focusing on the Vision for Space Exploration. These initiatives provide a clear indication that a very capable ground-test stand at Edwards Air Force Base will be beneficial to support the testing of future access-to-space vehicles. To meet the demand of full integration testing of rocket-powered vehicles, the NASA Dryden Flight Research Center, the Air Force Flight Test Center, and the Air Force Research Laboratory have combined their resources in an effort to restore and upgrade the original X-15 Rocket Engine Test Facility to become the new Rocket Vehicle Integration Test Stand. This report describes the history of the X-15 Rocket Engine Test Facility, discusses the current status of the facility, and summarizes recent efforts to rehabilitate the facility to support potential access-to-space flight-test programs. A summary of the capabilities of the facility is presented and other important issues are discussed.

Reusable Rocket Engine Maintenance Study

NASA Technical Reports Server (NTRS)

Macgregor, C. A.

1982-01-01

Approximately 85,000 liquid rocket engine failure reports, obtained from 30 years of developing and delivering major pump feed engines, were reviewed and screened and reduced to 1771. These were categorized into 16 different failure modes. Failure propagation diagrams were established. The state of the art of engine condition monitoring for in-flight sensors and between flight inspection technology was determined. For the 16 failure modes, the potential measurands and diagnostic requirements were identified, assessed and ranked. Eight areas are identified requiring advanced technology development.