Developments in REDES: The Rocket Engine Design Expert System

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

Approximate theoretical performance evaluation for a diverging rocket

A simplified combustion model, which is motivated by available performance studies on the diverging rocket reactor, has been used as basis for an engine performance evaluation. Comparison with conventional rocket configurations shows that an upper performance limit for the diverging reactor is comparable with performance estimates for engines using an adiabatic work cycle. Development of the diverging reactor for engine applications may, however, offer some advantages for very hot, high-energy, propellant systems

Introduction to the problem of rocket-powered aircraft performance

An introduction to the problem of determining the fundamental limitations on the performance possibilities of rocket-powered aircraft is presented. Previous material on the subject is reviewed and given in condensed form along with supplementary analyses. Some of the problems discussed are: 1) limiting velocity of a rocket projectile; 2) limiting velocity of a rocket jet; 3) jet efficiency; 4) nozzle characteristics; 5) maximum attainable altitudes; 6) ranges. Formulas are presented relating the performance of a rocket-powered aircraft to basic weight and nozzle dimensional parameters. The use of these formulas is illustrated by their application to the special case of a nonlifting rocket projectile

Hybrid rocket performance

A hybrid rocket is a system consisting of a solid fuel grain and a gaseous or liquid oxidizer. Figure 1 shows three popular hybrid propulsion cycles that are under current consideration. NASA MSFC has teamed with industry to test two hybrid propulsion systems that will allow scaling to motors of potential interest for Titan and Atlas systems, as well as encompassing the range of interest for SEI lunar ascent stages and National Launch System Cargo Transfer Vehicle (NLS CTV) and NLS deorbit systems. Hybrid systems also offer advantages as moderate-cost, environmentally acceptable propulsion system. The objective of this work was to recommend a performance prediction methodology for hybrid rocket motors. The scope included completion of: a literature review, a general methodology, and a simplified performance model

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

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

Advanced ramjet concepts program

Uniquely advantageous features, on both the performance and weight sides of the ledger, can be achieved through synergistic design integration of airbreathing and rocket technologies in the development of advanced orbital space transport propulsion systems of the combined cycle type. In the context of well understood advanced airbreathing and liquid rocket propulsion principles and practices, this precept of synergism is advanced mainly through six rather specific examples. These range from the detailed component level to the overall vehicle system level as follows: using jet compression; achieving a high area ratio rocket nozzle; ameliorating gas generator cycle rocket system deficiencies; using the in-duct special rocket thrust chamber assembly as the principal scramjet fuel injection operation; using the unstowed, covered fan as a duct closure for effecting high area ratio rocket mode operation; and creating a unique airbreathing rocket system via the onboard, cryogenic hydrogen induced air liquefaction process

Combustor design and analysis using the Rocket Combustor Interactive Design (ROCCID) methodology

The ROCket Combustor Interactive Design (ROCCID) Methodology is a newly developed, interactive computer code for the design and analysis of a liquid propellant rocket combustion chamber. The application of ROCCID to design a liquid rocket combustion chamber is illustrated. Designs for a 50,000 lbf thrust and 1250 psi chamber pressure combustor using liquid oxygen (LOX)RP-1 propellants are developed and evaluated. Tradeoffs between key design parameters affecting combustor performance and stability are examined. Predicted performance and combustion stability margin for these designs are provided as a function of the combustor operating mixture ratio and chamber pressure

A method for comparing the performance of power plants for vertical flight

A new method of power plant selection for vertical flight is proposed. It can be used to determine whether the performance of a rocket design can be improved by substituting for the rocket motor a different power plant such as a ramjet. Calculations indicate that there are advantages in using the ramjet provided the power plant can be made to operate under rapid acceleration and at high altitudes

Injector element characterization methodology

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

Nuclear thermal rocket nozzle testing and evaluation program

Performance characteristics of the Nuclear Thermal Rocket can be enhanced through the use of unconventional nozzles as part of the propulsion system. The Nuclear Thermal Rocket nozzle testing and evaluation program being conducted at the NASA Lewis is outlined and the advantages of a plug nozzle are described. A facility description, experimental designs and schematics are given. Results of pretest performance analyses show that high nozzle performance can be attained despite substantial nozzle length reduction through the use of plug nozzles as compared to a convergent-divergent nozzle. Pretest measurement uncertainty analyses indicate that specific impulse values are expected to be within + or - 1.17 pct