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Search indexed NASA NTRS and DOE OSTI research on propulsion, heat transfer, battery materials and energy systems. Follow report and document links to the original sources.

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At least 1,531 records · Page 85
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NASA NTRS1973

The Space Shuttle Main Engine

The Space Shuttle Main Engine is the highest performance large rocket engine being developed in the world. This engine will be used to put the Space Shuttle orbiter vehicle into earth orbit and will be the workhorse propulsion system of the 1970s and 1980s. The propulsion requirements demand maximum performance from the liquid oxygen/hydrogen propellants with a minimum hardware weight. The system is being developed at low cost with a required early flight schedule of 1978. The engine system and components are based on the latest technology using advanced materials and high-pressure combustion. Engine weight and performance are discussed. This paper describes engine design, the program and present progress, the technology of advanced components, experimental results, and development areas underway.

Sanchini, D. J.↗
NASA NTRS2015

Combustion Stability Verification for the Thrust Chamber Assembly of J-2X Developmental Engines 10001, 10002, and 10003

The J-2X engine, a liquid oxygen/liquid hydrogen propellant rocket engine available for future use on the upper stage of the Space Launch System vehicle, has completed testing of three developmental engines at NASA Stennis Space Center. Twenty-one tests of engine E10001 were conducted from June 2011 through September 2012, thirteen tests of the engine E10002 were conducted from February 2013 through September 2013, and twelve tests of engine E10003 were conducted from November 2013 to April 2014. Verification of combustion stability of the thrust chamber assembly was conducted by perturbing each of the three developmental engines. The primary mechanism for combustion stability verification was examining the response caused by an artificial perturbation (bomb) in the main combustion chamber, i.e., dynamic combustion stability rating. No dynamic instabilities were observed in the TCA, although a few conditions were not bombed. Additional requirements, included to guard against spontaneous instability or rough combustion, were also investigated. Under certain conditions, discrete responses were observed in the dynamic pressure data. The discrete responses were of low amplitude and posed minimal risk to safe engine operability. Rough combustion analyses showed that all three engines met requirements for broad-banded frequency oscillations. Start and shutdown transient chug oscillations were also examined to assess the overall stability characteristics, with no major issues observed.

Morgan, C. J.↗
NASA NTRSDate not supplied

Cryogenic Extension of NASA Species Polynomials Using Hydrogen and Oxygen at Stoichiometry

NASA has been conducting research in Rotating Detonation Rocket Engines (RDREs) for several years. A recent test at the Marshall Spaceflight Center has successfully shown operation of a RDRE that ran for 251 seconds which had combustion chamber cryogenic inlet conditions [1]. However, an overall pressure gain was not observed in the performance of the engine. Such an indication could mean an optimal chamber design is yet to be discovered. Such a design could significantly reduce parasitic pressure loses and improve overall engine efficiency beyond that of conventional rockets. To investigate different chamber configurations, new CFD capability is needed to accurately capture cryogenic thermophysical and transport property data. Such an effort is currently part of NASA’s Early Career Initiative (ECI) program. This paper focuses on the thermophysical property mixture model approach that makes use of NASA’s polynomial fits of species and another cryogenic model chosen from NIST’s Refprop program. Presented in this paper are results of a 1D denotation CFD simulation of stoichiometric Hydrogen and Oxygen mixture at a cryogenic upstream condition.

Combustion↗
NASA NTRS2001

Simulation of MC-1 Engine on Real-Time Station

The MC-1 rocket engine is a new, 60,000-pound-thrust engine designed to boost small spacecraft carrying payloads weighing up to 500 pounds. The engine was designed in-house at NASA's Marshall Space Flight Center (MSFC) in Huntsville, Alabama, and built by SUMMA Technology, Inc. A vital part of the success of the engine development was the verification of the Propulsion System Controller (PSC) used to control the MC-1 engine during development testing at test facilities in Mississippi and California. The MC-1 engine simulation software was developed on the Applied Dynamics, Inc.'s Real-Time Station (RTS) computer system (ESL) to verify the PSC's hardware and software performance in the Marshall Avionics System Testbed's (MAST) Engine Simulation Lab at the MSFC. The engine model includes the simulation of pressure transducers, thermocouple sensors and valve-positions.

Ly, William↗
NASA NTRS2006

Design and Analysis of a Turbopump for a Conceptual Expander Cycle Upper-Stage Engine

As part of the development of technologies for rocket engines that will power spacecraft to the Moon and Mars, a program was initiated to develop a conceptual upper stage engine with wide flow range capability. The resulting expander cycle engine design employs a radial turbine to allow higher pump speeds and efficiencies. In this paper, the design and analysis of the pump section of the engine are discussed. One-dimensional meanline analyses and three-dimensional unsteady computational fluid dynamics simulations were performed for the pump stage. Configurations with both vaneless and vaned diffusers were investigated. Both the meanline analysis and computational predictions show that the pump will meet the performance objectives. Additional details describing the development of a water flow facility test are also presented.

Dorney, Daniel J.↗
NASA NTRS2014

Combustion Stability Characteristics of the Project Morpheus Liquid Oxygen/Liquid Methane Main Engine

The Project Morpheus liquid oxygen (LOX) / liquid methane rocket engines demonstrated acousticcoupled combustion instabilities during sea‐level ground‐based testing at the NASA Johnson Space Center (JSC) and Stennis Space Center (SSC). High‐amplitude, 1T, 1R, 1T1R (and higher order) modes appear to be triggered by injector conditions. The instability occurred during the Morpheus‐specific engine ignition/start sequence, and did demonstrate the capability to propagate into mainstage. However, the instability was never observed to initiate during mainstage, even at low power levels. The Morpheus main engine is a JSC‐designed ~5,000 lbf‐thrust, 4:1 throttling, pressure‐fed cryogenic engine using an impinging element injector design. Two different engine designs, named HD4 and HD5, and two different builds of the HD4 engine all demonstrated similar instability characteristics. Through the analysis of more than 200 hot fire tests on the Morpheus vehicle and SSC test stand, a relationship between ignition stability and injector/chamber pressure was developed. The instability has the distinct characteristic of initiating at high relative injection pressure drop (dP) at low chamber pressure (Pc); i.e., instabilities initiated at high dP/Pc at low Pc during the start sequence. The high dP/Pc during start results during the injector /chamber chill‐in, and is enhanced by hydraulic flip in the injector orifice elements. Because of the fixed mixture ratio of the existing engine design (the main valves share a common actuator), it is not currently possible to determine if LOX or methane injector dP/Pc were individual contributors (i.e., LOX and methane dP/Pc typically trend in the same direction within a given test). The instability demonstrated initiation characteristic of starting at or shortly after methane injector chillin. Colder methane (e.g., sub‐cooled) at the injector inlet prior to engine start was much more likely to result in an instability. A secondary effect of LOX sub‐cooling was also possibly observed; greater LOX sub‐ cooling improved stability. Some tests demonstrated a low‐amplitude 1L‐1T instability prior to LOX injector chill‐in. The Morpheus main engine also demonstrated chug instabilities during some engine shutdown sequences on the flight vehicle and SSC test stand. The chug instability was also infrequently observed during the startup sequence. The chug instabilities predictably initiated at low dP/Pc at low Pc. The chug instabilities were always self‐limiting; startup chug instabilities terminated during throttle‐up and shutdown chug instabilities decayed by shutdown termination.

Melcher, J. C.↗