NASA LogoNTRSNTRS - NASA Technical Reports ServerSearchmore_vertCollectionsAboutNewsHelpLogin Back to ResultsMetallized Gelled Propellants: Oxygen/RP-1/Aluminum Rocket Heat Transfer and Combustion MeasurementsA series of rocket engine heat transfer experiments using metallized gelled liquid propellants was conducted. These experiments used a small 20- to 40-lb/f thrust engine composed of a modular injector, igniter, chamber and nozzle. The fuels used were traditional liquid RP-1 and gelled RP-1 with 0-, 5-, and 55-percentage by weight loadings of aluminum particles. Gaseous oxygen was used as the oxidizer. Three different injectors were used during the testing: one for the baseline O(2)/RP-1 tests and two for the gelled and metallized gelled fuel firings. Heat transfer measurements were made with a rocket engine calorimeter chamber and nozzle with a total of 31 cooling channels. Each chamber used a water flow to carry heat away from the chamber and the attached thermocouples and flow meters allowed heat flux estimates at each of the 31 stations. The rocket engine Cstar efficiency for the RP-1 fuel was in the 65-69 percent range, while the gelled 0 percent by weight RP-1 and the 5-percent by weight RP-1 exhibited a Cstar efficiency range of 60 to 62% and 65 to 67%, respectively. The 55-percent by weight RP-1 fuel delivered a 42-47% Cstar efficiency. Comparisons of the heat flux and temperature profiles of the RP-1 and the metallized gelled RP-1/A1 fuels show that the peak nozzle heat fluxes with the metallized gelled O2/RP-1/A1 propellants are substantially higher than the baseline O2/RP-1: up to double the flux for the 55 percent by weight RP-1/A1 over the RP-1 fuel. Analyses showed that the heat transfer to the wall was significantly different for the RP-1/A1 at 55-percent by weight versus the RP-1 fuel. Also, a gellant and an aluminum combustion delay was inferred in the 0 percent and 5-percent by weight RP-1/A1 cases from the decrease in heat flux in the first part of the chamber. A large decrease in heat flux in the last half of the chamber was caused by fuel deposition in the chamber and nozzle. The engine combustion occurred well downstream of the injector face based on the heat flux estimates from the temperature measurements.Document ID19960045830Acquisition SourceLegacy CDMSDocument TypeConference PaperAuthorsPalaszewski, Bryan*(NASA Lewis Research Center Cleveland, OH United States)*Zakany, James S.*(NYMA, Inc. Brook Park, OH United States)*Date AcquiredSeptember 6, 2013Publication DateAugust 1, 1996Subject CategoryPropellants And FuelsReport/Patent NumberAIAA Paper 96-2622E-10400NASA-TM-107309NAS 1.15:107309Meeting InformationMeeting: Joint Propulsion ConferenceLocation: Lake Buena Vista, FLCountry: United StatesStart Date: July 1, 1996End Date: July 3, 1996Sponsors:American Society for Electrical Engineers,American Inst. of Aeronautics and Astronautics,American Society of Mechanical Engineers,Society of Automotive Engineers, Inc.Accession Number96N32698Funding Number(s)CONTRACT_GRANT: NAS3-27186PROJECT: RTOP 252-50-01Distribution LimitsPublicCopyrightWork of the US Gov. Public Use Permitted.Available DownloadsNameType19960045830.pdfSTIcloud_downloadcontent_copyvisibilityRelated RecordsThere are no records associated with this record.visibility_offNo Preview Available
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