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NASA NTRS · Conference Paper
Dual throat engine design for a SSTO launch vehicle
Attribution
This is the abstract and citation. Full text lives at NASA NTRS — we link out rather than host. All credit to the authors and Legacy CDMS.
Abstract
Verbatim from NASA NTRS. Not paraphrased, not summarized.
A propulsion system analysis of a dual fuel, dual throat engine for launch vehicle application was conducted. Basic dual throat engine characterization data are presented to allow vehicle optimization studies to be conducted. A preliminary baseline engine system was defined. Dual throat engine performance, envelope, and weight parametric data were generated over the parametric range of thrust from 890 to 8896 KN (200K to 2M lb-force), chamber pressure from 6.89 million to 34.5 million N/sq m (1000 to 5000 psia) thrust ratio from 1.2 to 5, and a range of mixture ratios for the two tripropellant combinations: LO2/RP-1 + LH2 and LO2/LCH4 + LH2. The results of the study indicate that the dual fuel dual throat engine is a viable single stage to orbit candidate.
Authors
- Obrien, C. J. Aerojet Liquid Rocket Co.
- Salmon, J. W. Aerojet Liquid Rocket Co.
Citation: Obrien, C. J., Salmon, J. W. (2019). Dual throat engine design for a SSTO launch vehicle. Legacy CDMS. NASA NTRS ID 19800021887. https://ntrs.nasa.gov/citations/19800021887 ↗