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Preliminary calibration of a generic scramjet combustorThe results of a preliminary investigation of the combustion of hydrogen fuel at hypersonic flow conditions are provided. The tests were performed in a generic, constant-area combustor model with test gas supplied by a free-piston-driven reflected-shock tunnel. Static pressure measurements along the combustor wall indicated that burning did occur for combustor inlet conditions of P(static) approximately equal to 19kPa, T(static) approximately equal to 1080 K, and U approximately equal to 3630 m/s with a fuel equivalence ratio approximately equal to 0.9. These inlet conditions were obtained by operating the tunnel with stagnation enthalpy approximately equal to 8.1 MJ/kg, stagnation pressure approximately equal to 52 MPa, and a contoured nozzle with a nominal exit Mach number of 5.5.
Document ID
19910011826
Acquisition Source
Legacy CDMS
Document Type
Contractor Report (CR)
Authors
Jacobs, P. A.
(NASA Langley Research Center Hampton, VA., United States)
Morgan, R. G.
(Queensland Univ. Saint Lucia (Australia)., United States)
Rogers, R. C.
(NASA Langley Research Center Hampton, VA., United States)
Wendt, M.
(Queensland Univ. Saint Lucia (Australia)., United States)
Brescianini, C.
(Queensland Univ. Saint Lucia (Australia)., United States)
Paull, A.
(Queensland Univ. Saint Lucia (Australia)., United States)
Kelly, G.
(Queensland Univ. Saint Lucia, Australia)
Date Acquired
September 6, 2013
Publication Date
March 1, 1991
Subject Category
Aircraft Propulsion And Power
Report/Patent Number
ICASE-16
NASA-CR-187539
AD-A234873
NAS 1.26:187539
Report Number: ICASE-16
Report Number: NASA-CR-187539
Report Number: AD-A234873
Report Number: NAS 1.26:187539
Accession Number
91N21139
Funding Number(s)
CONTRACT_GRANT: NAS1-18605
PROJECT: RTOP 505-90-52-01
CONTRACT_GRANT: NAGW-674
Distribution Limits
Public
Copyright
Work of the US Gov. Public Use Permitted.
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