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In-stream measurements of combustion during Mach 5 to 7 tests of the Hypersonic Research Engine (HRE)Results of in-stream combustion measurements taken during Mach 5 to 7 true simulation testing of the Hypersonic Research Engine/Aerothermodynamic Integration Model (HRE/AIM) are presented. These results, the instrumentation techniques, and configuration changes to the engine installation that were required to test this model are described. In test runs at facility Mach numbers of 5 to 7, an exhaust instrumentation ring which formed an extension of the engine exhaust nozzle shroud provided diagnostic measurements at 10 circumferential locations in the HRE combustor exit plane. The measurements included static and pitot pressures using conventional conical probes, combustion gas temperatures from cooled-gas pyrometer probes, and species concentration from analysis of combustion gas samples. Results showed considerable circumferential variation, indicating that efficiency losses were due to nonuniform fuel distribution or incomplete mixing. Results using the Mach 7 facility nozzle but with Mach 6 temperature simulation, 1590 to 1670 K, showed indications of incomplete combustion. Nitric oxide measurements at the combustor exit peaked at 2000 ppmv for stoichiometric combustion at Mach 6.
Document ID
19930066107
Document Type
Conference Paper
Authors
Lezberg, Erwin A. (NASA Lewis Research Center Cleveland, OH, United States)
Metzler, Allen J. (NASA Lewis Research Center Cleveland, OH, United States)
Pack, William D. (NASA Lewis Research Center Cleveland, OH, United States)
Date Acquired
August 16, 2013
Publication Date
June 1, 1993
Subject Category
AIRCRAFT PROPULSION AND POWER
Report/Patent Number
AIAA PAPER 93-2324
Meeting Information
AIAA, SAE, ASME, and ASEE, Joint Propulsion Conference and Exhibit(Monterey, CA)
Distribution Limits
Public
Copyright
Other