Oxygen-Rich Combustion Experiments in a LOX/GH2 Uni-element Rocket

S. A. Rahman, H. M. Ryan, S. Pal, R. J. Santoro

Propulsion Engineering Research Centerand

Department of Mechanical Engineering

The Pennsylvania State University

University Park, PA 16802

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BackmoundCombustion characteristics of a LOX/GH2 swirl coaxial injector element have been examined up to very

high oxidizer to fuel ratios in a research rocket chamber at Penn State University's Cryogenic Combustion Lab. The

single-element tests demonstrate that, for injector element flowrates comparable to those of booster engine injectors,

ignition, stable combustion, and good performance can be achieved with LOX at O/F ratios as high as 170.

Operation of injectors at such high O/F ratios is a highly desirable element of candidate cryogenic

propulsion systems for next-generation Reusable Launch Vehicles (RLV). Oxygen-rich preburners, supplying low

temperature exhaust gases to the turbine drives, have the potential to minimize cost, weight, and operational

complexity of advanced rocket engines. Fundamental data at the single-element level, such as that reported here, is a

component of an industry-wide oxygen-rich combustion technology program for RLV propulsion. Recent progress

is summarized in this presentation.

Research Objectives

Research efforts are directed towards understanding specific technical issues that must be resolved to

minimize the risk and cost associated with developing oxygen-rich rocket preburners. The experiments concentrate

on hot-fire uni-element tests to demonstrate concepts which can be incorporated into hardware design and

development. Two concepts under consideration are direct injection of propellants at high O/F, and stoichiometric

injection followed by downstream injection of LOX to achieve the high O/F. The specific results given here address

the performance, ignition, combustion stability, and wall heat transfer aspects of a direct-injection swirl coaxial

element design operating at high O/F.

Cxu_Jlt_me.m_Experiments with direct-injection at high O/F have been conducted in an optically-accessible uni-element

rocket test chamber of 2 inch square cross-section (1.1 ft. length) with LOX/GH2 propellants. A swirl coaxial

injector element, characterized under both cold-flow and hot-fire, was used to atomize the LOX. LOX flowrates we_held constant in the experiments while O/F ratio was achieved by varying the hydrogen fuel flowrate. A gaseous

hydrogen/oxygen torch was used to ignite the main flow.

A series of experiments has been completed where O/F ratio was varied from 5 to 170, while simultaneous

measurements were made of high frequency pressure oscillations and wall heat transfer. Chamber pressures for this

series were nominally 300 psia, and data was obtained at both upstream and downstream locations within the rocket

chamber. The results show that wall heat transfer is greatly reduced for high O/F combustion. Pressure oscillations

are also at a low level, approximately 1% of chamber pressure, for the entire range of O/F.

Further characterization of the direct-injection high O/F scheme is planned, and will involve non-intrusive

measurement of spray penetration and the spray flame temperature.

Acknowledgment

The oxygen-rich studies are sponsored by NASA Marshall Space Flight Center under NASA Agreement

No. NCC8-46. The swirl coaxial injector design/characterization work is supported by Dr. Mitat Birkan of the AirForce Office of Scientific Research, Air Force Systems Command, under grant number F49620-93-1-0365.

515

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