Takuo Onodera

Effect of Liquid Disintegration on Flow Instability in a Recessed Region of a Shear Coaxial Injector

A liquid jet disintegration phenomenon in the recessed region of a shear coaxial injector was examined to find the relation between the difference of the recess depth and flow instability which may induce combustion instability of liquid rocket engines. An injector with a rectangular cross section and a recess for a central post, which modeled the shear coaxial injector element employed in liquid rocket engines, was constructed. The injector was made of transparent acrylic glass to allow observation of the disintegration phenomenon in the recess. Cold-flow tests with water and nitrogen gas with ambient pressures of 0.2, 0.3 and 0.4 MPa were conducted. Results showed that a condition arose in which the flow in the recess was choked by two-phase flow. The choked flow was accompanied by vibration of the central post which caused a significant change of the disintegration pattern from moderate disintegration to violent disintegration. A similar transition from a fiber-type flow to a super-pulsating disintegration flow reported by Chigier and Reitz was also observed under non-choked conditions for a coaxial injector without a recessed region. The boundary of the transition was found to depend on certain values of ReL/(WeG) 0.5 for each recess depth, including the two-phase choked flow condition. This means that the transition from a fibertype flow to super-pulsating disintegration leads to the transition from a non-choked flow to a choked flow.

Combustion Instability Phenomena Observed During Cryogenic Hydrogen Injection Temperature Ramping Tests for Single Coaxial Injector Elements

For LOX/LH2 shear coaxial injectors, it is well-known that high-frequency combustion instabilities may occur when the injection temperature of hydrogen decreases below a certain value, but the mechanism of the initiation of combustion instability with a coaxial injector is still not clear. In the present study, firing tests were conducted with five types of single shear coaxial injector elements by using LOX and LH2 as propellants to further investigate the mechanism of the initiation of combustion instability during temperature-ramping changes during hydrogen injection. Results showed that unstable combustion was initiated when the hydrogen injection temperature decreased to less than a certain cryogenic temperature. The combustion instabilities observed in the present firing tests are discussed and classified into three different types.

Hot-Firing Test of Methane-Fueled Rocket Engine under High Altitude Condition

Fundamental research activities have been conducted on LOX/Methane rocket engine to achieve higher performance for upper stage system of a middle class launch vehicle. The effects of combustion chamber pressure and mixture ratio on engine performance were demonstrated using a subscale engine test model in the High Altitude Test Stand. Highly expanded nozzle, area ratio of 71 was used in the HATS test to obtain thrust increment generated in the nozzle part. The combustion chamber pressure was changed as a parameter from 1 MPa to 3 MPa, mixture ratio was changed as a parameter from 2.8 to 3.6. Test results shows that specific impulse and thrust performance are increasing with combustion chamber pressure. On the other hand, effects of mixture ratio were small within the test condition.

Status of Experimental Research on High Performance Methane-Fueled Rocket Thrust Chamber

For commercial vehicles, hydrogen and kerosene are widely used as fuel. Methane also has potential for use as a liquid rocket propellant. The authors have designed three types of injectors, which are to be evaluated with subscale hot-firing tests. Flow structure and combustion conditions are observed by single-element combustion visualization hot-firing test. By TDK analysis, nozzle efficiency can be increased by increasing combustion chamber pressure. Verification of analytical results by subscale hot-firing tests is also planned. Data obtained by both test campaigns will contribute to the advance of a high-performance methane engine design.

Studies on Combustion Instability for Liquid Propellant Rocket Engines

To build a framework of a prediction tool for injection-coupled combustion instability with coaxial-type injectors, Hutt and Rocker’s methodology including Crocco’s n-τ model was applied. This linear stability analysis considers a coupling among LOX/fuel flow-path acoustics, chamber responses, and Crocco’s combustion characteristics. To validate the tool, LOX/methane subscale firing tests, which were performed and reported by NASA, were analyzed. The injection-coupled oscillating combustion, which was occurred at 5 kHz, was selected as unstable case. A stable combustion case was also selected for comparison purposes. Injection and combustion characteristics, which amplify the oscillation, were found to be dominated by the LOX-post acoustic characteristics. Chamber responses, which decay the oscillation, were estimated with two approaches: (a) onedimensional acoustic analysis with Natanzon’s methodology with a short-nozzle approximation, and (b) threedimensional acoustic analysis with a commercial software, ACTRAN. The stability was evaluated with AFC (amplitude-frequency characteristics) diagram, in which the injection and combustion characteristics and chamber responses are compared with regard to amplitudes. As a result, significant differences were not seen in the AFC diagrams between the unstable and stable cases (both analyses showed “unstable”). Further investigations for the chamber responses are needed to evaluate the potential instabilities of the system, correctly. In addition, we need to introduce a phase relationship into the tool to understand the underlying physical phenomena.

Experimental and Numerical Study on Characteristics of Fuel Mixers for a Liquid Rocket Engine

Some rocket engines have a fuel mixer upstream of the injector to mix two hydrogen flows of different temperatures. In the mixing process, this fuel mixer may generate large fluctuations of flow properties, which in turn may increase combustion pressure fluctuations. Therefore, fuel mixers must be designed carefully to prevent such large fluctuations. In addition, fuel mixers must have good mixing characteristics and be free of large flow property fluctuations even at off-design points when rocket engines require deep-throttling capability. In this study, we experimentally and numerically investigated the effects of fuel mixer configuration in a rocket engine on the downstream flow properties. In the experiments, we used three different mixer models with different cryogenic hydrogen injection hole configurations (small holes, large holes, and a mixture of both size holes), and conducted experiments using cryogenic hydrogen and gaseous hydrogen under different flow conditions corresponding to engine throttling. The mixers with large injection holes showed better mixing characteristics than the mixer with smaller holes even under conditions of throttling.

REDUCTION OF COMBUSTION PRESSURE FLUCTUATIONS IN ROCKET ENGINE

While many studies to date on combustion instability in rocket engines have focused on high-frequency pressure fluctuations, there have been few studies on low-frequency pressure fluctuations in combustion chambers. Lowfrequency pressure fluctuations are thought to occur near natural frequencies of launch vehicles and can trigger vehicle vibrations. To determine the conditions where low-frequency pressure fluctuations increase, we reanalyzed combustion test data previously obtained at NAL and found that pressure fluctuations tend to increase at low LOX injection velocities. Based on this finding, we constructed two kinds of modified injector elements, one with a smaller flow area cross section and the other with a tapered LOX post. We then performed sub-scale model experiments with these elements. Experimental results showed these modifications to be effective in reducing low-frequency pressure fluctuations in combustion chambers. In addition, using a one-dimensional model for flow in the recess of the injector element, the influence of the tapered LOX post on atomization of the LOX jet was investigated from the point of view of an increase of momentum ratio of propellants. From calculations, it was found that LOX jet expansion in the recess has a major effect on the increase of the momentum ratio.