Masaki Sasaki

Control of Transition between Two Working Modes of a Dual-bell Nozzle by Gas Injection

*† ‡ For a dual-bell nozzle to be employed in a future reusable launch vehicle, active control of working mode transition is essential. To establish a technique for such control, the effectiveness of secondary injection was investigated by conducting cold-flow tests. The secondary injection was introduced from the inflection position of the nozzle contour perpendicular to the thrust axis. As a result, the secondary injection successfully changed the transition condition. Based on the results, a system and a sequence to realize the active control of the transition between two working modes of a dual-bell nozzle are herein discussed.

Combustion and Heat Transfer Modeling in Regeneratively Cooled Thrust Chambers (Optimal Solution Procedures for Heat Flux Estimation of a Full-Scale Thrust Chamber)

Combustion flowfields in GH2/LOX sub-scale calorimeter chambers with multi-injector elements and full-scale thrust chamber are investigated using Reynolds-Averaged NavierStokes simulation, in which the finite rate chemistry with the H2/O2 detailed reaction mechanism is taken into account. The computed wall heat flux distributions are compared to that of the simplified cases to reduce a computational cost. The considered simplifications are a presence of reaction and a number of injector rows. At first, these simplifications are validated in the simulation of sub-scale chambers. The reaction is essential for the prediction of heat flux because it makes change the species distribution in a thermal boundary layer on a thrust chamber wall. A heat flux using a combustion simulation with only outermost injectors shows a good agreement with that with an original configuration near a face plate. On the other hand, it overestimates the heat flux around nozzle and throat parts. It is clarified that this overestimate comes from the shortage of unburned hydrogen near a chamber wall in the simplified method. Next, the simplification of the number of injector rows are applied to the simulation of full-scale thrust chambers. The effectiveness of this simplification for the prediction of wall heat flux is revealed. The optimal solution by using of the simplification is proven to be effective for the prediction of heat flux in a full-scale thrust chamber.

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.

Experimental study on combustion stability characteristics of non-swirl and swirl coaxial injectors

To clarify the difference of combustion stability characteristics between conventional coaxial injectors and swirl-coaxial injectors, an experimental study with liquid oxygen (LOX) /hydrogen and LOX/methane propellants was conducted using thrust chambers with nominal thrusts of 7 and 14 kN at chamber pressures from 2.3 to 10 MPa. Stability rating tests were done with a bomb disturbance technique. The results indicate that the magnitude of induced pressure disturbance with swirl-coaxial injectors is smaller than that with coaxial injectors at the same injection velocity ratio, Vf/Vo. The difference of this stability characteristic is considered to mainly result from the size of the LOX droplets produced. That is, the amplitude of induced maximum pressure disturbance with the same size of bomb appears to be smaller with injectors which produce a fine LOX spray. The increase of fuel density was found to cause a generation of acoustic mode oscillations, a trend independent of the size of LOX droplets. INTRODUCTION Since the occurrence of unstable combustion causes severe damage to rocket engines, a full understanding of this phenomenon is essential to the development of new rocket engines or improvement of existing engines. LOX/hydrogen rocket engines, in general, have employed coaxial injectors, which are known to be simple and characterized by good chamber compatibility and good combustion stability. With these conventional coaxial injectors, LOX is injected at low velocity through a center orifice, and gaseous fuel from an annulus gap around this center LOX orifice is injected with high velocity. With this configuration, atomization and mixing of propellants is achieved by the action of shearing force between gaseous fuel and LOX. Thus, the LOX stream is * Researcher, Rocket Combustion Laboratory t Senior researcher, Rocket Combustion Laboratory J Head, Rocket Combustion Laboratory, Senior Member

Hot-gas-side Heat Transfer Characteristics of a Ribbed Combustor

For expander cycle liquid rocket engines, the enhancement of the heat transfer between combustion gas and regenerative coolant is one of the key issues. The adoption of a ribbed combustor is one possible method of enhancing heat transfer. Hot firing tests of sub-scale combustors with hot gas side wall ribs were conducted to evaluate enhancement of hot gas side wall heat extraction. Based on the firing test results, the influence of combustion pressure and mixture ratio on heat transfer enhancement was evaluated as the basic heat transfer characteristics with ribs and the scale effect was also examined. Nomenclature Pc = chamber pressure MR = mixture ratio ηC* = efficiency of characteristic exhaust velocity φ = rib efficiency

Observation of LOX/Hydrogen Combustion Flame in a Rocket Chamber During Chugging Instability

To obtain concrete information on the mechanism of unstable combustion of liquid oxygen-hydrogen rocket engines, a rectangular rocket chamber with four glass windows was developed. The chamber was designed to simulate a 100-kN-sized cylindrical rocket chamber. Combustion tests were conducted at a chamber pressure of 1.7MPa. Combustion flames and oxygen jets were visualized with a high-speed video at a rate of 4,000 frame/s during low frequency unstable combustion. Oxygen jet images with a backlight, combustion flame and intensity of combustion flame were obtained. Stability analysis based on the double time lag model by Szuch was conducted to assist the understanding of the mechanism of unstable combustion.

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.