nghyeon Seo

Effects of Injector Recess and Chamber Pressure on Combustion Characteristics of Liquid–Liquid Swirl Coaxial Injectors

Combustion characteristics such as combustion performance and combustion stability have been studied experimentally using a small liquid rocket thrust chamber with 19 liquid–liquid swirl coaxial injectors. Data were obtained from static pressure, temperature, and dynamic pressure sensors installed in propellant manifolds and the combustion chamber. While changing the recess length of the injector, characteristic velocity and pressure fluctuation data were collected and analyzed. In addition, chamber pressure was varied between 42 and 54 bar, which covers the sub- and supercritical pressures of oxygen. The results show that the longer recess length generally promotes combustion performance and the spray interaction between injectors in the multielement combustor increases the characteristic velocity. When the chamber pressure is above the critical pressure of oxygen, the recess length scarcely affects the pressure fluctuation. However, when the chamber pressure is below the critical pressure, the shift from external mixing to internal mixing of oxidizer and fuel sheets by the variation of recess length significantly degrades combustion stability and induces strong low-frequency instability. Accordingly, the effects of both recess length and operating chamber pressure must be taken into consideration when designing liquid–liquid swirl coaxial injectors.

Combustion Dynamics and Stability of a Fuel-Rich Gas Generator

The dynamic characteristics of fuel-rich combustion have been studied using an experimental combustor simulating a gas generator for a liquid rocket engine. The combustor burns liquid oxygen and fuel (Jet A-1) at a mixture ratio of about 0.32 and a chamber pressure ranging from 4.10 to 7.24 MPa, which covers subcritical to supercritical pressures of oxygen. For the investigation of combustion dynamics, pressure fluctuation measurements using piezoelectric sensors have been a major probe throughout the present study. Two different types of injector heads equipped with biliquid swirl coaxial injectors and either a short nozzle or a turbine manifold nozzle have been used in the study. Fuel-rich combustion of both injector heads with the short nozzle revealed pressure pulsation at frequencies of about 128 Hz, which attenuates along with an increase of a chamber pressure. Combustion tests with the turbine manifold nozzle conducted at chamber pressures lower than the oxygen critical pressure showed combustion instabilities at a frequency of 330 Hz, which has been identified as a longitudinal resonant mode by a linear acoustic analysis. The combustion instabilities seem to be induced by inherent pressure fluctuations from the biliquid swirl coaxial injector when the chamber pressure is below the oxygen critical pressure.

Fuel-Rich Combustion Characteristics of Biswirl Coaxial Injectors

An experimental study was performed to investigate the combustion characteristics of liquid–liquid swirl coaxial injectors in fuel-rich conditions. Liquid oxygen and kerosene (Jet A-1) were burned in a range of mixture ratios (0.29–0.41) and chamber pressures (46–65 bar) in a gas generator for a liquid rocket engine. An injector head was connected to a water-cooled chamber and a short nozzle with or without an extension pipe between the chamber and the nozzle. The extension pipe acoustically simulated a turbine inlet manifold. The injector head had 37 identical swirl coaxial injectors. It is found that the characteristic velocity and combustion gas temperature are seldom influencedby the extension pipe, but are only functions of themixture ratio. The dynamic pressure data show that the combustion instability in the fuel-rich gas generator equipped with biswirl coaxial injectors can be significantly affected by themixture ratio and also by the extension pipe, which influences the resonant frequency in the chamber.

Experimental Verification for Acoustic Damping Enhancement by Gaps in Injector-Formed Baffles

The effect of baffle gaps on the damping enhancement of a liquid rocket engine combustor has been elucidated through a series of tests, which include cold acoustic tests under both atmospheric and simulated viscous conditions and simulated combustion tests. The injector-formed baffles, which consist of an array of protruded coaxial injectors, were found to have a much greater acoustic damping effect than conventional planar baffles. For several axial baffle lengths, an optimal acoustic damping capacitance has been achieved at a 0.1 ∼ 0.2 mm baffle gap. The reason there exists an optimal baffle gap is thought to be mainly due to the viscous dissipation at the surface of the injector-formed baffles. Consequently, the axial baffle length can be reduced by taking advantage of the optimal baffle gap, providing a possible solution to the thermal cooling problem persistent with the baffle. Moreover, these optimum characteristics can provide some guidelines for manufacturing and assembling of the baffled injectors in rocket combustors.

Combustion Stability Assessment of Double Swirl Coaxial Injectors Using Simulant Propellants

*This study is related to conduct model combustion tests applying various double swirl coaxial injectors to identify their combustion stability characteristics. Gaseous oxygen and mixture gas of methane and propane have been used as simulant propellants. A model combustion chamber was designed for its first tangential frequency to be corresponded to that of a full-scale thrust chamber. They were manufactured four different kinds of injector heads with five elements to examine their stability characteristics. The main idea of the experiment is that a propellant mixing mechanism is considered as a dominant factor significantly affecting combustion stability in the full-scale thrust chamber. Self-excited dynamic pressure measurements in the model combustion chamber show different combustion stability characteristics with respect to a recess length of an oxidizer post. The test result shows that the coupling between combustion phenomena and the 1T frequency in the model combustion chamber becomes strengthened according to the increase of a recess ratio.

The Nonlinear Combustion Instability Prediction of Solid Rocket Motors

The prediction of combustion instability is important to avoid an obvious threat to the structural safety and the motor performance because it affects the apparent response function of the propellant, the burning rate, and a mean flow Mach number at the local surface. The combustion instability occurs in case acoustic waves were coupled with the combustion/flow dynamic frequency. In this paper, an acoustic instability model is derived from the nonlinear wave equation for analysing acoustic dynamics in solid rocket motors. The chamber pressure and burning rate effects on combustion instability have been investigated.

Flow Control Characteristics of Cavitating Venturi in a Liquid Rocket Engine Test Facility

The flow rate control of a cavitating venturi has been investigated with downstream pressure variation. A set of cavitating venturies for a liquid rocket engine thrust chamber firing test facility have been designed and manufactured. The flow characteristics of the cavitating venturies have been analyzed by experimental and computational methods. Results showed that constant mass flow rate condition was established by the cavitation inside the venturi. However, upstream pressure less than the actual design pressure of the cavitating venturi could not supply a constant flow rate.

Investigation of self-excited combustion instabilities in two different combustion systems

The objective of this paper is to characterize dynamic pressure traces measured at self-excited combustion instabilities occurring in two combustion systems of different hardware. One system is a model lean premixed gas turbine combustor and the other a fullscale bipropellant liquid rocket thrust chamber. It is commonly observed in both systems that low frequency waves at around 300Hz are first excited at the onset of combustion instabilities and after a short duration, the instability mode becomes coupled to the resonant acoustic modes of the combustion chamber, the first longitudinal mode for the lean premixed combustor and the first tangential mode for the rocket thrust chamber. Low frequency waves seem to get excited at first since flame shows the higher heat release response on the lower frequency perturbations with the smaller phase differences between heat release and pressure fluctuations. Nonlinear time series analysis of pressure traces reveals that even stable combustion might have chaotic behavior with the positive maximum Lyapunov exponent. Also, pressure fluctuations under combustion instabilities reach a limit cycle or quasi-periodic oscillations at the very similar run conditions, which manifest that a self-excited high frequency instability has strong nonlinear characteristics.

Characteristics of Dynamic Pressures in Liquid Rocket Thrust Chambers

The objective of this paper is to characterize dynamic pressure traces measured at various experimental conditions of liquid rocket thrust chambers. Stability rating test and self-excited combustion instability results for subscale and full-scale thrust chambers show various aspects of dynamic behavior of the pressure field depending on combustion system hardware and operating conditions. External perturbations facilitated by a shock wave generated from an explosion of solid material are able to alter the stable combusting flow to an unstable one when it has lack of dynamic stability margin. The direct comparison of pressure traces for self-excited and perturbed high frequency instabilities suggests that coupling mechanisms behind these two cases are totally different from each other. Naturally occurring combustion instability seems to be first coupled with a low frequency wave that affects flame more vigorously with the smaller phase difference between heat release rate and acoustics than a high frequency wave. However, for the artificial perturbation, the combusting flow triggered by a shock wave becomes located at physical conditions irrelevant to those before the perturbation.

Development of 30-Ton f LOx/Kerosene Rocket Engine Combustion Devices(II) - Gas Generator

The development process of a gas generator for a 30-tonf pump-fed space liquid rocket engine is described. Starting from the development of an injector, followed by subscale and full-scale test specimens, the development of LOx/kerosene fuel-rich gas generator has been concluded successfully. Various analytical methods have been utilized in the course of design and the performance requirements have been verified experimentally through ignition tests, combustion performance and stability assessment tests and duration tests. The gas generator has proven its workability and stability within a defined operation window of varying chamber pressure and mixture ratio and demonstrated compliance to the performance and life time requirements.