Liquid Rocket Engine Combustor Research Articles

Method of Predicting Stability Against Acoustic Vibrations in Liquid Rocket Engine Combustors Based on Combustion Noise

Method of Predicting Stability Against Acoustic Vibrations in Liquid Rocket Engine Combustors Based on Combustion Noise

Scaling of oxygen-methane reacting coaxial jets using x-ray fluorescence to measure mixture fraction

Stoichiometric mixing length Ls of reacting coaxial jet flames is a critical scaling parameter for liquid rocket engine combustors. Previous studies have shown that Ls for shear coaxial flames can be scaled like their nonreacting counterparts using a nondimensional momentum flux ratio J. In addition, stoichiometric mixing lengths of reacting and nonreacting coaxial jets collapse upon a single line by altering J using an effective outer flow gas density. This effective density is calculated from a modified version of the equivalence principle, originally developed by Tacina and Dahm [1, 2] and accounts for the effects of heat release on mixing. However, previous studies also required a second nonphysical scaling constant Sc for the reacting jets, which is not predicted by the equivalence principle [3]. It was originally hypothesized that Sc is attributed to the limitation of hydroxyl (OH) planar laser-induced fluorescence, which only infers Ls. Direct quantitative measurement of conserved scalar fields using conventional optical diagnostics is difficult due to the lack of a tracer that easily fluoresces, survives high temperature oxygen flames, and is not dominated by quenching effects. To measure a conserved scalar field, this work implements x-ray fluorescence of Kr and Ar tracers to obtain quantitative mixture fraction fields. From these mixture fraction fields, stoichiometric mixing lengths for two CH4/O2 flames are calculated and scaled against nonreacting coaxial mixing lengths using the equivalence principle. By directly measuring the stoichiometric mixing length, it is established that the additional constant is a byproduct of the OH measurement technique and the equivalence principle fully captures the scaling. Comparison with high-fidelity simulation of the flame further supports this conclusion. In addition to further strengthening this scaling method, this work represents the first use of x-ray fluorescence to make quantitative conserved scalar measurements in turbulent flames.

Longitudinal combustion instability of a pintle injector for a liquid rocket engine combustor

The longitudinal combustion instability characteristics of a pintle injector for a liquid rocket engine combustor are investigated experimentally. The effects of the propellant total momentum ratio (TMR) on the combustion oscillation characteristics and the characteristic exhaust velocity (C*) efficiency are investigated. The time scales of the combustion oscillations are compared with various phenomena in the combustion chamber in order to identify dominating phenomena of each combustion oscillation mode. The target combustion pressure and the propellant mixture ratio are 0.5 MPa and 1.45, respectively. High-speed optical measurements for the CH* chemiluminescence and backlit spray images are also conducted to investigate the oscillating combustion behaviors of the pintle injector. The amplitude of the pressure oscillation (P’) increases with decreasing TMR. When the TMR is 0.73, the P'is approximately 50 % of the average combustion pressure, whereas the P’ decreases to 20 % of the average combustion pressure when the TMR is increased to 2.35. Frequency characteristics of the combustion oscillations are affected by the TMR. When the TMR is smaller than unity, the pressure oscillations at the chamber acoustic natural frequencies are dominant, and the coupling between the heat release rate and the combustion chamber acoustics could occur. When the TMR is increased, low frequency oscillations whose frequencies are smaller than the fundamental natural frequency of the combustion chamber are observed. The time scale of the observed oscillation is close to the residence time of the combustion gas in the combustion chamber, and the convective process could cause the low frequency oscillations.

Experimental investigation on thermochemical phenomena in SiFRP

This study focuses on understanding and modeling the physical phenomena that occur in degraded zones of silica-phenolic (SiFRP) materials under exposure to high-temperature gasses when applied to a liquid rocket engine (LRE) combustor. Although understanding and modeling these phenomena is considered essential in designing an LRE combustor, few studies on these fields can be found in the available literature. Basically, it is well known that when ablators are heated, a pyrolysis reaction proceeds in them, forming three distinct zones: a charred, a decomposed, and a virgin zone. The obtainable information for the thermal response of SiFRP in ground-firing tests is classified in two categories. The first category involves the equilibrium state characteristics after a long time has elapsed following burnout. This refers to the degraded thickness distribution, which reflects 3D information (the combustor’s inner surface x the thickness direction) regarding the heat load distribution over the entire combustor’s inner surface, owing to the highly insulating nature of SiFRP. The second category involves the transient characteristics of the propagation of the degraded zones in SiFRP, which can be detected using an ultrasonic testing (UT) method. In this paper, the progress of in-depth phenomena of SiFRP and their physical variations were intentionally studied. Our aim was to clarify and specify the quantitative threshold values of the interface points that characterize each degraded zone and the UT reflection point, and then express these values in terms of physical quantities that could appear in a numerical analysis.

Geometric and number effect on damping capacity of Helmholtz resonators in a model chamber

An acoustic cavity was selected as a stabilization device to control high-frequency combustion instabilities in gas turbines or liquid rocket engine combustors, and the acoustic damping capacity of the acoustic cavity was investigated for various geometric configurations under atmospheric non-reacting conditions. The tuning frequency of the acoustic cavity and the acoustic responses of a model chamber with a single acoustic cavity were studied first. Damping capacity was initially quantified through the frequency width of two split modes and the amplitude-damped ratio. The results showed that the cavity with the largest orifice area or the shortest orifice length was the most effective in acoustic damping of the harmful resonant mode. The effect of the number of cavities on acoustic damping capacity was also studied. Damping capacity was improved by increasing the number of cavities. For a better evaluation of acoustic damping capacity, two quantified parameters; the acoustic absorption, meaning the damping efficiency, and acoustic conductance, meaning the acoustic power loss, were introduced. The case was observed that has had insufficient loss of acoustic power in spite of having the highest absorption efficiency. As a result, fine geometric tuning for the acoustic cavity is required for the sufficient passive control. Also, the choice of the number of cavities is important to optimize the damping efficiency and absolute damping loss in consideration of the restriction of the cavity volume.

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 performance of bipropellant liquid rocket engine combustors with fuel-impingement cooling

In order to obtain an accurate combustion analyses which are important in the thruster design of modern advanced liquid rocket engine, flow analysis should be conducted from the injector phase down to the propulsive nozzle throat. Thus, in the present study, flow analysis for the axisymmetric thrust chamber of an OMV(exp 3) installed with a pintle-type ring-shaped injector and a conical convergent nozzle is conducted. Liquid monomethyl hydrazine (MMH) and nitrogen tetroxide (NTO) storable bipropellants are used as fuel and oxidizer sources. An optimum injected fuel and oxidizer droplet-size combination is proposed. Finally, the results obtained are presented. 4 refs.

Hot-gas-side and coolant-side heat transfer in liquid rocket engine combustors

The objectives of this article are to develop a multidisciplinary, computational methodology to predict the hot-gas-side and coolant-side heat transfer in film cooling assisted, regeneratively cooled liquid rocket engine combustors, and to use it in parametric studies to recommend optimized design of the coolant channels for a developmental combustor. An integrated numerical model which incorporates computational fluid dynamics (CFD) for the hot-gas thermal environment, and thermal analysis for the liner and coolant channels, was developed. This integrated CFD/thermal model was validated by comparing predicted heat fluxes with those of hot-firing test and industrial design methods for a 40-k calorimeter thrust chamber and the Space Shuttle Main Engine main combustion chamber. Parametric studies were performed for the advanced main combustion chamber to find a strategy for a proposed coolant channel design.

Bipropellant combustion in a liquid rocket combustion chamber

A comprehensive computer simulation code is developed for the combustion-flow analysis of a bipropellant liquid rocket engine combustor. A bipropellant group combustion model is built in the computer code to facilitate the determination of various combustion modes, including the combustion of oxidizer droplets, fuel droplets and gas mixture. Several physical submodels are also adopted to account for the interaction between bipropellant droplets and the gas-phase flow. An axisymmetric cylindrical liquid rocket engine combustion chamber installed with two annular-ring injectors is selected to demonstrate the capabilities of the present computer simulation code. Two types of injection processes are analyzed. One is characterized by a non-premixed spray process and the other is a premixed spray process. Selected numerical computations show that the premixed spray results in a better utilization of combustion chamber space and gives a higher combustion efficiency in general. But finer-droplet sprays do not necessarily lead to a higher combustion efficiency as traditionally speculated.