High-frequency Combustion Instability Research Articles

Analysis of unstable combustion caused by the Gas–Liquid two-phase flow in small hypergolic propellant engines

Hypergolic bipropellant engines are extensively used in spacecraft operations. In these engines, the combustion reaction is initiated by the impinging jets of the oxidizer and fuel in the liquid phase. The most commonly used hypergolic fuels, hydrazine and monomethylhydrazine, as well as the oxidizer, nitrogen tetroxide, are all in liquid state under room temperature and atmospheric pressure conditions. However, as these engines are operated in a space vacuum, a portion of the propellant inevitably evaporates because of low-pressure boiling immediately after engine start-up. The mixing of gas with liquid propellant results in unstable combustion, and minimizing this effect is critical for the engine design. Particularly, preventing high-frequency combustion instability is essential as it can be detrimental to the engine. This paper presents a mechanism that activates high-frequency combustion instability, realized by analyzing the pressure within the combustion chamber during artificially induced unstable combustion. Furthermore, the methodology for establishing operational constraints based on the proposed mechanism is clarified, along with the method for collecting test data. The study findings provide important indicators for the design criteria of bipropellant engines, contributing to the diversification and complexity of space development programs.

Interaction of acoustic pressure and heat release rate fluctuations in a model rocket engine combustor.

We experimentally clarify the interaction of acoustic pressure and heat release rate fluctuations during a transition to high-frequency combustion instability in a model rocket engine combustor. The dynamical state of acoustic pressure fluctuations undergoes a transition from high-dimensional chaotic oscillations to strongly correlated limit cycle oscillations. The coherent structure in the heat release rate field emerges with the initiation of weakly correlated limit cycle oscillations. The effect of the heat release rate on acoustic pressure fluctuations predominates during high-dimensional chaotic oscillations. In contrast, the effect of acoustic pressure on the heat release rate fluctuations markedly increases during the correlated limit cycle oscillations. These are reasonably shown by an ordinal pattern-based analysis involving the concepts of information theory, synchronization, and complex networks.

Driving Mechanisms in Low-Order Modeling of Longitudinal Combustion Instability

Several test cases in the literature have shown that both transverse and longitudinal high-frequency combustion instability can be driven by the injector dynamics. In these cases, pressure oscillations result in fluctuations in propellant mass flow rate, which yields pulsing heat release. This fundamental mechanism is the focus of the present work, with the aim of including this effect in a quasi-1D nonlinear model of Euler equations suited to studies of longitudinal combustion instability. In particular, the injection dynamics is represented through a simplified formulation, which is the core of the proposed response function. The analysis also addresses the influence of combustion efficiency on the main characteristics of the resulting limit cycle (frequency and amplitude). The obtained model is tested comparing the quasi-1D simulations against the experimental data of the continuously variable resonance combustor available in the literature, considering three different geometrical configurations, with different lengths of the oxidizer post. The proposed formulation is capable of reasonably reproducing the unstable behavior, as well as providing a simple model that explains the mechanism that leads to a low average combustion efficiency during unstable operation.

Injection-Coupling Instabilities in the BKD Combustor: Acoustic Analysis of the Isolated Injectors

Injection coupling is a well-known cause of high-frequency combustion instability in hydrogen/liquid oxygen () rocket engines. This type of instability is commonly explained by the two-way coupling between the dynamics of the combustion chamber and the injection system. Recent experimental studies of the BKD combustor, however, suggest that the LOX injector could be self-excited and driving the acoustic mode of the combustion chamber. To assess the feasibility of this mechanism, here, we study both experimentally and theoretically the acoustic stability of the LOX injector isolated from the combustion chamber. The experimental study was performed in a water facility mimicking the conditions of a single LOX injector. The water injector was then modeled using an acoustic network analysis, where the transfer matrix of the LOX injector inlet orifice was computed numerically using a linear approach. The analysis successfully predicts the experimental peak in unsteady pressure, revealing that the LOX injector can be self-excited. The instability was found to be driven by the whistling of the orifice at the inlet of the injector coupled with the second longitudinal acoustic mode of the LOX post tube.

Complex-network analysis of high-frequency combustion instability in a model single-element rocket engine combustor

We study the spatio-temporal dynamics of high-frequency combustion instability in a model single-element rocket combustor using an acoustic energy flux-based spatial network. The acoustic energy source collapses by the formation of small communities with weak connection when the flame edge is attached to the injector rim. In contrast, large communities with strong connection are formed in the shear layer between the oxygen and hydrogen jets when the flame edge is detached from the injector rim, which has a significant impact on driving combustion instability. The switching between the attachment and detachment of the flame edge during combustion instability can be explained by the spectral-clustering-based transition network constructed from the pressure and flow velocity of the hydrogen jet at the injector exit, and the temperature near the injector rim.

Experimental study on high-frequency combustion instability of liquid-propellant rocket engines using off-design combustion model

Combustion tests were conducted to investigate the high-frequency combustion instability of liquid propellant rocket engines using a small experimental model. The model simulates the base space around the propellant injectors of rocket engines. In this study, we presumed that off-design combustion occurred in the base space created by the injected propellants and injector faceplate, and investigated the propagation and repetition of combustion in the base space. The model has two base spaces in which the propagation of combustion was examined. 2-Propanol and gaseous oxygen were used as propellants. 2-Propanol was used as a substitute fuel to study the injection and combustion dynamics of high-density rocket propellants. The test results revealed that (1) the pressure changed up to 20 times higher than the pressure prior to combustion within a millisecond; (2) the flame propagated from one base space to another; (3) combustion was repeated several times, and (4) the pressure outside the base space was dozens of percent to several times greater. The propagated and repeated combustions, as in (2) and (3), can produce a pressure change of a higher frequency than the frequency due to a single combustion. The experimental results indicated that off-design combustion in the base space could be a mechanism for high-frequency combustion instability.

Self-excited instability regimes of a confined turbulent jet flame at elevated pressure

The dynamic response of a confined, premixed turbulent jet flame is investigated at high thermal power densities (∼25 MW/m2/bar) and turbulence levels (Rejet∼ 5 ×105). As equivalence ratio and inlet jet velocity are varied at these conditions, multiple instability modes coexist: a low-amplitude instability (p′/pC≲ 9%) with longitudinal-mode fluctuations (1L and 2L) and two high-amplitude (p′/pC≲ 20%), high-frequency, transverse instability modes. While the axial modes are ubiquitous across the operational envelope, the transverse mode selection is sensitive to the equivalence ratio and reactant jet velocity. A linear stability analysis (LSA) of the confined base flow is performed to explore the flame perturbation in response to the density and temperature gradients, and the shear-layer instabilities in the flow. The high-frequency combustion instabilities are associated with a combined azimuthal hydrodynamic mode of the reactant jet, (1) at the combustion chamber near field and, (2) downstream in the fully developed region of the combustor. An excellent agreement was observed between the convectively unstable modes identified by the temporal LSA and the self-excited combustion instabilities in the experiment.

Effect of baffle injectors on the first-order tangential acoustic mode in a cylindrical combustor

Combustion instability is a very important issue in the development of the propulsion systems used in aerospace. It is very important to associate the high frequency combustion instabilities with the acoustic characteristics of the combustion chamber. In this paper, the effects of various baffle injectors which were installed on the injector faceplate on the first-order tangential acoustic mode were investigated theoretically and experimentally. The effects of the gap between adjacent injectors on the first-order tangential acoustic mode in a cylindrical chamber were considered. The acoustic admittance of the injectors was derived. The results showed that the amplitude and frequency of the first-order tangential acoustic mode increase with the increase in the gap between adjacent injectors, but decrease with the increase in the number and height of the baffles. The baffle injectors have a greater influence on the amplitude and frequency of the first-order tangential acoustic mode than the baffle blades.

Impact of shear-coaxial injector hydrodynamics on high-frequency combustion instabilities in a representative cryogenic rocket engine

The excitation mechanism of a thermoacoustic instability in a 42-element research rocket thrust chamber with representative operating conditions with respect to European cryogenic rocket engines is investigated in detail. From previous research it was known that the chamber 1T mode can be excited by persistent heat release rate oscillations which are modulated by the resonant modes of the liquid oxygen injectors. The excitation source of the longitudinal injector eigenmodes is investigated in this study. Fibre-optical probes measuring the OH* dynamics from the recess volume of two injectors showed additional frequency content which could neither be explained by the chamber acoustics, nor the acoustics of the injection system. Instead, the temporal evolution of these frequencies correlate with the oxidizer flow velocity. In this work we show that the additional flame modulation originates from a hydrodynamic effect in the injection system. Even though the exact process cannot be precisely identified, an effect designated orifice whistling at the injector inlet orifice seems to be a likely candidate. Combining the new results with previous publications about this combustor, it is now possible to explain past and present observations in terms of the hydrodynamic and thermoacoustic conditions which are necessary for the combustion instability to appear. The conditions, which lead to an injection-driven excitation of the 1T mode are matching frequencies of the 2L mode of the injectors and the chamber 1T mode as well as a Strouhal number between 0.2 and 0.4 based on the length and flow velocity of the injector inlet orifice.

Numerical simulation of similarities between rotating detonation and high-frequency combustion instability under two mixing schemes

This paper presents an experimental study on rotating detonations in a hollow combustor with the slit-orifice nozzle. The experimental results reveal that the propagation speed of detonation waves increases with the rise of mass flow rates and is greater than the Chapman–Jouguet detonation speed (VC−J). Furthermore, numerical simulations of rotating detonation in a non-premixed three-dimensional cylindrical combustor have been conducted based on a multispecies reacting code. The influence of two mixing schemes—that is, slit-orifice and coaxial injector—on detonation waves are studied to determine whether the characteristics of detonation waves tend toward high-frequency combustion instability due to changes in the mixing scheme. It is found that the slit-orifice scheme’s detonation speed, pressure, and temperature are significantly higher than those of the coaxial injector scheme. In particular, the detonation speed of the former reaches 124% of the VC−J, while that of the latter is only 80.5% of the theoretical value. The numerical results reveal that the low-speed detonation is caused by the deterioration of the hydrogen (H2)/air mixing conditions. Moreover, the flow-field structures of two mixing schemes were comparable, both containing transverse detonation waves, oblique shocks, contact surfaces, and wedge-shaped reactant regions. Furthermore, the Rayleigh index analysis showed that the unsteady heat release was in phase with the pressure fluctuations, amplifying the pressure. Therefore, it is suggested that high-frequency combustion instability may be a manifestation of rotating detonation waves under poor mixing conditions.

Analysis of Tangential Combustion Instability Modes in a LOX/Kerosene Liquid Rocket Engine Based on OpenFOAM

Self-excited high frequency combustion instability (HFCI) of first-order tangential (1T) mode was observed in a staged-combustion LOX/Kerosene liquid rocket engine numerically. Two different kinds of 1T patterns, standing wave mode and traveling wave mode, were captured in the present work. In the nominal operation condition, the ratio of oxygen-to-fuel (O/F) was 2.5. Propellant was evenly distributed in all injectors and no HFCI occurred. The chamber pressure obtained from the numerical simulation and experiment showed a good agreement, which validated the numerical model. When the mass flow of fuel for two injectors was modified, severe HFCI occurred. The pressure wave node was located at a fixed diameter, showing a 1T standing wave mode. As the O/F was set 4.4 and the propellant distribution was completely uniform, the numerical result yielded a 1T wave node featured a spinning behavior, which was a traveling 1T wave mode. Once the HFCI arose, no matter what standing mode or spinning mode, the pressure and heat release oscillated totally in phase temporally and coupled spatially. The heat release from combustion was fed into the resonant acoustic mode. This was the thermoacoustic coupling process that maintained the HFCI.

A Hybrid Real/Ideal Gas Mixture Computational Framework to Capture Wave Propagation in Liquid Rocket Combustion Chamber Conditions

The present work focuses on the development of new mathematical and numerical tools to deal with wave propagation problems in a realistic liquid rocket chamber environment. A simplified real fluid equation of state is here derived, starting from the literature. An approximate Riemann solver is then specifically derived for the selected conservation laws and primitive variables. Both the new equation of state and the new Riemann solver are embedded into an in-house one-dimensional CFD solver. The verification and validation of the new code against wave propagation problems are then performed, showing good behavior. Although such problems might be of interest for different applications, the present study is specifically oriented to the low order modeling of high-frequency combustion instability in liquid-propellant rocket engines.

High frequency combustion instability control by discharge plasma in a model rocket engine combustor

To suppress the high frequency combustion instability of rocket engines, quasi-direct current (quasi-DC) discharge plasma is considered in combination with special arrangements in a rocket engine combustor. After establishing an internal reacting flow field model of a rocket combustor and a phenomenological plasma model, the influence of quasi-DC discharge plasma on high frequency combustion instability is numerically investigated. The results show that the characteristics of the high frequency combustion instability in the combustor are captured very well through the simulation. Based on the DES turbulence model, a positive feedback process is clearly observed for the self-excited combustion instability. The mean temperature and pressure of the combustor are not sensitive to the occurrence of quasi-DC discharge plasma. The pressure histories of the combustor and their characteristic oscillating frequencies are not affected by the plasma when only one actuator is in operation. However, the oscillating amplitudes of the pressure are reduced within a certain time when several actuators are in operation. The heat release becomes more diffusive and more eddies form around the injector faceplate with an increasing number of plasma filaments. Mainly through the reduction of local heat release, the plasma is capable of affecting the high frequency combustion instability in the combustor. To better control high frequency combustion instability, the optimal arrangement is a double-plasma filaments with symmetrical actuators.

Single pulse combustion test of high-frequency instability of rocket engine

Combustion tests were conducted to investigate the high-frequency combustion instability of liquid-propellant rocket engines. An experimental apparatus was designed to examine the rapid and large pressure increase that occurs near the propellant injector, based on the mechanism of the off-design combustion model. The propellants were 2-propanol and gaseous oxygen. The propanol jet created a semi-enclosed base space near the injection port. Pressure in the semi-enclosed space due to combustion was 1.5–3 times higher than the pressure prior to combustion. The duration of the pressure increase ranged from 0.03 to 2 ms. The pressure increase and its duration agreed with the features of the high-frequency combustion instability of liquid-propellant rocket engines. The model temperatures affected the pressure increase and the duration of pressure change.

Experimental investigation of self-excited combustion instabilities in a small Earth storable bipropellant rocket combustor

A 1000 N scale liquid rocket combustion experiment was designed and conducted to investigate high-frequency combustion instabilities in small Earth storable bipropellant engines. The propellants were injected through unlike impingement injectors. Based on the Hewitt stability correlation, the first tangential combustion instabilities could be spontaneously initialized. As combustion instabilities grew, standing tangential modes gradually shifted to spinning modes. This eventually yielded limit-cycle instabilities with peak-to-peak amplitudes of 40 bar, about four times the static combustor pressure. One single pressure wave characterized by steep rising and falling edges rotated around the chamber at supersonic velocity during spinning mode, which could be defined as a detonation-like wave. However, the pressure oscillated in-phase in the axial direction of the combustor. The frequency of the spinning mode was about 10.9 kHz and was minimally influenced by operation conditions. In a series of test campaigns, the instabilities boundary was identified by the parameters of the Hewitt correlation coefficient (HCC) and the oxidizer/fuel mixing ratio (ROF). When the HCC of oxidizer was less than 0.101 and the HCC of fuel was less than 0.075, self-excited tangential combustion instabilities could be developed easily. The combustion instability with moderate amplitude was not observed.

Comparison of blockage performance of pressure wave in structural baffle injector and fluidic baffle injector

The role of the structural baffle injectors and blades in the liquid rocket engine is to block the transverse pressure waves that are caused by combustion instability. Although the protection of the liquid rocket system from high-frequency combustion instability is essential, there are side effects such as increased weight of the rocket and thermal effect. In this study, a fluidic baffle injector was applied to the simulant spray system, expecting that it would operate with the same performance as a structural baffle injector. The development of the additive manufacturing technology allows for the designing of various geometrical injector arrays such as the combination of a gas-centered swirl injector and shear coaxial injector. This study aimed to compare the blockage performance of a structural baffle injector with that of a fluidic baffle injector using damping capacity and various injector arrays. The damping capacity was high in the structural baffle injector because the pressure wave was completely blocked. However, the amplitude was similar in the shear coaxial injector. This means that the blockage performance of the shear coaxial injector as a fluidic baffle injector is noteworthy.

A compressible two-phase flow framework for Large Eddy Simulations of liquid-propellant rocket engines

Atomization of liquid jets is a key feature of many propulsion systems, such as jet engines, internal combustion engines or liquid-propellant rocket engines (LRE). As it controls the characteristics of the spray, atomization has a great influence on the complex interaction between phenomena such as evaporation, turbulence, acoustics and combustion. In this context, Computational Fluid Dynamics is a promising way to bring better understanding of dynamic phenomena involving atomization, such as e.g. high-frequency combustion instabilities in LRE. However the unsteady simulation of primary atomization in reactive compressible two-phase flows is very challenging, due to the variety of the spatial and temporal scales, as well as to the high density, velocity and temperature gradients which require robust and efficient numerical methods. To address this issue, a numerical strategy is proposed in this paper, which is able to describe the dynamics of the whole chain of mechanisms from the liquid injection to its atomization and combustion. Primary atomization is modeled by a coupling between a homogeneous diffuse interface model and a kinetic-based Eulerian model for the spray. This strategy is successfully applied to the unsteady simulation of an operating point of the Onera’s Mascotte test bench, representative of one coaxial injector of LRE operating under subcritical conditions. The dynamics of the liquid core is retrieved and the flame shape as well as Sauter mean diameters are in good agreement with experimental results. These results demonstrate the ability of the strategy to deal with the harsh conditions of cryogenic combustion, and provide a promising framework for future studies of combustion instabilities in LRE.

Large-eddy simulation of an air-assisted liquid jet under a high-frequency transverse acoustic forcing

The present contribution falls in the scope of high-frequency combustion instabilities occurring in liquid rocket engines. Under subcritical operating conditions, numerical simulations have to render the effects of acoustic waves on the atomisation of liquid jets since these may impact the stability of the engine. Therefore, the present contribution aims at evaluating the ability of a particular numerical strategy adapted to the simulation of two-phase flows to render these interaction mechanisms. The selected strategy is based on the coupling between a diffuse interface method for the simulation of large liquid structures, and a kinetic-based Eulerian model for the description of droplets. The numerical simulation of an air-assisted liquid jet submitted to a transverse acoustic modulation is performed. The flattening of the liquid core under acoustic constraints is retrieved and induces an intensification of its stripping. In addition, thanks to appropriate coupling source terms, the modification of the spray shape as well as periodic oscillations of droplets are retrieved. The numerical strategy is thus proved to be adapted to deal with atomised liquid jets under transverse acoustic modulation and can be used for future numerical studies of high-frequency combustion instabilities.

Effect of syngas composition on high frequency combustion instability in a non-premixed turbulent combustor

An experimental investigation on the effect of syngas composition (H2/CO) on thermodynamic instability of a bluff-body combustor is studied. Three syngas compositions, namely, 25%H2–75%CO, 50%H2–50%CO and 75%H2–25%CO, and pure hydrogen and 75%H2–25%CH4 are acoustically characterized by varying the inlet Reynolds number, Re, in the range of 2200–8100. The combustor is observed to undergo two modes of thermo-acoustic instability across the above Re range for all syngas compositions and pure hydrogen, whereas for the H2CH4 composition, the instability manifests at a single frequency ∼130 Hz. For all the syngas compositions, the pressure oscillations exhibit low frequency (130 Hz) at low Re, and bifurcate to higher frequency (∼500 Hz) at mid-range Re values, while pure hydrogen exhibits bifurcation from the low frequency to the first harmonic (∼260 Hz). The high frequency observed with the syngas compositions is at the third harmonic natural acoustic mode of the combustor and matches thrice the Strouhal frequency of the hydrodynamic oscillations of the shear layer separating from the bluff body. Further investigation into the nature of the dynamic behaviour across the listed fuels is performed using simultaneous time resolved imaging of OH* and CO2* chemiluminescence and two-component particle image velocimetry. The peak in the CO2* chemiluminescent intensity is found to be axially staggered downstream of that of OH*. This is due to sequential oxidation of H2 and CO, the former producing OH as an intermediate consumed by the latter. Spatio-temporal analysis of the time-resolved data reveals that the presence of the flame in the shear layer of the bluff-body plays a significant role in modulating the high frequency acoustic excitation. Further, a spatial Rayleigh map is derived based on both the OH* and CO2* chemiluminescence, which reveals the bluff-body shear layer dynamics to yield the key driving and damping regions in the present study.

The Effects of Non-uniform Distribution of Oxidizer Flow on High-Frequency Combustion Instability

The numerical calculation of three-dimensional unsteady combustion for the combustion chamber of LOX/kerosene high pressure staged combustion rocket engine was carried out. By changing the offset ratio of oxygen mass flow rate in the edge area of the injector face, computational studies were conducted to investigate the effects of non-uniform distribution of oxidizer flow on combustion instability for a liquid-propellant rocket engine. The calculation results show that the offset ratio of oxygen mass flow rate changes the distribution of heat release in the combustion chamber. Within a certain range of offset ratio, the non-uniform distribution degree of oxidizer flow enhances the coupling between the pressure and heat release. As a result, it leads to an increase in the pressure oscillation amplitude in the combustion chamber. However, if the offset ratio is too large, the oxygen-fuel ratio will be too small in some regions, which will reduce coupling between the pressure and heat release and increase the damping of combustion instability.