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![Schematic of an Acoustic Resonance Absorber for Liquid Rocket Engines: Taken and edited from [20]](profile/John-Bennewitz/publication/268472794/figure/fig9/AS:570573286592512@1513046680163/Schematic-of-an-Acoustic-Resonance-Absorber-for-Liquid-Rocket-Engines-Taken-and-edited.png)
![Figure 3: Schematic Diagrams for Common Transverse Instability Modes [10]](profile/John-Bennewitz/publication/268472794/figure/fig3/AS:570573282398214@1513046679369/Schematic-Diagrams-for-Common-Transverse-Instability-Modes-10_Q320.jpg)
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... these side lobes act as Helmholtz Resonators which take the acoustic energy out of a specific resonance; the frequencies at which these absorbers resonate depend on the volume of the side lobe and are independent of geometry. Thus, typical resonance absorbers, such as the one seen in Figure 9, have simple geometric shapes [19]. [20] In order to maximize the dampening effect of resonance absorbers, these acoustic side lobes are located at an acoustic pressure anti-node of the instability. ...
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... In order to identify the primary mechanism that generates combustion oscillations, we have derived above the theoretical basis for the occurrence of exhaust velocity bifurcation in the combustion chamber of a rocket engine. We know from Bernoulli's principle that a change in the velocity of a fluid will cause a change in the flow pressure [20]. Therefore, we can deduce that the velocity bifurcation inside the rocket combustion chamber, coupled with repeated cycles of propellant combustion, will set up a flow perturbation which generates a pressure oscillation inside the chamber [21] [22] [23]. ...
... Apart from the main activity, a simple study on the resonant acoustic coupling has been undertaken in order to explain some of the anomalies observed in the measured jet noise spectrum. A hybrid rocket engine's combustion and post-combustion chambers can act as resonators, and pressure fluctuations within the chambers, resulting for example from the combustion process, may excite resonance modes of the chambers [16][17][18]. When the lengths of the two chambers are in a precise ratio, the respective longitudinal modes may be coupled [19], enhancing the acoustic efficiency of the sound transmission outside the nozzle for some frequencies of the spectrum (i.e., the natural frequencies of resonant cavities). ...
... In fact, pressure fluctuations within the chambers, e.g., due to the combustion process, can propagate toward the boundary and then be reflected back toward the flame. These waves combine to produce acoustic pressure and velocity oscillations in the vicinity of the boundary; If these acoustic fluctuations are able to alter the combustion rate with the correct phase, they will be converted into higher amplitude acoustic disturbances [16]. ...
... Normalized relative sound power spectrum as a function of Strouhal number. (a) Experimental data for standard chemical rockets with single nozzle (1.56 to 31,100 kN)[15][16][17][18]20,21]. (b) Data interpolation on the basis of case study input information[11]. ...
A rocket’s turbulent jet radiates intense acoustic waves, which are an acoustic load for structural components like payload, launch structure, and rocket avionics, and impact communities near the launch site. Therefore, a careful characterization of the acoustic field produced by a rocket engine can provide crucial information during the design phase. In particular, this deals with improving the understanding of the acoustics of low-thrust hybrid rocket engines. Since an accurate jet noise detection around the entire launch site is time-consuming and extremely cost-prohibitive, a fast and reliable predictive tool is invaluable. For this purpose, a semi-empirical model was employed, using the exhaust plume property and geometric characteristics of the nozzle as input. Experimental data collected during a firing test campaign, conducted in the framework of HYPROB-NEW project by the Italian Aerospace Research Center, were decisive to discuss the validity of the model also for low-thrust hybrid propulsion and support the goodness of the noise curves and metrics estimated for nearby regions and provide considerations about the implications of engine geometric characteristics on noise emissions.
... If the burning rate is swayed by these acoustic oscillations, the severity of the instabilities inside the reaction slot increases, hence causes enhanced oscillation of flame. As a result of this unstable heat release fluctuations, increased amplitude acoustic disturbances are developed, accelerating the build-up of instability 44) . ...
... In modern annular combustors [37] , fuel injectors are often distributed in the azimuthal direction. In liquid rocket engines, the combustion process involves a series of processes, including injection, atomization, vaporization, mixing and chemical reactions, which elongates the reaction zone and makes the heat release rate nonuniform in the longitudinal and tangential direction [38] . Therefore, the heat source is always distributed in practical systems. ...
Self-excited thermoacoustic oscillations are favorable in thermoacoustic engines, but unwanted in many combustion systems due to the detrimental outcomes, including the structure vibration and overloaded thermal flux to combustor wall. To explore some proper methods to modify thermoacoustic oscillations, understanding the mechanism of thermoacoustic instability is needed. In the present work, the transition to instability in a Rijke-type thermoacoustic system with axially distributed heat source is explored experimentally. A silicon/ceramic tube is used as the acoustic resonator, and the distributed heat source is designed by wounding electric wires over two ceramic rings. The influences of the characteristics of heat source (including heating power, heater length and heater location), mass flow rate, and tube materials on the nonlinear dynamic behaviors and stability boundaries of the Rijke-type thermoacoustic system are systematically evaluated. Large-amplitude limit cycle is observed, and subcritical and supercritical bifurcations are present in the thermoacoustic system. For the system with fixed mass flow rates, the critical heat power for transition to instability is found to be approximately linearly proportional to the heater length, while the strength of pressure oscillations responds nonlinearly. In addition , the system with a thin heater is more prone to undergo subcritical bifurcation. As for the influence of the mass flow rate, it has also been found that increasing the mass flow rate would enhance the nonlinearity of the coupling between the unsteady heat release rate and acoustic waves, which leads to a high possibility of occurrence of subcritical bifurcation. For all the heater length studied, there is an optimal mass flow rate where the critical heat power for the transition to instability is minimal. Last, the results show that the thermoacoustic system with ceramic tube has a narrower stability region at small air flow rate and is more prone to subcritical bifur-cation than the silicon tube.
... The resultant asymmetric mass distribution was found to be favorable to suppress the first tangential instability mode in rocket engines. 33 Given these complex outcomes of misaligned impinging jets, Gadgil and Raghunandan 34 conducted an experimental study on the misalignment effect under Re ranging from 9000 to 30 000 in which the droplet size information was collected at a single site 50 mm downstream the impingement point. The results show that as the misalignment ratioê(ratio between the misalignment distance and the jet diameter) increases, the liquid sheet elongates while the droplet size diminishes. ...
This study numerically investigated the atomization characteristics of misaligned impinging jets, with the misalignment ratio ê ranging between 0 and 0.2, by employing the volume of fluid method with an adaptive mesh refinement algorithm. The results show that the droplet Sauter mean diameter varies non-monotonically with ê and reaches the minimum value, which implies the best atomization performance, at ê=0.1 under operating conditions concerned in the present work. Meanwhile, the moderately misaligned impingement also leads to a more uniform spatial dispersion of the atomized fragments and droplets. These unique spray behaviors can be attributed to the instability and disintegration of the liquid sheet formed upon jet impingement, as evident from the non-monotonic dependence of the breakup length of the liquid sheet on the misalignment ratio ê. Analyses on the velocity fluctuation and vorticity distribution further suggest that the misalignment alters the intrinsic instability mode of the liquid sheet by introducing a lateral stretch effect, which diverts the peak streamwise momentum away from the centerline. The current finding indicates that misalignment tuning could be a promising optimization and control technique in propellant mixing and atomization.
... The first spinning tangential mode is proven, historically, to be most harmful inside the combustion chamber of a liquid rocket engine. The hazard is due to increased heat transfer to the walls of the chamber, caused by combustion products travelling unrestrained around its circumference, as noted by Bennewitz and Frederick (2013). Therefore, it was chosen as the acoustic mode to suppress inside the combustion chamber of the herein developed liquid rocket engine. ...
The basic design of a rocket engine injector and combustion chamber for saturated nitrous oxide and liquid ethanol is presented. At first, an oxidant-fuel mixture is selected based on a thermochemical analysis that explores several existing options and other combinations that have not yet been studied. As a result, nitrous oxide is chosen as an oxidant and ethanol as fuel. Then a simplified methodology is proposed for the design of a pressure-swirl injector responsible for ethanol. Computational fluid dynamics is used to verify the validity of the above-mentioned analysis, using Volume of Fluid (VOF). For the nitrous oxide injector, the flash-boiling phenomenon is investigated, verifying its importance for the ongoing project. The effect is treated analytically using the Dyer model to account for non-equilibrium thermodynamics. Simplified zero-dimensional and one-dimensional combustion models are explored as tools to design the rocket combustion chamber. Furthermore, combustion instability due to acoustic phenomena is studied, with the first spinning tangential mode being computed for the herein developed motor and an ensemble of acoustic cavities being developed to suppress the aforementioned mode. Finally, a diagram of the static test bench which will be used to validate the injectors and the designed engine is also presented.
... Although it cannot yet be accurately predicted, it can be understood: (Zucrow and Hoffman, 1977) presents the underlying theory behind combustion instability, deriving the acoustical modes of oscillation (inside a cylindrical cavity) that may become unstable. On the other hand, (Bennewitz and Frederick, 2013) discuss traditional and novel methods for suppressing these instabilities, specifically passive techniques, like baffles. (Kim, 2010) reveals itself a useful source on the design of small acoustic cavities, which are another type of passive control devices. ...
... The unsteady heat release is responsible for giving rise to acoustic perturbations (in the form of pressure waves), which travel along the combustion chamber until they are reflected at the boundary and come back towards the injector plate, possibly amplifying flame oscillations if with correct phasing, leading to instability growth (Bennewitz and Frederick, 2013). ...
... Its characteristic time is on the order of 10 −3 seconds, close to that of acoustic oscillations. Chemical time (proportional to chemical kinetic rates of reaction) in fuel rich combustion is also around this order of magnitude, thus making combustion itself a possible driver of instability by interacting with acoustic modes (Bennewitz and Frederick, 2013). ...
This report presents the basic design of a rocket engine injector and combustion chamber for saturated nitrous oxide and liquid ethanol, as well as details of the construction and operation of the engine in which the injectors will be used. At first, an oxidant-fuel mixture is selected based on a thermochemical analysis that explores several existing options and other combinations that have not yet been studied. As a result, nitrous oxide is chosen as an oxidant and ethanol as fuel. Then a simplified methodology is proposed for the design of a pressure-swirl injector responsible for ethanol. Computational fluid dynamics is used to verify the validity of the above-mentioned analysis, using Volume of Fluid (VOF). For the nitrous oxide injector, the flash-boiling phenomenon is investigated, verifying its importance for the ongoing project. The effect is treated analytically using the Dyer model to account for non-equilibrium thermodynamics. Simplified zero-dimensional and one-dimensional combustion models are explored as tools to design the rocket combustion chamber. Furthermore, combustion instability due to acoustic phenomena is studied, with the first spinning tangential mode being computed for the herein developed motor and an ensemble of acoustic cavities being developed to suppress the aforementioned mode. Finally, a preliminary diagram of the static test bench which will be used to validate the injectors and the designed engine is also presented.
... On the other hand, the rationale behind attributing the Rayleigh heat addition criterion to explain HFI processes in rocket engines is because there are indeed, oftentimes, clearly linear processes that tend to operate at very similar frequencies [58,[74][75][76] . This is seen in symmetric sinusoidal pressure oscillations at high frequencies [77][78][79][80][81] . The problem arises when the same process is also used to explain the patently non-linear, non-isentropic detonation-like processes that are also categorized under the umbrella of high frequency instabilities [65,82] . ...
... Considering the content provided in the prior sections of this paper, not surprisingly, rocket engines are also susceptible to one or more waves rotating about the combustor circumference; it is called first, second-tangential spinning mode, and so on [59] , as evident in Fig. 32 . Like RDCs, the direction of these waves is random and prefers either clockwise or counterclockwise direction to rotate [78] . In RDCs, smaller fuel orifice area [167] also supports multiple detonation waves inside the combustor. ...
Rotating detonation combustors (RDC) are at the forefront of pressure gain combustion (PGC) research, utilizing one or more azimuthally spinning detonation waves, an intrinsically unsteady process, to effect a stagnation pressure rise across the device. The prospective step-increase in efficiency, simplicity of design without the requirement for mechanical actuations and the ease of assembly make it an especially promising technology that could be integrated into existing propulsion and power generation architectures. This is coupled with the significant complexity of the detonation-based multi-axis flow field and the associated combustion modes and coupling mechanisms. The current paper is an overview of the research done worldwide to address some of the challenges and questions pertaining to the physics of RDC operation. When appropriate, notable parallels are drawn to the phenomena of low and high frequency instabilities in solid and liquid rockets that have been recognized as the most severe hindrance to their operation.
... The rationale behind attributing the Rayleigh heat addition criterion to explain HFI processes in rocket engines is because there are oftentimes, clearly linear processes that tend to operate at very similar frequencies [17,[29][30][31]. This is seen as symmetric sinusoidal pressure oscillations at high frequencies [32][33][34][35][36]. The problem arises when the same process is also used to explain the patently nonlinear, non-isentropic detonation-like processes that are also categorized under the umbrella of high frequency instabilities [37,38]. ...
Rotating detonation combustors (RDC) are at the forefront of pressure gain combustion (PGC) research. The simplicity in design and the ease of assembly makes it a promising technology that could be integrated into existing combustor architectures. This is, however, coupled with the considerable complexities of the detonation-based flow field, and the associated modes and coupling mechanisms. The current paper is an overview of the research done at the University of Cincinnati to address some of the challenges and questions pertaining to the physics of RDC operation. Issues such as combustor geometry, injection schemes and mixing, varied reactants behavior and modes of RDC operation are discussed. The effects of pressurization of the combustor, along with other detonation enhancement strategies are also deliberated upon. When appropriate, parallels are drawn to the phenomena of high frequency combustion instabilities to address the similarities in observations between the two fields.
... U NDERSTANDING the physical mechanisms that contribute to the effectiveness of combustion instability mitigation techniques is of interest when designing combustors. Combustion instability control may be achieved by either providing sufficient damping of the energy that drives the acoustic modes or by breaking the coupling between unsteady heat release and acoustic pressure oscillations [1]. ...
A linear modal analysis is undertaken to investigate the effects of acoustic modulation at the inlet boundary on the longitudinal instability modes of a dump combustor. This study complements an accompanying experimental investigation that demonstrates combustion instability control through single-frequency acoustic modulation at the inlet [Bennewitz, J. W., Frederick, R. A., Jr., Cranford, J. T., Lineberry, D. M., “Combustion Instability Control Through Acoustic Modulation at the Inlet Boundary: Experiments,” Journal of Propulsion and Power, Vol. 31, No. 6 (2015), pp. 1672-1688]. The modal analysis employs acoustically consistent matching conditions instead of the conventional mass, momentum, and energy balances. A specific impedance boundary condition at the inlet is derived through a mass-spring-damper model of a speaker diaphragm that provides the acoustic modulation. The speaker model constants are obtained from an apparatus consisting of a speaker attached to a short hard-wall-terminated duct. At first, the modal analysis is shown to predict a naturally unstable first longitudinal mode in the absence of acoustic modulation, consistent with the spontaneously excited combustion instability mode observed experimentally. Subsequently, a detailed investigation involving variation of the modulation frequency from 0 to 2500 Hz and a mean combustor temperature from 1248 to 1685 K demonstrates the unstable to stable transition of a 2300–2500 Hz first longitudinal mode. The model-predicted mode stability transition is consistent with experimental observations, thereby supporting the premise that inlet acoustic modulation is a means to control high-frequency combustion instabilities. From the modal analysis, it may be deduced that the inlet impedance provides a damping mechanism for instability suppression.
