Rocket Thrust Chambers Research Articles

Combustion Stability Analysis of Rocket Engines with Resonators Based on Nyquist Plots

The paper is concerned with the formulation and validation of an efficient, flexible, and reliable method to assess the influence of resonator rings on combustion dynamics in rocket thrust chambers. The chamber is modeled as a network of low-order acoustic elements that describe three-dimensional waves in cylindrical duct segments with nontrivial boundary conditions and mean flow. The main elements responsible for (thermo)acoustic driving and damping, such as time-lagged fluctuating heat release, injector plate, and short exit nozzle are considered. Particular attention is given to the formulation of a semianalytical resonator ring model, which is capable of handling transverse waves, mode coupling, and mode scattering. A Nyquist plot method based on control theory, which is considerably more efficient than classical root finding techniques, is used to identify the eigenmodes and growth rates of the system. After validating the Nyquist method against analytical results, the stabilizing influence of resonator rings is demonstrated through a representative test case. The key mechanisms induced by the resonator ring that lead to stabilization are also discussed. The proposed method is capable to predict linear stability and give insight in the three-dimensional pressure distribution with very moderate computational effort that allows parametric studies.

Multidisciplinary simulation of a regeneratively cooled thrust chamber of liquid rocket engine: Turbulent combustion and nozzle flow

The present study proposed a unified framework to simulate multi-physical processes which are crucial for trade-off design of liquid rocket thrust chambers among propulsive performance, regenerative cooling, and pressure budget. As the first part, a turbulent combustion model based on the flamelet approach was developed to effectively incorporate detailed chemistry of high hydrocarbon fuel, turbulent mixing, enthalpy loss, and pressure variations within the nonadiabatic nozzle flow. In order to correctly capture the convective heat transfer and viscous friction in the turbulent boundary layer at the chamber wall, an advanced low-Reynolds number turbulence model is adopted in an axisymmetric compressible RANS solver, which is interactively coupled with a cooling analysis module for the conjugate heat transfer and hydraulics through the regenerative cooling channels. The present method has been applied to an actual regeneratively cooled thrust chamber and compared with measurement of hot-firing tests in terms of specific impulse, characteristic velocity, and thrust coefficient. Based on the numerical results, the effects of additional fuel cooling injection and wall friction on the propulsive parameters are discussed.

3D fluid–structure interaction analysis of a typical liquid rocket engine cycle based on a novel viscoplastic damage model

SUMMARYIn many space missions, expandable or reusable launch systems are used. In this context, the reliable design of liquid rocket engines (LREs) is a key issue. In the present paper, we present a novel combination of numerical schemes. It is applied to model the extreme physical phenomena a typical LRE undergoes during its loading cycles. The numerical scheme includes a partitioned fluid–structure interaction (FSI) algorithm in combination with a unified viscoplastic damage model. This allows the complex description of the material response under cyclic thermomechanical loading taking place in LREs. In this regard, we focus on the response of the cooling channel wall that is made from copper alloys. For the coupled FSI analysis, the individual domains of the rocket thrust chamber are modeled by a 3D parametrized approach. The well‐established single field solver codes, DLR TAU for the hot gas and ABAQUS FE software for the structural domain, are coupled via the inhouse developed simulation environment ifls. Ifls provides the necessary algorithms for a partitioned coupling approach such as individual code steering, data interpolation, time integration and iteration control. Finally, the results of an FSI analysis of a complete engine cycle are presented. They show the potential of the new numerical scheme for the lifetime prediction of LREs.Copyright © 2013 John Wiley & Sons, Ltd.

Injector Head Transpiration Cooling Coupled with Combustion in Subscale Thrust Chamber

Transpiration cooling coupled with combustion was investigated in an liquid rocket thrust chamber with a transpiration-cooled injector plate. A numerical model was developed using the real gas equation of state. The and combustion process was modeled by the eddy dissipation concept model, which includes the detailed chemical reaction mechanisms in turbulent flows. The permeability and the inertia coefficient of the metal mesh, porous media used for the injector plate were obtained experimentally. The simulation results for the porous plate temperatures and fluxes compare well with hot firing experiments, giving reliable predictions of the combustion, flow, and heat transfer processes in the liquid rocket thrust chamber. The model was also used to investigate the effects of the inlet conditions and the plate material on the transpiration cooling. The results show that locations near the chamber wall and the ignition hole on the plate surface have higher temperatures than other locations. The relatively small thermal conductivity of the porous media is also shown to decrease the plate surface temperature, but increase the thermal stress in the plate.

연료 막냉각을 적용한 액체로켓 연소기의 연소/냉각 성능 간 Trade-off 해석

액체로켓 추력실의 성능 예측 및 초음속 노즐부 형상 설계에 활용 중인 in-house 해석 도구를 이용하여 재생냉각 연소기에 대한 성능/냉각 통합해석을 수행하였으며, 막냉각 유량 및 외곽 분사기열의 혼합비 변화에 따른 연소 성능과 냉각 성능 간 trade-off 경향을 고찰하였다. 향후 막냉각 및 주요 설계 인자의 최적화 도구로 활용될 수 있도록 개발 연소기에 대한 시험 결과와의 비교 등을 통하여 수치해석 도구를 검증/개선해나갈 계획이다. Performance of a liquid rocket thrust chamber with regenerative cooling scheme has been numerically analyzed using in-house CFD code which can predict combustion/cooling performance and provide nozzle design parameters. This paper investigates trade-offs between combustion and cooling performance with varying amount of fuel directly injected into the chamber wall to form cooling films and mixture ratios for the peripheral injectors. Further efforts to verify/improve the simulation methodology including comparison with the firing test results are planned to make it a reliable tool to optimize the film cooling and other major design parameters.

Enhanced Cooling of a Liquid-Fueled Rocket Thrust Chamber by Metal Foams

Enhanced Cooling of a Liquid-Fueled Rocket Thrust Chamber by Metal Foams

Effects of Recess Length on Discharge Coefficients of Swirl Coaxial Injectors

An experimental study has been performed to investigate the effects of the recess length on the discharge coefficients of swirl coaxial injectors. Liquid oxygen and kerosene (Jet A-1) were burned in a range of mixture ratios (1.97–3.31) and chamber pressures (43–80 bar) in small liquid rocket thrust chambers. Each thrust chamber was equipped with 19 liquid–liquid swirl coaxial injectors. While changing the recess length of the injectors, static pressure, temperature, and flow rate data were collected and analyzed to calculate the discharge coefficient. The discharge coefficients obtained from hot-firing tests are compared with those from cold-flow tests. Operating conditions such as oxidizer-to-fuel mixture ratio and chamber pressure have some influence on the discharge coefficients. The recess length of the injectors significantly affects the discharge coefficients in the hot-firing tests, but not in the cold-flow tests. The results show that the longer recess length gradually reduces the discharge coefficients. Especially when the liquid oxygen film collides with the kerosene film inside the fuel nozzle, the discharge coefficients of the outer fuel-side injectors decrease more remarkably than those of the inner oxidizer-side. It is thought that the flame anchored inside the recessed region in the swirl coaxial injector confines propellant flows, and the internal mixing of propellants strongly disturbs the outer fuel flow. Since the discharge coefficient is directly related to the injection pressure drop, the effects of recess length on discharge coefficients must be carefully considered when designing liquid–liquid swirl coaxial injectors.

Enhanced Cooling of a Liquid-Fueled Rocket Thrust Chamber by Metal Foams

DOI: 10.2514/1.B34422 As human beings pursue the deeper space exploration, advanced propulsion systems that deliver super performance and possess increased durability for the next generation spacecraft will be needed, despite current limitations on materials. A promising technology is to fill the current open cooling channels with open-cell metal foams to drastically enhance the cooling of the thrust chamber. After applying the methods developed for the feasibility evaluation, it is found that under equal pressure drop and coolant mass flow rate, the actual amount of increase in the heat transfer coefficient for the foamed channel over the open channel is approximately the same for any given pressure drop, while the percent heat transfer coefficient increase actually decreases with increasing pressuredropfrom99%at841 kPa=mto48%at3090 kPa=m.Itisalsoshownthatunderequalpressuredropand coolant mass flow rate in an actual rocket condition, the heat transfer coefficient of the foamed channel is 49581 W=m 2 K,whichismorethan110%enhancement,comparedwith23464 W=m 2 Kfortheopenchannel.That means that, under the same pressure drop and mass flow rate, the foamed channel has the capability to significantly enhance the cooling for the rocket engine thrust chamber.

Experimental investigations of transpiration cooling applied to C/C material

In today’s rocket thrust chambers extremely high heat loads are apparent. Reducing these loads in combination with a reduction of the thrust chamber’s structural weight can be achieved with transpiration cooling applied to composite carbon/carbon materials (C/C). This article presents the results from transpiration cooling tests in subsonic main-flow at Mach numbers between Mg = 0.3 and Mg = 0.7 while total temperatures of up to Tt,g=523 K were reached. As coolant air was used as a reference, combined with tests using argon and helium. Four different C/C samples were investigated, varying the thickness and porosity of the material. During the tests, detailed information of the surface temperature of the transpiration cooled C/C samples was obtained using infrared thermography. Then, the cooling efficiency of the porous wall at different main-flow conditions was assessed. It was pointed out, that the main-flow total temperature, wall thickness and material charge or porosity do not influence the cooling behavior in the range tested, whereas the main-flow heat load and coolant used show a significant impact. Temperature profiles within the sample were recorded in addition using thermocouple measurements to investigate the thermal behavior and the assumption of thermal equilibrium within the wall in more detail. Based upon these observations, it is shown that models from literature can be used for the prediction of transpiration cooling effects with foreign gas injection for the investigated materials if the non-adiabatic test environments are taken into account.

Recent Experimental Efforts on High-Pressure Supercritical Injection for Liquid Rockets and Their Implications

Pressure and temperature of the liquid rocket thrust chambers into which propellants are injected have been in an ascending trajectory to gain higher specific impulse. It is quite possible then that the thermodynamic condition into which liquid propellants are injected reaches or surpasses the critical point of one or more of the injected fluids. For example, in cryogenic hydrogen/oxygen liquid rocket engines, such as Space Shuttle Main Engine (SSME) or Vulcain (Ariane 5), the injected liquid oxygen finds itself in a supercritical condition. Very little detailed information was available on the behavior of liquid jets under such a harsh environment nearly two decades ago. The author had the opportunity to be intimately involved in the evolutionary understanding of injection processes at the Air Force Research Laboratory (AFRL), spanning sub- to supercritical conditions during this period. The information included here attempts to present a coherent summary of experimental achievements pertinent to liquid rockets, focusing only on the injection of nonreacting cryogenic liquids into a high-pressure environment surpassing the critical point of at least one of the propellants. Moreover, some implications of the results acquired under such an environment are offered in the context of the liquid rocket combustion instability problem.

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.

Numerical Analysis of Deterioration in Heat Transfer to Near-Critical Rocket Propellants

Liquid propellants, which are typically used for regenerative cooling of rocket thrust chambers, can flow in channels at supercritical pressures and in the neighborhood of pseudocritical temperature (near-critical fluid). This could be for instance the case for the envisioned liquid-oxygen/liquid-methane engines with chamber pressures larger than about 50 bar. When the fluid is in such a near-critical condition, deterioration in heat transfer can occur if the heat transfer level is higher than a threshold value. Aiming to improve flow prediction capabilities for the design of such systems, the present study is devoted to numerical simulations of near-critical fluids flowing in uniformly heated straight tubes. After code validation against experimental data of near-critical-hydrogen flow, numerical simulations of near-critical-methane flow in heated tubes are carried out, each characterized by a different wall heat flux. Results are discussed in detail and the near-critical-methane flow condition that exhibits the heat transfer deterioration is identified and emphasized.

Effect of temperature fluctuations on high frequency acoustic coupling

This article considers the influence of fluctuations in temperature on acoustic resonance characteristics of a combustion chamber. This analysis is concerned with acoustic coupling experienced in high frequency instabilities. Experiments carried out on a multiple injector combustor (MIC) operating with liquid oxygen and gaseous methane are used to simulate on a model scale the situation prevailing in a liquid rocket thrust chamber. External acoustic modulations are generated by periodically blocking an auxiliary nozzle with a rotating toothed wheel. By continuously changing the rotation velocity, it is possible to apply a linear frequency sweep and obtain the response of the system to an external excitation. It is found that the chamber quality factor under hot fire conditions is reduced with respect to the one measured under cold flow conditions. This is shown to be correlated with the level and spatial extension of temperature fluctuations inside the chamber. Temperature fluctuations induce variations in the speed of sound which in turn perturb the eigenfrequencies of the system and modify its response. This effect is modeled in a simplified way by considering a second order dynamical system featuring an eigenfrequency which fluctuates around its mean value. The dynamics is analyzed by making use of simulations and by the method of averaging. It is shown that when the eigenfrequency fluctuates, the system behaves as if it had an augmented level of dissipation and that this induces a reduction in the quality factor. It is concluded that this mechanism could have an important impact on the development of oscillations in liquid rocket engines.

이중 와류 동축형 분사기의 연소안정성 평가에 관한 연구

A liquid rocket thrust chamber should have a high confidence in its combustion performance and combustion stability. Expecially, the injector of which function is spraying and mixing propellants plays an important role in getting the confidence. This study was carried out to evaluate combustion stability of a double swirl coaxial injector by using the model similarity method. Besides, in case of a baffle which was used to improve combustion stability, the length and gap effects of the baffle were investigated.

추진제의 비균일 혼합분포를 고려한 액체로켓 추력실의 성능 예측기법 개발

추진제의 비균일 혼합분포를 고려한 액체로켓 추력실의 성능 예측기법 개발

액체로켓 연소기 재생냉각 챔버 구조설계

재생냉각형 액체로켓 연소기의 냉각채널 설계에 대한 탄-소성 구조설계를 수행하였다. 구조해석에 필요한 데이터를 얻기 위하여 여러 가지 온도에서 재생냉각 챔버 제작에 사용하는 구리합금의 단축인장시험을 수행하였다. 재료시험 결과 구리합금은 브레이징 공정 후에 연화되어 강도 값이 매우 저하되며 온도가 증가함에 따라 유동 응력 값이 더 작게 나타났다. 재료시험 데이터를 이용하여 냉각채널의 구조해석을 수행한 결과 채널 내부 냉각유체에 의한 내부압력보다 고온의 연소가스에 의한 열 하중에 의하여 채널의 변형이 심하게 발생함을 확인하였다. 따라서 기계적인 하중을 견딜 수 있는 한도 내에서 냉각채널의 두께를 감소시켜 열 하중을 최소화함으로서 연소기의 무게 감소, 냉각성능 향상, 그리고 구조적인 안정성을 향상시킬 수 있다는 것을 알 수 있었다. Elastic-plastic structural analysis for regenerative cooling chamber of liquid rocket thrust chamber is performed. Uniaxial tension test is also conducted for the copper alloy in order to get material data necessary for the structure analysis. The results of uniaxial tension test reveal that copper alloy become ductile after brazing process and flow stress becomes lower as temperature becomes higher. As a result of structural analysis using the material data, the deformation of cooling channel is more increased by thermal load than by internal pressure of cooling fluid. Therefore, the results of analysis show that structural stability and cooling performance of combustion thrust chamber which is designed to endure mechanical load and minimized a channel thickness are improved by decreased thermal load as possible.

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.

Time- and Length Scales of Combustion in Liquid Rocket Thrust Chambers

This paper summarizes experimental and computational results of research on propellant mixing and combustion in high-pressure LOX/GH2 rocket combustors. Hot-fire tests and numerical experiments are utilized to get more detailed information about the physics of mixing and combustion under high-pressure (transcritical) conditions. Time- and length scales of the reacting shear layer are identified by results from numerical simulation and experiments. The comparison between numerical and experimental results show the same tendencies of the distribution of length scales within the reactive shear layder. It was ascertained that the length scales are increasing downstream of the injector face plate. Qualitative and quantitative data about the flowfield (morphology, velocity, species distribution, etc.) at different boundary conditions are presented.

Thermophysics Characterization of Kerosene Combustion

A one-formula surrogate fuel formulation and its quasi-global combustion kinetics model are developed to support the design of injectors and thrust chambers of kerosene-fueled rocket engines. This surrogate fuel model depicts a fuel blend that properly represents the general physical and chemical properties of kerosene. The accompanying gaseous-phase thermodynamics of the surrogate fuel is anchored with the heat of formation of kerosene and verified by comparing a series of one-dimensional rocket thrust chamber calculations. The quasi-global combustion kinetics model consists of several global steps for parent fuel decomposition, soot formation, and soot oxidation and a detailed wet-CO mechanism to complete the combustion process. The final thermophysics formulations are incorporated with a computational fluid dynamics model for prediction of the combustion efficiency of an unielement, tripropellant combustor and the radiation of a kerosene-fueled thruster plume. The model predictions agreed reasonably well with those of the tests.

REAL GAS EFFECTS SIMULATION IN A MODEL GH2/O2 ROCKET THRUST CHAMBER

A CFD-model for a multicomponent reacting mixture of real gases at supercritical conditions is integrated into the KIVA-II code to analyze high pressure GH2/O2 ignition processes in a model single injector rocket thrust chamber. The model incorporates mixture thermodynamic anomalies accounted for in terms of the Redlich-Kwong equation for O2, while another species are treated as ideal, perfect gases, the k − ε turbulent submodel for compressible media, and equilibrium and finite-rate combustion chemistry. The detailed mechanism of H2/O2 oxidation is used to simulate ignition initiated by "hot spots". The "dense" gas formulation, in contrast to the ideal gas model, predicts mixing enhancement within the cylindrical part of the thrust chamber leading to a decrease in the oxygen jet core length, more stringent requirements for mixture flammability and reduced combustion efficiency. Using equilibrium chemistry, the model does not predict ignition and pressure rise time histories comparable to the results of calculation with finite-rate kinetics.