Liquid Rocket Thrust Chambers Research Articles

Numerical Investigation on Radiative Heat Loads in Liquid Rocket Thrust Chambers

In liquid rocket engines, wall heat transfer analyses are of primary importance for the thrust chamber design. Numerical investigations usually derive the convective heat flux from computational fluid dynamics simulations, whereas the radiative counterpart is often neglected. The goal of the present paper is a systematic numerical investigation on the role of radiative heat transfer on the overall thermal loads occurring in the thrust chamber of liquid rocket engines. To this purpose, a numerical code for thermal radiation computations is coupled to a computational fluid dynamics solver. The results of the numerical setup are discussed and compared to the available literature’s numerical data. Then, the radiative contribution to the overall heat flux in the thrust chamber of oxygen/hydrogen and oxygen/methane liquid rocket engines is investigated through a wide parametric analysis. The results show that, in typical operating conditions, the radiative heat load is proportionally greater in methane-fueled engines than in hydrogen-fueled engines, and that the radiative contribution to the overall wall heat flux becomes increasingly relevant for large-scale low-chamber-pressure engines. A methodology for a quick preliminary evaluation of the radiative wall heat flux in liquid rocket thrust chambers is finally presented and verified against coupled simulations.

Convective and Radiative Wall Heat Transfer in Liquid Rocket Thrust Chambers

The relative weight of convective and radiative wall heat transfer in liquid rocket engine thrust chambers is estimated by means of dedicated computational fluid dynamics tools. In particular, although the convective wall heat flux is directly evaluated from the solution of the flowfield, an appropriate computational tool, based on the discrete transfer method, is developed to evaluate the radiative wall heat flux. Numerical results are compared against experimental data from the literature, concerning sub-scale thrust chambers fed with either oxygen/methane or oxygen/hydrogen. Then, the effect of varying the chamber pressure or the wall temperature is investigated via a parametric analysis. Results indicate a contribution of radiation up to 15% of the total wall heat flux for oxygen/methane and up to 11% for oxygen/hydrogen propellants.

Chemical Reaction Effects on Heat Loads of CH4/O2 and H2/O2 Rockets

Cooling of liquid rocket thrust chamber walls to allowable solid material temperatures induces near-wall chemical reactions, which are known to have an important role on the heat transfer from the hot gas to the wall. In this study, the contribution of near-wall chemical reactions to heat flux is investigated and quantified by suitable numerical analyses. Numerical results are first compared to literature experimental data of wall heat flux in subscale calorimetric thrust chambers for both oxygen/methane and oxygen/hydrogen propellant combination. Then, a parametric analysis is carried out varying chamber pressure, wall temperature, and propellant combination. This study highlights that oxygen/methane combustion products are more subject to near-wall recombination phenomena. They provide an increase of wall heat flux between 20 and 30% with respect to the frozen flow model evaluation, whereas in the case of oxygen/hydrogen, the wall heat flux increase due to recombination reactions is between 7 and 14%.

Effective Modeling of Conjugate Heat Transfer and Hydraulics for the Regenerative Cooling Design of Kerosene Rocket Engines

The present work is motivated to develop 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. In this paper, an effective modeling of conjugate heat transfer and hydraulics through the regenerative cooling passage has been performed to quantitatively evaluate detailed cooling designs, including spirally twisted channels and bidirectionally branched circuit, as well as to provide the wall heat flux to a compressible reacting flow solver in an interactively coupled manner. The kerosene fuel used as coolant is modeled by a three-component physical surrogate, and the fluid properties required for calculation of a Nusselt number correlation and empirical resistance coefficients are computed over the entire thermodynamic states from compressed liquid to supercritical fluid using the NIST SUPERTRAPP. The present method has been applied to an actual regeneratively cooled thrust chamber and validated against measurement of hot-firing tests in terms of temperature increase and pressure drop of the coolant through the cooling passages. Based on the numerical results, supplementary effects of peripheral fuel cooling injection and thermal barrier coating are addressed.

Effects of injector recess on heat flux in a combustion chamber with cooling channels

An experimental study has been carried out to study the effects of recess length in swirl coaxial injectors on heat transfer in a combustion chamber with cooling channels. Liquid oxygen and kerosene (Jet A-1) were burned in a range of mixture ratios (2.12–3.07) and chamber pressures (57.3–82.3 bar) in small liquid rocket thrust chambers. Each thrust chamber had nineteen liquid–liquid swirl coaxial injectors. While changing the recess length of the injectors and operating conditions, temperatures of the cooling water passing through the cooling channels were collected and analyzed to calculate the heat flux in the combustion chamber. As expected, operating conditions such as oxidizer-to-fuel mixture ratio and chamber pressure had a large influence on the heat flux in the combustion chamber. The recess length of the injectors also significantly affected the heat flux in the combustion chamber, which showed that the longer recess length gradually increased the heat flux. It is thought that the variation of flame structure with respect to the recess length in swirl coaxial injectors was the main reason. Since heat flux is a very important factor in designing cooling channels, the effects of recess length on heat flux in a combustion chamber must be carefully considered.

Scaling theory for liquid propellant rocket thrust chambers

With the advent of the very large liquid propellant rocket, it has become necessary, if possible, to derive a rational scaling theory for combustion chamber design so as to enable relatively simple and economical initial tests to be carried out on small scaled models using scaled parameters of propellant mass flows, pressure etc., and from these to predict operating and design data for the full scale rocket. Owing to the complex and interdependent nature of the aerothermo-chemical processes in the chamber involving evaporation, diffusion and chemical reaction, the similarity criteria must necessarily extend over several, non-dimensional parameters, but it is still possible to evolve relatively simple rules for correlating the design and performance of the model and large scale motors-as shown by Penner-Tsien and Crocco. The paper concludes with a discussion of the accuracy and practical feasibility of such scaling rules.

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.

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.

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.

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.

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.

Optimization of mixture ratio distribution in liquid propellant rocket thrust chamber

An optimization problem is set up for determining the propellant mixture ratio distribution in the liquidpropellant rocket thrust chamber that would yield maximum specific impulse. The problem is solved using the method of Lagrangian Multipliers. The results show that uniform mixture ratio distribution in the combustion chamber will give a maximum value of specific impulse for mixture ratios greater than a critical value. Normal operating mixture ratios for liquid-propellant rocket engines are well above this critical value.

液体ロケット燃焼室内熱負荷に関する研究

液体ロケット燃焼室内熱負荷に関する研究