Progress in Technology of Main Liquid Rocket Engines of Launch Vehicles in China

Liquid propellant rocket engines for a launch vehicle are an essential aerospace technology, representing the advanced level of hi-tech in a country. In recent years, China’s aerospace industry has made remarkable achievements, and liquid rocket engine technology has also been effectively developed. In this article, the development processes of China’s liquid rocket engines are discussed. Then, the performance features of China’s new generation liquid rocket engines as well as the flight tests of the new-generation launch vehicles are introduced. Finally, the development direction and the most recent progress of the next generation large-thrust liquid rocket engine is presented.

Research on Key Technologies for Reusable Liquid Rocket Engines

Based on current research,the development trend of reusable liquid rocket engines was analyzed.Key technologies and research focuses of the reusable liquid rocket engine have been analyzed and summarized,and then suggestions on the development of future key technologies are proposed.

Numerical and Experimental Characterizations of SiFRP Ablator for the Application to Liquid Rocket Engine Combustors

The ablative material is supposed to be one of good candidates for LRE （liquid rocket engine） combustion chamber to achieve both high reliability and low cost and a numerical analysis for the ablator is considered to be a potentially efficient tool to reduce cost as well. So far, ablators have been successfully applied for many SRM （solid rocket motors）, but the application to LRE is still quite limited in Japan. The authors believe that this is primarily because of the unpredictable nature of the heat load from combustion gases to the combustor wall. Indeed, reliable thermal design of ablative combustion chamber, namely reliable prediction of thermal performance, needs both reliable heat load model and reliable ablator response model. This paper elaborates our research activities and our recent research findings.

Analysis of combustion instability via constant volume combustion in a LOX/RP-1 bipropellant liquid rocket engine

Turbulent two-phase reacting flow in the chamber of LOX/RP-1 bipropellant liquid rocket engine is numerically investigated in this paper. The predicted pressure and mean axial velocity are qualitatively consistent with the experimental measurements. The self-excited pressure oscillations are obtained without any disturbance introduced through the initial and boundary conditions. It is found that amount of abrupt pressure peaks appear frequently and stochastically in the head regions of the chamber, which are the important sources to drive and strengthen combustion instability. Such abrupt pressures are induced by local constant volume combustion, because local combustible gas mixtures with high temperature are formed and burnt out suddenly due to some fuel droplets reaching their critical state in a rich oxygen surrounding. A third Damkhler number is defined as the ratio of the characteristic time of a chemical reaction to the characteristic time of a pressure wave expansion to measure the relative intensity of acoustic propagation and combustion process in thrusters. The analysis of the third Damkhler number distributions in the whole thrust chamber shows that local constant volume combustion happens in the head regions, while constant pressure combustion presents in the downstream regions. It is found that the combustion instability occurs in the head regions within about 30 mm from the thruster head.

Genetic Algorithm to Optimize the Design of Main Combustor and Gas Generator in Liquid Rocket Engines

A genetic algorithm was used to develop optimal design methods for the regenerative cooled combustor and fuel-rich gas generator of a liquid rocket engine. For the combustor design, a chemical equilibrium analysis was applied, and the profile was calculated using Rao's method. One-dimensional heat transfer was assumed along the profile, and cooling channels were designed. For the gas-generator design, non-equilibrium properties were derived from a counterflow analysis, and a vaporization model for the fuel droplet was adopted to calculate residence time. Finally, a genetic algorithm was adopted to optimize the designs. The combustor and gas generator were optimally designed for 30-tonf, 75-tonf, and 150-tonf engines. The optimized combustors demonstrated superior design characteristics when compared with previous non-optimized results. Wall temperatures at the nozzle throat were optimized to satisfy the requirement of 800 K, and specific impulses were maximized. In addition, the target turbine power and a burned-gas temperature of 1000 K were obtained from the optimized gas-generator design.

Conceptual Design for a Kerosene Fuel-rich Gas-generator of a Turbopump-fed Liquid Rocket Engine

A design method for a kerosene fuel-rich gas-generator of a liquid rocket engine using turbopumps to supply propellant was performed at a conceptual level. The gas-generator creates hot gases, enabling the turbine to operate the turbopumps. A chemical non-equilibrium analysis and a droplet vaporization model were used for the estimation of the burnt gas properties and characteristic chamber length. A premixed counter-flow flame analysis was performed for the prediction of the burnt gas properties, namely the temperature, the specific heat ratio and heat capacity, and the chemical reaction time. To predict the vaporization time, the Spalding model, using a single droplet in convective condition, was used. The minimum residence time in the chamber and the characteristic length were calculated by adding the reaction time and the vaporization time. Using the characteristic length, the design methods for the fuel-rich gas-generator were established. Finally, a parametric study was achieved for the effects of the O/F ratio, mass flow rate, chamber pressure, initial droplet temperature, initial droplet diameter and initial droplet velocity.

Verification on Spray Simulation of a Pintle Injector for Liquid Rocket Engine

The pintle injector used for a liquid rocket engine is a newly re-attracted injection system famous for its wide throttle ability with high efficiency. The pintle injector has many variations with complex inner structures due to its moving parts. In order to study the rotating flow near the injector tip, which was observed from the cold flow experiment using water and air, a numerical simulation was adopted and a verification of the numerical model was later conducted. For the verification process, three types of experimental data including velocity distributions of gas flows, spray angles and liquid distribution were all compared using simulated results. The numerical simulation was performed using a commercial simulation program with the Eulerian multiphase model and axisymmetric two dimensional grids. The maximum and minimum velocities of gas were within the acceptable range of agreement, however, the spray angles experienced up to 25% error when the momentum ratios were increased. The spray density distributions were quantitatively measured and had good agreement. As a result of this study, it was concluded that the simulation method was properly constructed to study specific flow characteristics of the pintle injector despite having the limitations of two dimensional and coarse grids.

Development of Preliminary Design Program for Combustor of Regenerative Cooled Liquid Rocket Engine

An integrated program was established to design a combustor for a liquid rocket engine and to analyze regenerative cooling results on a preliminary design level.Properties of burnt gas from a kerosene-LOx mixture in the combustor and rocket performance were calculated from CEA which is the code for the calculation of chemical equilibrium.The heat transfer of regenerative cooling was analyzed by using SUPERTRAPP code for coolant properties and by one-dimensional correlations of the heat transfer coefficient from the combustor liner to the coolant.Profiles of the combustors of F-1 and RS-27A engines were designed from similar input data and the present results were compared to actual data for validation.Finally,the combustors of 30 tonf class,75 tonf class and 150 tonf class were designed from the required thrust,combustion chamber,exit pressure and mixture ratio of propellants.The wall temperature,heat flux and pressure drop were calculated for heat transfer analysis of regenerative cooling using the profiles.

Coupled Lagrangian impingement spray model for doublet impinging injectors under liquid rocket engine operating conditions

To predict the effect of the liquid rocket engine combustion chamber conditions on the impingement spray, the conventional uncoupled spray model for impinging injectors is extended by considering the coupling of the jet impingement process and the ambient gas field. The new coupled model consists of the plain-orifice sub-model, the jet-jet impingement sub-model and the droplet collision sub-model. The parameters of the child droplet are determined with the jet-jet impingement sub-model using correlations about the liquid jet parameters and the chamber conditions.The overall model is benchmarked under various impingement angles, jet momentum and offcenter ratios. Agreement with the published experimental data validates the ability of the model to predict the key spray characteristics, such as the mass flux and mixture ratio distributions in quiescent air. Besides, impinging sprays under changing ambient pressure and non-uniform gas flow are investigated to explore the effect of liquid rocket engine chamber conditions. First, a transient impingement spray during engine start-up phase is simulated with prescribed pressure profile. The minimum average droplet diameter is achieved when the orifices work in cavitation state, and is about 30% smaller than the steady single phase state. Second, the effect of non-uniform gas flow produces off-center impingement and the rotated spray fan by 38°. The proposed model suggests more reasonable impingement spray characteristics than the uncoupled one and can be used as the first step in the complex simulation of coupling impingement spray and combustion in liquid rocket engines.

Study of pressure surge during priming phase of start transient in an initially unprimed pump-fed liquid rocket engine

In this paper, transient phenomenon during start up process of a pump fed liquidrocket engine is investigated through numerical simulation. The engine studied in this workis designed such that engine systems are not wetted with propellant until the engine is com-manded to start. This is achieved by positioning the valves for propellant admission at the inter-face of test stand/flight stage and the engine. To evaluate engine performance during starttransient for such systems, unsteady flow simulation was conducted using Method of Charac-teristics and equations for priming. The same has been reported in this work. The results indi-cated a brief period of abrupt pressure rise at pump upstream after opening of the propellantadmission valves, during the process of priming of engine systems at valve downstream.The peak pressure obtained was significantly higher than the propellant tank pressure as wellas the steady state pump suction pressure. The transitory pressure rise was found to occurdue to flow resistance at impeller inlet caused by formation of a forced vortex for orientingthe flow through impeller blades during off design transient regime. The maximum pressureat pump upstream, as computed from start transient simulation, was used as a design inputfor pump inlet feed lines. The engine was realized and subsequently qualified in a ground test facility. Hot test data obtained for pressure and flow rate during transient regime were found tobe in good agreement with the simulation results.

Experiment, acoustic model for the self-oscillation of coaxial swirl injector and its influence to combustion of liquid rocket engine

During the experiment of gas/liquid coaxial swirl injector conducted with air and water under atmosphere environment, it is observed that the injector may selfoscillate. The self oscillation periodically occurs and vanishes with the increasing velocity of the gas flow.A theoretical model is presented based on the experiment investigation. Simulation of the acoustic process has been performed and conclusions consistent with the experiment can be drawn from the theoretical model, which explains the exPeriment phenomena quite well. At last, the comparison between phenomena of the self oscillation and some experiments of LRE indicates that some instability phenomena in oxygen/hydrogen propellant rocket engine may be the related to self oscillation in coaxial injectors

Regenerative Cooling for Liquid Rocket Engines

Heat transfer in the thrust chamber is of great importance in the design of liquid propellant rocketengines. Regenerative cooling is an advanced method which can ensure not only the proper runningbut also higher performance of a rocket engine. The theoretical model is complicated, it relates to fluiddynamics, heat transfer, combustion, etc... In this papers a regenerative cooling model is presented.Effects such as radiation, heat transfer to environment, variable thermal properties and coking areincluded in the model. This model can be applied to all kinds of liquid propellant rocket engines aswell as similar constructions. The modularized computer code is completed in the work.

Gas film/regenerative composite cooling characteristics of the liquid oxygen/liquid methane (LOX/LCH4) rocket engine

The thermal protection of rocket engines is a crucial aspect of rocket engine design.In this paper,the gas film/regenerative composite cooling of the liquid oxygen/liquid methane(LOX/LCH4)rocket engine thrust chamber was investigated.A gas film/regenerative composite cooling model was developed based on the Grisson gas film cooling efficiency formula and the one-dimensional regenerative cooling model.The accuracy of the model was validated through experiments conducted on a 6 kg/s level gas film/regenerative composite cooling thrust chamber.Additionally,key parameters related to heat transfer performance were calculated.The results demonstrate that the model is sufficiently accurate to be used as a preliminary design tool.The temperature rise error of the coolant,when compared with the experimental results,was found to be less than 10%.Although the pressure drop error is relatively large,the calculated results still provide valuable guidance for heat transfer analysis.In addition,the performance of composite cooling is observed to be superior to regenerative cooling.Increasing the gas film flow rate results in higher cooling efficiency and a lower gas-side wall temperature.Furthermore,the position at which the gas film is introduced greatly impacts the cooling performance.The optimal introduction position for the gas film is determined when the film is introduced from a single row of holes.This optimal introduction position results in a more uniform wall temperature distribution and reduces the peak temperature.Lastly,it is observed that a double row of holes,when compared to a single row of holes,enhances the cooling effect in the superposition area of the gas film and further lowers the gas-side wall temperature.These results provide a basis for the design of gas film/regenerative composite cooling systems.

Numerical study of operational processes in a GOx-kerosene rocket engine with liquid film cooling

Combustion process inside kerosene-GOx rocket combustor with kerosene Alm cooling is studied,and a modeling approach is proposed.The paper suggests to use the Lagrangian particle tracking technique to model fuel film behavior while the continuous fluid is simulated via the Navier-Stokes system of Favre-averaged equations.The approach is validated over the 12 experimental regimes by the criterions of characteristic velocity and pressure,ence on the adiabatic wall temperatures and relatively low impact on the pressure.In general,phenomena,the calculation of operational processes becomes fast and robust yet precise en-the design process.

Dynamic interaction between clustered liquid propellant rocket engines under their asynchronous start-ups

A nonlinear mathematical model of the low-frequency dynamics of the clustered multi-engine rocket propulsion system has been developed and the computations of the engine transient processes during the start-ups of the four-engine propulsion system with a shared feed system have been made applied.Based on propulsion system start-up modeling the influence of the connectivity of engines in a cluster on the starting characteristics of individual engines is shown.In particular,an advanced nonlinear mathematical model of the pump cavitation phenomena is a distinctive feature of the mathematical model.The computation results showed that the asynchronous engines start-ups during rocket lift-off lead to severely nonlinear engine transients and clustered engine thrust misbalance.The influence of the rocket engines asynchronous start-ups on the clustered feed system transients depends on many factors,mainly on from the clustered feed system low-frequency dynamics,the magnitude of the disturbance and the phase difference between disturbances acting on different branches of the feed system.The deep lingering dips in the flow rate and pressure transients are possible due to the nonlinear dynamic interaction of the engines.In case of great pressure dips at the pump inlet(up to the pressure of saturated vapors during significant periods of start-up time)the cavitation breakdowns of the pumps of one or more engines from the cluster are possible.This can disrupt the operation of the entire propulsion system and leads to the failure of the launch vehicle mission.

Investigation of Novel Hydrogen/Oxygen Thruster for Orbital Maneuver in Space Station

Hydrogen and oxygen orbital maneuver thruster, based on gas-dynamic resonance ignition, is a new liquid rocket propulsion technology, and is especially applicable to space station. By means of theoretic thermodynamic calculation of the hydrogen and oxygen thruster, combined with the experimental exploration on the coaxial hydrogen and oxygen resonance ignition, a scheme of the thruster head configuration is designed as the combination of a coaxial hydrogen/oxygen resonance igniter and an oxygen augmentation injector. Through ignition tests on coaxial hydrogen/oxygen resonance igniter characterization, the thruster head ignition tests have been conducted successfully in sequence of resonance ignition and oxygen augmentation combustion. Finally, the thruster ground tests are successfully carried out in forms of single impulse, successive double impulses and 3.0 seconds continuous running, which verify the reliability and feasibility of the thruster. The response time of the thruster starting is restricted within 0.2 second.

Numerical simulation of axial liquid film cooling in rocket combustor

Numerical simulation has been done for liquid film cooling in liquid rocket combustor.Multiple species of axial Navier-Stokes equations have been solved for liquid-film / hot-gas flow field,and k-εequations have been used for compressible turbulent flow.The results of the model agree well with the results of software FLUENT.The results show that :(1) Liquid film can decrease the wall heat flux and temperature effectively,and the cold border area formed by the film covers the whole combustor and nozzle wall.(2) The turbulent viscosity is higher than the physical viscosity,and its biggest value is in the border area of the convergent area in nozzle.The effect of turbulent flow on the whole simulation field can not be ignored.(3) The mass fraction of kerosene at the film inlet is 1,but it decreases along the nozzle wall and achieves its lowest value at the outlet.However,the mass fraction of kerosene near the wall is the biggest at any axial location.

Improving the performance of LOX/kerosene upper stage rocket engines

Improved liquid rocket engine cycles were proposed and analyzed via comparison with existing staged combustion and gas-generator cycles.The key features of the proposed cycles are regenerative cooling of thrust chamber by oxygen and subsequent use of this oxygen for driving one or two oxygen pumps.The fuel pump(s)are driven in a conventional manner,for example,using a fuel-rich gas-generator cycle.Comparison with staged combustion cycle based on oxygen-rich pre-burner showed that one of the proposed semi-expander cycles has a specific impulse only on 0.4%lower while providing much lower oxygen temperature,more efficient tank pressurizing system and built-in roll control.This semi-expander cycle can be considered as a more reliable and cost-effective alternative of staged combustion cycle.Another semi-expander cycle can be considered as an improvement of gas-generator cycle.All proposed semi-expander cycles were developed as a derivative of thrust chamber regenerative cooling performed by oxygen.Analysis of existing oxygen/kerosene engines showed that replacing of kerosene regenerative cooling with oxygen allows a significant increase of achievable specific impulse,via optimization of mixture ratio.It is especially the case for upper stage engines.The increasing of propellants average density can be considered as an additional benefit of mixture ratio optimization.It was demonstrated that oxygen regenerative cooling of thrust chamber is a feasible and the most promising option for oxygen/kerosene engines.Combination of oxygen regenerative cooling and semi-expander cycles potentially allows creating the oxygen/kerosene propulsion systems with minimum specific impulse losses.It is important that such propulsion systems can be fully based on inherited and well-proven technical solutions.A hypothetic upper stage engine with thrust 19.6 kN was chosen as a prospective candidate for theoretical analysis of the proposed semi-expander cycles.The newly-developed software RECS was used for the comparative analysis of engine cycles.

New Power for Boosting China's Aerospace Industry

China's new-generation launch vehicle LM-5 successfully completed its maiden launch in November 2016.Among the new technologies applied in the launch vehicle,four types of liquid rocket engines attracted extensive attention.These engines feature advanced concepts and technologies such as a staged combustion cycle and expander cycle.The engines are the results of hard effort of more than ten years,which is also an epitome of the development history of China's aerospace industry.This paper gives a brief introduction to the technological schemes,main parameters,development process and application of the four types of engines that powered the new-generation launch vehicle.Finally,proposals for new liquid propulsion technology development in the future in China are presented.

空化诱导轮内流动不稳定性的最大似然估

The article illustrates the application of Bayesian estimation to the identification of flow instabilities,with special reference to rotating cavitation,in a three-bladed axial inducer using the unsteady pressure readings of a single transducer mounted on the casing just behind the leading edges of the impeller blades.The typical trapezoidal pressure distribution in the blade channels is parametrized and modulated in time and space for theoretically reproducing the expected pressure generated by known forms of cavitation instabilities(cavitation auto-oscillations and higher-order surge cavitation modes,n-lobed subsynchronous/synchronous/super-synchronous rotating cavitation).The Fourier spectra of the theoretical pressure so obtained in the rotating frame are transformed in the stationary frame,frequency broadened to better approximate the experimental results,and parametrically fitted by maximum likelihood estimation to the measured auto-correlation spectra.Each form of instability generates a characteristic distribution of sidebands in addition to its fundamental frequency.The identification makes use of this information for effective detection and characterization of multiple simultaneous flow instabilities with intensities spanning over about 20 db down to about 4 db signal-to-noise ratios.The same information also allows for effectively bypassing the aliasing limitations of traditional cross-correlation methods in the discrimination of multiple-lobed azimuthal instabilities from the measurements returned by arrays of equally spaced sensors.The method returns both the estimates of the model parameters and their standard deviations,providing the information needed for the assessment of the statistical significance of the results.The proposed approach represents therefore a promising tool for experimental research on flow instabilities in high-performance turbopumps.