Technical Innovation of LH2/LOX Rocket Engines in China

This paper provides a detailed introduction to and analysis of the course of China's technological innovation in liquid hydrogen/liquid oxygen(LH2/LOX)rocket engines from a historical point of view.It starts with the investigation of LH2/LOX rocket engines by relevant departments of the Chinese Academy of Sciences in the 1960s and their preliminary achievements.Then,the policy decision concerning LH2/LOX engine development,the project approval of the Long March-3(Chang Zheng-3,CZ-3)rocket,and the process of developing LH2/LOX engines are analyzed in detail,followed by an introduction to and summary of the development situation and technical innovation characteristics of China's LH2/LOX engines as they grew from 4 tons to 8 tons,and finally to 50 tons.Finally,the paper briefly analyzes the innovation experience connected with China's LH2/LOX engines.

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.

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.

Damage localization effects of the regeneratively-cooled thrust chamber wall in LOX/methane rocket engines

To investigate the damage localization effects of the thrust chamber wall caused by combustions in LOX/methane rocket engines, a fluid-structural coupling computational methodology with a multi-channel model is developed to obtain 3-demensioanl thermal and structural responses.Heat and mechanical loads are calculated by a validated finite volume fluid-thermal coupling numerical method considering non-premixed combustion processes of propellants. The methodology is subsequently performed on an LOX/methane thrust chamber under cyclic operation. Results show that the heat loads of the thrust chamber wall are apparently non-uniform in the circumferential direction. There are noticeable disparities between different cooling channels in terms of temperature and strain distributions at the end of the hot run phase, which in turn leads to different temperature ranges, strain ranges, and residual strains during one cycle. With the work cycle proceeding, the circumferential localization effect of the residual strain would be significantly enhanced. A post-processing damage analysis reveals that the low-cycle fatigue damage accumulated in each cycle is almost unchanged, while the quasi static damage accumulated in a considered cycle declines until stabilized after several cycles. The maximum discrepancy of the predicted lives between different cooling channels is about 30%.

Thermal-structural analysis of regeneratively-cooled thrust chamber wall in reusable LOX/Methane rocket engines

To predict the thermal and structural responses of the thrust chamber wall under cyclic work,a 3-D fluid-structural coupling computational methodology is developed.The thermal and mechanical loads are determined by a validated 3-D finite volume fluid-thermal coupling computational method.With the specified loads,the nonlinear thermal-structural finite element analysis is applied to obtaining the 3-D thermal and structural responses.The Chaboche nonlinear kinematic hardening model calibrated by experimental data is adopted to predict the cyclic plastic behavior of the inner wall.The methodology is further applied to the thrust chamber of LOX/Methane rocket engines.The results show that both the maximum temperature at hot run phase and the maximum circumferential residual strain of the inner wall appear at the convergent part of the chamber.Structural analysis for multiple work cycles reveals that the failure of the inner wall may be controlled by the low-cycle fatigue when the Chaboche model parameter c3= 0,and the damage caused by the thermal-mechanical ratcheting of the inner wall cannot be ignored when c3〉 0.The results of sensitivity analysis indicate that mechanical loads have a strong influence on the strains in the inner wall.

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.

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.

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.

A New Simulation Method for 3D Propellant Grain Burn Analysis of Solid Rocket Motor

Describes a new computer program (Regress-3D) to simulate the regression of complex 3D grain cavity and calculate the burning surface area. It has a large region of applicability in solid rocket motor design and has made new improvements compared with other available codes. User can easily and rapidly build his initial grain shapes and then obtain geometric information of his design. Considering with the calclulting results, redesigning can be performed as desire until reaching at the satisfied result. Advantages and disadvantages of this method are also discussed.

Experimental Investigation on Performance of Pulse Detonation Rocket Engine Model

The PDRE test model used in these experiments utilized kerosene as the fuel, oxygen as oxidizer, and nitrogen as purge gas. The solenoid valves were employed to control intermittent supplies of kerosene, oxygen and purge gas. PDRE test model was 50 mm in inner diameter by 1.2 m long. The DDT （deflagration to detonation transition） enhancement device Shchelkin spiral was used in the test model. The effects of detonation frequency on its time-averaged thrust and specific impulse were experimentally investigated. The obtained results showes that the time-averaged thrust of PDRE test model was approximately proportional to the detonation frequency. For the detonation frequency 20 Hz, the time-averaged thrust was around 107 N, and the specific impulse was around 125 s. The nozzle experiments were conducted using PDRE test model with three traditional nozzles. The experimental results obtained demonstrated that all of those nozzles could augment the thrust and specific impulse. Among those three nozzles, the convergent nozzle had the largest increased augmentation, which was approximately 18%, under the specific condition of the experiment.

Ignition characteristics and combustion performances of a LO_2/GCH_4 small thrust rocket engine

A 500 N model engine filled with LO2/GCH4 was designed and manufactured.A series of ignition attempts were performed in it by both head spark plug and body spark plug.Results show that the engine can be ignited but the combustion cannot be sustained when head spark plug applied as the plug tip was set in the gaseous low-velocity zone with thin spray.This is mainly because flame from this zone cannot supply enough ignition energy for the whole chamber.However,reliable ignition and stable combustion can be achieved by body spark plug.As the O/F ratio increases from 2.61 to 3.49,chamber pressure increases from 0.474 to 0.925 MPa and combustion efficiency increases from 57.8%to 95.1%.This is determined by the injector configuration,which cannot produce the sufficiently breakup of the liquid oxygen on the low flow rate case.

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.

An Electro-Hydraulic Actuator for the TVC of a Throttlable Kerolox Rocket Engine

An electro-hydraulic actuator for the thrust vector control(TVC)of a throttlable kerolox rocket engine is introduced in this paper.The creative feature is an integrated hydraulic power drive unit,where a constant speed kerosene motor is used to draw high pressure kerosene from the engine and to drive a constant pressure variable displacement piston pump,acting as the power supply for the actuator.Its operational mechanism,to accommodate the varying pressure from the turbo-pump of a throttling engine,lies in a pressure-reducing flow regulator inserted at the motor inlet.Another key point is that the displacement of the motor is reasonably bigger than the pump so that a sufficiently wide range of pressures can be adapted.Modeling analysis and flight test results were well matched,which show the outstanding performance of this novel type actuator.

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.

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.

Studies on Aerodynamic Behavior and Performance of Aerospike Nozzles

Experimental and numerical studies are carried out on a 6-cell tile-shaped aerospike nozzle, a 6-cell aerospike nozzle with round-to-rectangle primary nozzles and a 1-cell linear aerospike nozzle. Good altitude compensation capacities and high efficiencies are obtained in the tests. The efficiencies of 6-cell tile-shaped aerospike nozzle and 1-cell linear aerospike nozzle at design altitude approach to 100 %, and that of 6-cell aerospike nozzle with round-to-rectangle primary nozzles in the same condition is about 95 % due to the imperfect cell contour and manufacturing defects. Numerical results are in good agreements with test data. The effects of ambient pressure on exhaust and then on base behavior are analyzed, The effects of variation in the amount of base bleed on performance are also examined in the tests.

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.

Adaptive Constrained On-board Guidance Technology forPowered Glide Vehicle

To make full use of expanded maneuverability and increased range,adaptive constrained on-board guidance technology is the key capability for a glide vehicle with a double-pulse rocket engine,especially under the requirements of desired target changing and on-line reconfigurable control and guidance.Based on the rapid footprint analysis,whether the new target is within the current footprint area is firstly judged.If not,the rocket engine ignites by the logic obtained from the analysis of optimal flight range by the method of hp-adaptive Gauss pseudospectral method(hp-GPM).Then,an on-board trajectory generation method based on powered quasi-equilibrium glide condition(QEGC)and linear quadratic regulator(LQR)method is used to guide the vehicle to the new target.The effectiveness of the guidance method consisted of powered on-board trajectory generation,LQR trajectory tracking,footprint calculation,and ignition time determination is indicated by some simulation examples.

An improved constant volume cycle model for performance analysis and shape design of PDRE nozzle

An improved constant volume cycle （CVC） model is developed to analyze the nozzle effects on the thrust and specific impulse of pulse detonation rocket engine （PDRE）. Theoretically, this model shows that the thrust coefficient/specific impulse of PDRE is a function of the nozzle contraction/expansion ratio and the operating frequency. The relationship between the nozzle contraction ratio and the operation frequency is obtained by introducing the duty ratio, by which the key problem in the theoretical design can be solved. Therefore, the performance of PDRE can be accessed to guide the preliminary shape design of nozzle conveniently and quickly. The higher the operating frequency of PDRE is, the smaller the nozzle contraction ratio should be. Besides, the lower the ambient pressure is, the larger the expansion ratio of the nozzle should be. When the ambient pressure is 1.013 × 105 Pa, the optimal expansion ratio will be less than 2.26. When the ambient pressure is reduced to vacuum, the extremum of the optimal thrust coefficient is 2.236 9, and the extremum of the specific impulse is 321.01 s. The results of the improved model are verified by numerical simulation.