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.

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.

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.

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.

Russian Nuclear Rocket Engine Design for Mars Exploration

This paper is to promote investigation into the nuclear rocket engine （NRE） propulsion option that is considered as a key technology for manned Mars exploration. Russian NRE developed since the 1950s in the former Soviet Union to a full-scale prototype by the 1990s is viewed as advantageous and the most suitable starting point concept for manned Mars mission application study. The main features of Russian heterogeneous core NRE design are described and the most valuable experimental performance results are summarized. These results have demonstrated the significant specific impulse performance advantage of the NRE over conventional liquid rocket engine （LRE） propulsion technologies. Based on past experience, the recent developments in the field of high-temperature nuclear fuels, and the latest conceptual studies, the developed NRE concept is suggested to be upgraded to the nuclear power and propulsion system （NPPS）, more suitable for future manned Mars missions. Although the NRE still needs development for space application, the problems are solvable with additional effort and funding.

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.

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.

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%.

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.

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.

Index allocation for a reusable LOX/CH4 rocket engine

Reusable rocket engines are the core components of reusable launch vehicles, and have thus become a major focus of aerospace engineering research in recent years. In practice, subsystem design is based on the overall index allocation of an engine;therefore, a multidisciplinary optimization approach is necessary. In this study, design of a reusable methane/liquid oxygen(LOX/CH4)rocket engine with a gas generator cycle was investigated using multidisciplinary optimization. Two parameters were chosen as design variables: pressure and fuel mix ratio of the main combustion chamber. Optimization objectives were specific impulse, structural mass, and life cycle cost of the reusable rocket engine, and constraints were assigned to each discipline according to rocket design requirements. Then, an optimization model was developed, and optimal design parameters were acquired for the LOX/CH4 rocket engine. The proposed method is effective for designing the index allocation of reusable rocket engines and takes into account the multidisciplinary nature of complex systems.

Operation of a Rotary-valved Pulse Detonation Rocket Engine Utilizing Liquid-kerosene and Oxygen

The pulse detonation rocket engine （PDRE） requires periodic supply of oxidizer, fuel and purge gas. A rotary-valve assembly is fabricated to control the periodic supply in this research. Oxygen and liquid aviation kerosene are used as oxidizer and fuel respectively. An ordinary automobile spark plug, with ignition energy as low as 50 mJ, is used to initiate combustion. Steady operation of the PDRE is achieved with operating frequency ranging from 1 Hz to 10 Hz. Experimentally measured pressure is lower than theoretical value by 13% at 1 Hz and 37% at 10 Hz, and there also exists a velocity deficit at different operating frequencies. Both of these two phenomena are believed mainly due to droplet size which depends on atomization and vaporiza-tion of liquid fuel.

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.