Liu Peijin

Experimental and Numerical Investigation of Active Cooling Ceramic Matrix Composite for Ramjet Propulsion System

Challenges of hypersonic airbreathing propulsion system are thermal protection, and, moreover, flow rate matching between fuel and coolant. Compare to conventional material, ceramic matrix composite structure was evaluated for the development of lightweight, active cooling composite panels for ramjet propulsion applications. Subscale active cooling and passive cooling refractory composite panel were designed, and experimented with an experimental facility that supplied lower heat transfer rate and an improved gas generator, which supplied hot gas of high-temperature and high heat transfer rate. The active cooling composite panel was tested with maximum surface temperature over 2000K for a cumulative time of 30s and with maximum surface temperature sustained over 3000K for 9.3s durations. A 1D transient heat transfer numerical simulation was carried out, and multi-layers structure with different materials was modeled. The result showed that the active cooling composite structure was available for high-temperature condition in hypersonic airbreathing propulsion.

Algorithm Study on Burning Surface Calculation of Solid Rocket Motor with Complicated Grain Based on Level Set Methods

We apply level set methods to create a new burning surface calculation algorithm, viewing burning surface as interface of different materials. New algorithm attains burning surface variation regulation of arbitrary configuration grain by tracing grain burning surface recession. In the full paper, we explain our new algorithm in detail, here we give only a briefing. Our explanation centers around three topics: (1) calculation of burning surface area of complicated 3D grain; (2)calculation of burning surface area of grain with flaw; (3)calculation of burning surface area of grain with changing burning rate. We used explicit weighted essentially nonoscillatory (WENO) discretization scheme and total variation diminishing (TVD) Runge-Kutta scheme in numerical simulation. Results show preliminarily that this new method has high calculation precision and good adaptability. This paper can be used to coupled grain burning surface recession process to the numerical simulation of SRM (solid rocket motor) chamber flowfield.

Integrated Flow Field Calculation of Combustor and Inlet in Variable Flow Ducted Rocket

AUSM arithmetic and King boron particle combustion model have been efficiently implemented to numerical investigate integrated combustion chamber and inlet for variable flow ducted rocket. The effect of the ducted rocket performance from increscent gas flow was investigated. Numerical results indicated that the increase of gas flow reduced performance of ducted rocket with fixed inlet. The rocket thrust is enhanced, yet special impulse descends and flight of missile is reduced. In order to improve the performance, several importance factors, such as flight trajectory and air flux should be calculated synthetically and inlet design point should be selected and optimized reasonably for inlet design of variable flow ducted rocket.

Influence of Primary Rocket Mass Flow Rate on Performance of RBCC

Three-dimensional two-phase numerical method and ground direct-connect test system were used to investigate the influence of primary rocket mass flow rate with different inflow velocity on RBCC performance. Numerical results indicated that: under different inflow condition, the increment of primary flow had different contribution to net thrust and specific impulse; under low mach inflow condition, the increase of primary flow had positive effect on secondary combustion and greatly boosted thrust; under high mach condition, increasing primary flow would chock flow path, and weakened the influence of primary mass flow rate changing on RBCC thrust. Experimental results validated the rule obtained by numerical investigation between primary flow and RBCC performance. Especially under high mach inflow condition, with the Mach number increasing, the contribution of increment of primary flow to RBCC performance was reduced. Both the numerical and experimental investigation results could show that whether in low mach or high mach inflow condition, an optimized primary mass flow rate existed, which was benefit for improving RBCC performance.

Numerical Simulation of Combustion Instability in a Model Ramjet Combustor

Ramjets exhibit undesirable combustion instabilities under certain operating conditions. In this paper, large-eddy simulation(LES) and premixed combustion model have been efficiently implemented to study combustion instabilities in ramjet combustors. The main objective is to predict dominant frequencies and various flow features of the combustors. Combustion instabilities in the ramjet dump combustor have been numerically simulated. Low-frequency, large-amplitude instabilities are observed, and instability frequencies and flame spread during various operating conditions are in good agreement with experimental observations. The results show large vortex structures dominate the flame propagation and vortex convection coupled unsteady heat release excites pressure oscillation in ramjet combustor.

Heat Transfer Analysis and Thermal Structure Design of RBCC Engines

RBCC (rocket-based combined cycle) engine has several operation modes (ejector mode, ramjet mode, scramjet mode and sometimes pure rocket mode) in the whole flight course. In different mode, inner flow field is different and also there are big differences of heat transfer characteristics. These characteristics make it difficult to design thermal structure. In this full paper, source term method is raised to simulate combustion without solving ingredient equations. Base on source term method, CFD technology is used to forecast strut and main flow path thermal environment when RBCC engine works on each operation mode. Wall heat flux is gotten by the simulation. According to the results of forecast, a thermal structure scheme of composite structure based on CMC materials is raised. Mathematical models about designing the thermal structure are established and solved. Temperature distribution of each structure layer is gotten, and results have proved the feasibility about this thermal structure. All of these works establish a base to make RBCC engines reusable finally.

Experimental Investigation of Rocket Ejector in SMC Combustion Mode

*† ‡ Experimental investigation is carried out to explore the factors affecting rocket ejector performance. This paper mainly investigates the effect of ejector combustor geometry and primary rocket chamber pressure to ejector rocket performance in SMC combustion mode, for which fuel rich gas oxygen/alcohol gas generator and geometry adjustable ejector combustor are adopted. The experimental results show that ejector combustor geometry is the key factor to ejector rocket relative thrust in SMC combustion mode, but primary rocket chamber pressure play a subordination role. By optimizing ejector combustor geometry, ejector rocket in SMC combustion mode can obtain comparable thrust augmentation.

Nozzle Admittance and Damping Analysis Using the LEE Method

Abstract The nozzle admittance is very important in the theoretical analysis of nozzle damping in combustion instability. The linearized Euler equations (LEE) are used to determine the nozzle admittance with consideration of the mean flow properties. The acoustic energy flux through the nozzle is calculated to evaluate the nozzle damping upon longitudinal oscillation modes. Then the parametric study, involving the nozzle convergent geometry, convergent half angle and nozzle size, is carried out. It is shown that the imaginary part of the nozzle admittance plays a non-negligible role in the determination of the nozzle damping. Under the conditions considered in this work (f*=1,000 Hz, de*=0.18 m), the acoustic energy flux released from the nozzle with a 30o convergent half angle is highest (30o: 6.0×104kgs−3

Numerical Simulation of Open Active Control of Low Frequency Combustion Instability Related to Vortex Shedding in Model Ramjet Dump Combustor

6.0 \times {10^4}{\rm{kg}}{{\rm{s}}^{- 3}}

Experimental rocket motor applied in evaluate the ablation performance of ablation resistance materials

, 45o: 5.2×104kgs−3