Cai Guo-biao

Numerical and experimental investigation on the effects of aft mixing chamber diaphragm in hybrid rocket motor

This paper focuses on the investigation of an aft mixing chamber diaphragm in a hybrid rocket motor. Both numerical and experimental researches are carried out to study its effects on the motor performances. The hybrid rocket motor with star fuel grain is utilized. The 90% hydrogen peroxide (HP) oxidizer and hydroxyl terminated polybutadiene (HTPB) based fuel are adopted as propellants. The diaphragm configuration settled in the aft mixing chamber includes four circular-holes with a uniform circumferential distribution. For both motors with and without the diaphragm, three-dimensional numerical simulations with gaseous combustions considered are carried out to study the flow field characteristics and motor performances. The comparison results show that the diaphragm can improve the mixing of the oxidizer and fuel. It has evident effect on increasing the motor efficiencies. Two firing experiments are conducted with full scale motors to investigate the effects of the diaphragm. The diaphragm used in the test is composed of a central steel framework and a closed thermal insulation layer. With the diaphragm employed, the combustion efficiency increases from 93.9% to 97.34% and the specific impulse efficiency increases from 80.77% to 87.28%, which verifies the positive effect of the diaphragm. Both numerical and experimental studies indicate that the scheme of the aft mixing chamber diaphragm proposed in the paper can improve the efficiencies of the hybrid rocket motor obviously.

Design and optimization of variable thrust hybrid rocket motors for sounding rockets

Besides safety and low-cost, the start/shutdown/restarting and throttling ability are the other two significant advantages of hybrid rocket motors (HRMs) compared with liquid and solid ones. In this study, a two-stage variable thrust and non-toxic 98%HP/HTPB hybrid rocket motor (VTHRM) is designed and applied in a sounding rocket, and the design parameters of the motor are analyzed and optimized. A computational program is developed to design the motor system structure, to predict the interior ballistics and the ballistic trajectory. A star grain and a wheel grain are compared. The design of experiment (DOE), variance analysis and the main effect analysis are employed to investigate the influence of the main design parameters on motor performance. The multidiscipline feasible (MDF) approach is applied to establish the optimization procedure after analyzing the system design structure matrix. A modified differential evolution algorithm is employed to maximize the load mass. The results indicate that the wheel grain could obtain a larger load mass and a lower length to diameter ratio, and that throttling markedly meliorates the motor and rocket performance. The conclusions drawn from the analysis and optimization could provide instructive guide and theoretical basis for engineering designs.

Single- and Multi-objective Optimization of Scramjet Components Using Genetic Algorithms Based on a Parabolized Navier-Stokes Solver

A 2D hypersonic inlet which can self-start at relay Mach number (Ma=3.5) is designed. Then single- or multi-objective optimization designs of the 2D inlet are carried out on cruise operation point (Ma=7.0) by Sequential Quadratic Programming (SQP), Multi-Island Genetic Algorithm (MIGA), and Multi-Objective Genetic Algorithms (MOGAs), i.e. Neighborhood Cultivation Genetic Algorithm (NCGA) and Nondominated Sorting Genetic Algorithm II (NSGA-II). In these optimizations, the inlet flowfields are calculated with the Single-Sweep Parabolized Navier-Stokes (SSPNS) algorithm codes which are proved highly accurate, highly efficient for supersonic/hypersonic flowfield simulations previously. Singleobjective optimization results show that pressure recovery maximum model is better than effective kinetic energy coefficient maximum model. Multi-objective optimization results reveal the tradeoffs among total pressure recovery, static pressure rise, and drag coefficients. Based on multi-objective design process, a multi-operation-point design and a design point determi-nation process are investigated. Results show if design point is set at cruise point, the off-design operation mass capture coefficients is relatively low; whereas if the design Mach number is 6.5, the 2D inlet will get good overall operation performances along the constant dynamic trajectory. The same optimization design methods are applied to design several 2D Single-Expansion-Ramp Nozzles (SERNs). Result of the single-objective optimization design shows that the objective, i.e. the thrust coefficient, is improved, but the corresponding lift coefficient is relatively low. In the 2- and 3-objective cases, tradeoffs of the thrust coefficient, the lift coefficient and the pitching moment coefficient, are obtained. Results show that cowl length and initial expansion angle of ramp influence the SERNs performance significantly. It is also shown that the shorter the cowl is, the higher the lift is; whereas the larger the initial expansion angle is, the smaller the pitching momentum is.

Scaling of the flowfield in a combustion chamber with a gas?gas injector

The scaling of the flowfield in a gas–gas combustion chamber is investigated theoretically, numerically and experimentally. To obtain the scaling criterion of the gas–gas combustion flowfield, formulation analysis of the threedimensional (3D) Navier–Stokes equations for a gaseous multi-component mixing reaction flow is conducted and dimensional analysis on the gas–gas combustion phenomena is also carried out. The criterion implies that the size and the pressure of the gas–gas combustion chamber can be changed. Based on the criterion, multi-element injector chambers with different geometric sizes and at different chamber pressures ranging from 3 MPa to 20 MPa are numerically simulated. A multi-element injector chamber is designed and hot-fire tested at five chamber pressures from 1.64 MPa to 3.68 MPa. Wall temperature measurements are used to understand the similarity of combustion flowfields in the tests. The results have verified the similarities between combustion flowfields under different chamber pressures and geometries, with the criterion applied.

Numerical and experimental study on shear coaxial injectors with hot hydrogen-rich gas/oxygen-rich gas and GH2/GO2

The influences of the shear coaxial injector parameters on the combustion performance and the heat load of a combustor are studied numerically and experimentally. The injector parameters, including the ratio of the oxidizer pressure drop to the combustor pressure (DP), the velocity ratio of fuel to oxidizer (RV), the thickness (WO), and the recess (HO) of the oxidizer injector post tip, the temperature of the hydrogen-rich gas (TH) and the oxygen-rich gas (TO), are integrated by the orthogonal experimental design method to investigate the performance of the shear coaxial injector. The gaseous hydrogen/oxygen at ambient temperature (GH2/GO2), and the hot hydrogen-rich gas/oxygen-rich gas are used here. The length of the combustion (LC), the average temperatures of the combustor wall (TW), and the faceplate (TF) are selected as the indicators. The tendencies of the influences of injector parameters on the combustion performance and the heat load of the combustor for the GH2/GO2 case are similar to those in the hot propellants case. However, the combustion performance in the hot propellant case is better than that in the GH2/GO2 case, and the heat load of the combustor is also larger than that in the latter case.

Experimental and Numerical Investigation of Large Mass Flow Rate Gas-Gas Injectors

†‡ This paper focuses on methodologies of investigation and design of large mass flow rate gas-gas injector. The experimental and CFD methods of GO2/GH2 single-element engine were presented at first. The combustor wall temperatures were measured using arrays of thermocouples instrumentation. The temperature increments along the axial direction had been obtained in experiments, which could indicate the general development and completion length of combustion. The combustion length shown in the results of numerical simulation of a typical shear coaxial injector at pressure of 3MPa and mass flow rate of 113g/s agreed with that of experiment. The typical CFD combustion flow field showed that the heat protection of shear-coaxial injector was well, but the combustion length was very long. The influences of key parameters on combustion were conducted, and the optimum-parameter shearcoaxial injector was designed and tested. The results of tests showed the optimum-parameter injector could obviously shorten the combustion length and the combustion was completed far before the throat at mass flow rate of 226g/s which was twice the mean mass flow rate of the SSME single main injector (SSMI). The results of CFD prediction showed that the optimum-parameter injector could undertake mass flow rate of 452g/s which was 4 times that of SSMI. The other new designs had been made based on studies of the shear-coaxial injector flow fields, and simulations and experiments were also conducted. The results showed that the O2 post expansion design could enhance the mixing and decrease the combustion length regardless of the expansion angle, H2 impinging design increased the combustion length, the 0.1mm thickness splitter plate rarely influenced the combustion length compared with 0.5mm, and the multi-hole design had a severe heat protection issue of injector faceplate.

Uncertainty analysis and probabilistic design optimization of hybrid rocket motors for manned lunar landing

To obtain a conceptual design for a hybrid rocket motor (HRM) to be used as the Ascent Propulsion System in the Apollo lunar module, the deterministic design optimization (DDO) method is applied to the HRM design. Based on the results of an uncertainty analysis of HRMs, an uncertainty-based design optimization (UDO) method is also adopted to improve the design reliability. The HRM design process, which is a multidisciplinary system, is analyzed, and a mathematical model for the system design is established to compute the motor performance from the input parameters, including the input variables and model parameters. The input parameter uncertainties are quantified, and a sensitivity analysis of the model parameter uncertainties is conducted to identify the most important model parameter uncertainties for HRMs. The DDO and probabilistic UDO methods are applied to conceptual designs for an HRM to be used as a substitute for the liquid rocket motor (LRM) of the Ascent Propulsion System. The conceptual design results show that HRMs have several advantages as an alternative to the LRM of the Ascent Propulsion System, including nontoxic propellant combination, small motor volume, and comparable functions, such as restarting and throating. Comparisons of the DDO and UDO results indicate that the UDO method achieves more robust and reliable optimal designs than the DDO method. The probabilistic UDO method can be used to develop better conceptual designs for HRMs.

Steady-state coupled analysis of flowfields and thermochemical erosion of C/C nozzles in hybrid rocket motors

A hybrid rocket can be used in various applications and is an attractive propulsion system. However, serious erosion of nozzles is common in motor firing operations, which could restrict the application of hybrid rocket motors. Usually, the serious erosion is attributed to the high concentration of oxidizing species in hybrid motors, while the details of flowfields in the motors are not paid special attention to. In this paper, first the thermochemical erosion of C/C nozzle is simulated coupled with the flowfields in a 98% H2O2/hydroxyl-terminated polybutadiene (HTPB) hybrid rocket motor. The simulation is made on a typical axisymmetric motor, including a pre-combustion chamber, an aft-combustion chamber and nozzle structures. Thermochemical reactions of H2O, CO2, OH, O and O2 with C are taken into account. Second, the change of flowfields due to fuel regression during motor firing operations is considered. Nozzle erosion in different flowfields is evaluated. Third, the results of nozzle erosion in the coupled simulation are compared with those under uniform and chemical equilibrium flow and motor firing test results. The results of simulation and firing tests indicate that the thermochemical erosion of nozzles in hybrid motors should be calculated coupled with flowfields in the motor. In uniform and chemical equilibrium flowfields, the erosion rate is overestimated. The diffusion flame in hybrid motors protects the nozzle surface from the injected oxidizer and high temperature products in flowfields, leading to a relatively fuel-rich environment above the nozzle. The influence of OH and the geometry of motor should also be considered in the evaluation of nozzle erosion in hybrid motors.

Simulation of two-phase plume field of liquid thruster

When the liquid propellant thruster works, its plume field would contain many propellant liquid droplets, especially at pulse state. Liquid droplets may move along with the gas flow and deposit on the components of spacecraft as contamination. The simulation of the plume field involving the gas molecules and liquid droplets is an important part in contamination studies of thruster plume. Based on the PWS software developed by Beihang University (BUAA), axial-symmetric two-phase direct simulation Monte Carlo (DSMC) method is used with the liquid droplet taken as a kind of solid particle. The computation of gas-to-particle effect and gas reflection on the particle surface are decoupled. The inter-particle collision is also considered. The gas parameters at nozzle exit of 120N engine after 20 ms pulse work are taken as the entrance condition of the numerical simulation. Four test cases are conducted for comparison of different collision modules. Simulation results show that the effects of liquid propellant droplets mainly concentrate near the axis line of engine. The particle-to-gas collision would cause evident differences in the gas field and subtle differences in the particle phase. The liquid droplets in the plume field are generally accelerated and convected by the gas molecules. The DSMC method is proved to be a feasible solver to numerically simulate the two-phase flow involving solid phase and rarefied gas flow.

Scaling of heat transfer in gas—gas injector combustor

The scaling of heat transfer in gas—gas injector combustor is investigated theoretically, numerically and experimentally based on the previous study on the scaling of gas—gas combustion flowfield. The similarity condition of the gas—gas injector combustor heat transfer is obtained by conducting a formulation analysis of the boundary layer Navier—Stokes equations and a dimensional analysis of the corresponding heat transfer phenomenon. Then, a practicable engineering scaling criterion of the gas—gas injector combustor heat transfer is put forward. The criterion implies that when the similarity conditions of inner flowfield are satisfied, the size and the pressure of gas—gas combustion chamber can be changed, while the heat transfer can still be qualitatively similar to the distribution trend and quantitatively correlates well with the size and pressure as q ∝ p0.8cd−0.2t. Based on the criterion, single-element injector chambers with different geometric sizes and at different chamber pressures ranging from 1 MPa to 20 MPa are numerically simulated. A single-element injector chamber is designed and hot-fire tested at seven chamber pressures from 0.92 MPa to 6.1 MPa. The inner wall heat flux are obtained and analysed. The numerical and experimental results both verified the scaling criterion in gas—gas injector combustion chambers under different chamber pressures and geometries.