Rocket Motor Performance Research Articles

An Approach to Analysing Erosive Characteristics of Two-Channel Combustion Chambers

To acquire the erosive characteristics of two-channel combustion chambers, a quasi-one-dimensional internal ballistics model is proposed. Combining the model with two different erosive burning models, the modified L-R expression, and the Mukunda-Paul expression, the flow parameters and the internal ballistics performance of a test solid rocket motor are computed. The results show good agreement with experimental data. According to the results, the more severe the erosion is, the earlier, longer, and gentler the tail-off stage becomes. During the tail-off stage, the dramatic drop of pressure leads to a low normal burning rate and makes it easier for erosion burning to occur. For this reason, notable erosive burning might appear during tail-off if the case-to-throat area ratio is extremely low. The results also show that flows in the inner channel and the outer channel are similar but not identical. This leads to different erosive burning behaviours in the two channels. Also, the two erosive burning rate models involved in this paper are compared. It seems that the M-P expression provides better results than the modified L-R expression does, since it reveals the threshold phenomena. Besides, the M-P expression has great advantages for its universality for most propellants and different SRM geometries.

Performance comparison of oxidizer injectors in a 1-kN paraffin-fueled hybrid rocket motor

To investigate the effects of oxidizer injection on the performance of hybrid rocket motors (HRMs), we have designed, manufactured and tested four types of injector: showerhead (SH), hollow-cone (HC), pressure-swirl (PSW) and vortex (VOR). This study is motivated by the fact that the experimental measurements of N2O/paraffin firings are poorly presented in the open literature. Besides few publications are dedicated to the characterization of novel types of injectors in hybrid rocket propulsion application such as HC, PSW, and VOR. It is advantageous that the study was conducted in the same motor configuration, with the advantage that it allowed to compare the performance of different types of injectors.This paper analyzes the influence of the oxidizer injector design on the main performance parameters, such as fuel regression rate, specific impulse and combustion efficiency. First, in order to observe injector spray qualities, a series of cold tests using liquid water and liquid nitrous oxide are carried out, providing a good understanding of the spray profiles. Then, the motor performance data is obtained by a series of firing tests using N2O as oxidizer and paraffin as fuel. Comparison of the various injectors data is made with the same average oxidizer mass flux and feeding pressure. The showerhead is used as a benchmark in this study. During this experimental analysis, the VOR injector exhibits the highest regression rate, followed by HC and SH. Because the assessment of the regression rate was not enough to explain all the effects of the injectors on the motor performance, firing tests of small-scale hybrid motor have been carried out. In terms of spray properties, the PSW exhibits significant differences compared with the other injectors; it generates the smallest Sauter Mean Diameter (SMD) in the formed spray and achieves good atomization. In spite of the fact that the PSW injector leads to the lowest regression rate, it provides good specific impulse, increases the oxidizer-to-fuel ratio (O/F), as well as a uniform and smooth consumption of the paraffin fuel grain. VOR leads to the highest specific impulse. In terms of stability, VOR, HC and SH injectors exhibit lower oscillations in the chamber pressure. Some observations are made on exhaust plume intensity developed during combustion, and in firing tests with SH a blow-out phenomenon occur often. Similarly to liquid engines, it is possible in hybrid motors to increase the global propulsive performance using alternative designs of the injection system.

Parametric Study on Taper-ended Tubular Solid Propellant Grains

The design of the solid propellant grain is a decisive aspect of the solid propellant rocket motor performance. Tubular grain design is a favorable design since it produces a high neutral (time-invariable) thrust-time profile. However, neutrality of tubular grains deteriorates as the aspect ratio of the grain deviates from an optimum value that is dependent on the web thickness. In some cases, the undesirable phenomenon of erosive grain burning may take place. One simple solution to restore neutrality is to add taper to the ends of the grain. Loss of motor filling comes as penalty for adding these tapered ends. The grain should thus be tailored to simultaneously satisfy both desired design objectives namely, neutrality and filling. The present paper aims to address the dependence of these two design objectives on the design of a taper-ended tubular grain. The designs that are likely to yield erosive burning are also addressed. A parametric study is conducted involving the aspect ratio of the grain, its web thickness, and the taper angles on both ends. These parameters were varied, respectively, in the ranges 0.5: 5, 0.1:0.9, and 10o:90o. It was concluded that neutrality and filling of these grains are two competing objectives. The maximum neutrality is achieved using a grain with aspect ratios, web thickness, front and rear taper of 2:2.5, 0.1, 50o, and 30o, respectively. In contrast, the maximum filling can be achieved using a grain with aspect ratios and web thickness of 1:3.5 and 0.8, respectively, with no taper at both ends. A compromised grain can have aspect ratio and web thickness of 2 and 0.8, respectively, with no taper at both ends.

Internal Ballistic Modeling of a Solid Rocket Motor by Analytical Burnback Analysis

The internal ballistic analysis of a solid rocket motor mainly relies on an accurate computation of the transient burning surface geometry of the propellant during combustion phase. A reliable and an efficient burnback analysis is crucial in order to predict the ballistic performance of a solid rocket motor in early design processes. Because analytical burnback solutions provide exact results in the most affordable time, derivation of analytical solutions for generic propellant geometries is of practical value and high importance for rocket designers. This paper summarizes a research effort for developing a computational tool based on burnback analytics for all possible different geometric configurations of a generic slotted-type propellant. After the burnback solutions were verified and implemented into an in-house code, this parametric and analytical tool is integrated into a zero-dimensional internal ballistic solver to enable an efficient overall computational framework to be employed for multidisciplinary design and optimization of solid rocket motors.

Experimental investigation of fuel composition and mix-enhancer effects on the performance of paraffin-based hybrid rocket motors

Experimental investigations have been performed for a lab-scale hybrid rocket motor with hydrogen peroxide and paraffin-based propellants. A new paraffin-based solid fuel with favorable mechanical strength, consisting of 50 wt% paraffin, 20 wt% PE wax, 18 wt% EVA, 10 wt% SA and 2 wt% carbon, was proposed. The thermal properties of the proposed paraffin-based fuel and its components were firstly evaluated using TG/DTG measurements. It was found that the decomposition rates of each of the components of this paraffin-based fuel follow the order SA > pure paraffin > paraffin-based fuel blends > PE > EVA. Measurements of the regression rate of this paraffin-based fuel were performed and compared with results found in the literature. This showed that the regression rate of this paraffin-based fuel is much higher than that of traditional polymer solid fuels such as HTPB fuels. Finally, by adding a protrusion, a cavity and swirl blades at the end and in the middle of the combustion chamber, experimental investigations into the effects of a mix-enhancer on the paraffin-based hybrid rocket motor were performed. These investigations are discussed in detail.

HYBRID OPTIMIZATION OF STAR GRAIN PERFORMANCE PREDICTION TOOL

ABSTRACTIn solid propellant rocket propulsion, the design of the propellant grain is a decisiveaspect. The grain design governs the entire motor performance and, hence, thewhole rocket mission. The ability to decide, during design phase, the proper graindesign that satisfies the predefined rocket mission with minimum losses is theultimate goal of solid propulsion experts. This study enables to predict the pressuretime curve of rocket motor with star grain configuration and also to optimize theperformance prediction tool through optimization methods to maximize its predictionefficiency. A hybrid optimization technique is used. Genetic Algorithm (GA) is firstimplemented to find the global optimum followed by Simulated Annealing (SA)optimization method to find the accurate local optimum. A program for predicting thepressure time curve of the rocket motor is created on MATLAB and then linked to GA- SA optimizers as an application on a case study. The purposed approach isvalidated against satisfying data. It is found that the developed optimized program iscapable of predicting rocket motor performance (including the effect of erosiveburning) with acceptable accuracy for preliminary design purposes.

Design and Implementation of a Thrust Vector Control (TVC) Test System

The rocket engines are tested statically to evaluate the performance of engine based upon thrust produced. One of the most important parameters of the rocket engine static testing evaluation is to measure the thrust produced by the engine. The thrust produced is measured using a Thrust Vector Control (TVC) test system which is a structural element equipped with load cells. In this study, a load sensor system was designed to measure the propulsion performance of a solid propellant rocket motor. The forces and moments of the rocket motor with respect to the six degrees of freedom of the test system were measured during firing. It is seen that the obtained experimental results and the analysis results are compatible with each other. The designed stand is capable of measuring axial thrust and lateral (misaligned) thrust components, and the rolling moment for rocket motors producing axial thrust up to 50 [kN].

Parametric investigation of secondary injection in post-chamber on combustion performance for hybrid rocket motor

The objective of this effort is to study the combustion performance of a hybrid rocket motor with the help of 3D steady-state numerical simulation, which applies 90% hydrogen peroxide as the oxidizer and hydroxyl-terminated polybutadiene as the fuel. A method of secondary oxidizer injection in post-chamber is introduced to investigate the flow field characteristics and combustion efficiency. The secondary injection medium is the mixed gas coming from liquid hydrogen peroxide catalytic decomposition. The secondary injectors are uniformly set along the circumferential direction of the post-chamber. The simulation results obtained by above model are verified by experimental data. Three influencing parameters are considered: secondary injection diameter, secondary injection angle and secondary injection numbers. Simulation results reveals that this design could improve the combustion efficiency with respect to the same motor without secondary injection. Besides, the secondary injection almost has no effect on the regression rate and fuel sueface temperature distribution. It is also presented that the oxidizer is injected by 8 secondary injectors with a diameter of 7–8 mm in the direction of 120°in post-chamber is identified as the optimized secondary injection pattern, through which combustion efficiency, specific impulse efficiency as well as utilization of propellants are all improved obviously.

Experimental and theoretical study on a dual-thrust rocket motor with subsonic intermediate nozzle

Dual-thrust rocket motors, with proper designs, can yield the desired boost-sustain thrust profiles for a variety of applications. Of the various designs that a dual-thrust rocket motor can have, the one with an intermediate nozzle provides a high boost-to-sustain thrust ratio as well as a stable operation. Depending on the internal conditions, the flow through the intermediate nozzle in the boost phase can be sonic, subsonic, or initially reversed. The present paper discusses the operation and performance of Dual-thrust rocket motors with subsonic intermediate nozzle. A mathematical model describing the internal ballistics of these dual-thrust rocket motors is derived from the basic governing equation. In addition, a set of experiments is conducted using a developed test motor to explore the physics of dual-thrust rocket motors with subsonic intermediate nozzle and assess the mathematical model validity. Results of the theoretical and experimental investigations explain most of the internal ballistics characteristics that were not discussed before in the open literature.

Effect of grain port length–diameter ratio on combustion performance in hybrid rocket motors

The objectives of this study are to develop a more accurate regression rate considering the oxidizer mass flow and the fuel grain geometry configuration with numerical and experimental investigations in polyethylene (PE)/90% hydrogen peroxide (HP) hybrid rocket. Firstly, a 2-D axisymmetric CFD model with turbulence, chemistry reaction, solid–gas coupling is built to investigate the combustion chamber internal flow structure. Then a more accurate regression formula is proposed and the combustion efficiency changing with the length–diameter ratio is studied. A series experiments are conducted in various oxidizer mass flow to analyze combustion performance including the regression rate and combustion efficiency. The regression rates are measured by the fuel mass reducing and diameter changing. A new regression rate formula considering the fuel grain configuration is proposed in this paper. The combustion efficiency increases with the length–diameter ratio changing. To improve the performance of a hybrid rocket motor, the port length–diameter ratio is suggested 10–12 in the paper.

Flame speed enhancement of solid nitrocellulose monopropellant coupled with graphite at microscales

The flame-speed-enhancement phenomenon of a solid monopropellant (nitrocellulose) using a highly conductive thermal base (graphite sheet) was demonstrated and studied both experimentally and theoretically. A propellant layer ranging from 20 μm to 170 μm was deposited on the top of a 20-μm thick graphite sheet. Self-propagating oscillatory combustion waves were observed, with average flame speed enhancements up to 14 times the bulk value. The ratio of the fuel-to-graphite layer thickness affects not only the average reaction front velocities but also the period and the amplitude of the combustion wave oscillations. To better understand the flame-speed enhancement and the oscillatory nature of the combustion waves, the coupled nitrocellulose-graphite system was modeled using one-dimensional energy conservation equations along with simple one-step chemistry. The period and the amplitude of the oscillatory combustion waves were predicted as a function of the ratio of the fuel-to-graphite thickness (R), the ratio of the graphite-to-fuel thermal diffusivity (α0), and the non-dimensional inverse adiabatic temperature rise (β). The predicted flame speeds and the characteristics of the oscillations agree well with the experimental data. The new concept of using a highly conductive thermal base such as carbon-based nano- and microstructures to enhance flame propagation speed or burning rate of propellants and fuels could lead to improved performance of solid and liquid rocket motors, as well as of the alternative energy conversion microelectromechanical devices.

비행제어를 위한 비행 중 고체로켓 추력 예측 방법

In this paper, an in-flight prediction method of thrust profiles for solid rocket motors is proposed. Actually, it is very difficult to have detailed information about the performance of the rocket motors beforehand because it is quite sensitive to combustion environments. To overcome this problem, we have developed an algorithm for generating in-flight prediction of rocket motor performance in realistic environments via a reference burnback profile and accelerations measured at a short time-interval just after launch. The performance is evaluated through a lot of flight test results.

ALUMINIZED POLYETHYLENE REGRESSION RATE EVALUATION FOR HYBRID ROCKET MOTOR

Hybrid rocket solid fuel grain regression rate is one of the significant values to accurately determine in the hybrid rocket motor (HRM) design parameters and performance prediction. The comprehensive method to measure and evaluate HRM solid fuel regression rate was investigated theoretically and experimentally .The mathematical model has been introduced and a computer program was built to describe HRM regression rate. A series of experimental tests using a small scale HRM employs Polyethylene (PE) and aluminized PE as solid fuel (up to 15% by mass metal powder additive) with different initial port diameter of fuel grain (5, 20 and 28mm) and gaseous oxygen (GO2) as oxidizer to evaluate the HRM regression rate performance. The measured experimental data are compared with results obtained from theoretical computation. The comparison shows good agreement, which proves the validity of the developed program. Regression rate of hybrid fuel grain was enhanced by addition of Al powder compared to basic configuration (0% Al) and it was found that it is more significant to use Al powder additive increased up to 7.5% by mass

Fluid–solid coupled simulation of the ignition transient of solid rocket motor

The first period of the solid rocket motor operation is the ignition transient, which involves complex processes and, according to chronological sequence, can be divided into several stages, namely, igniter jet injection, propellant heating and ignition, flame spreading, chamber pressurization and solid propellant deformation. The ignition transient should be comprehensively analyzed because it significantly influences the overall performance of the solid rocket motor. A numerical approach is presented in this paper for simulating the fluid–solid interaction problems in the ignition transient of the solid rocket motor. In the proposed procedure, the time-dependent numerical solutions of the governing equations of internal compressible fluid flow are loosely coupled with those of the geometrical nonlinearity problems to determine the propellant mechanical response and deformation. The well-known Zeldovich–Novozhilov model was employed to model propellant ignition and combustion. The fluid–solid coupling interface data interpolation scheme and coupling instance for different computational agents were also reported. Finally, numerical validation was performed, and the proposed approach was applied to the ignition transient of one laboratory-scale solid rocket motor. For the application, the internal ballistics were obtained from the ground hot firing test, and comparisons were made. Results show that the integrated framework allows us to perform coupled simulations of the propellant ignition, strong unsteady internal fluid flow, and propellant mechanical response in SRMs with satisfactory stability and efficiency and presents a reliable and accurate solution to complex multi-physics problems.

Horizontal jet characteristics of an underwater solid rocket motor at the beginning of working

Compared with the conventional solid rocket motor which works in atmosphere, the underwater solid rocket motor has different performance characteristics, because the buoyancy of water cannot be neglected, while in the atmosphere only gravity is present. In order to get a deep understanding of the mechanism of thrust oscillation characteristics and of the jet structure of underwater solid rocket motor, a 3-D numerical simulation using the volume of fluid (VOF) model is given in this paper to investigate the horizontal jet structure and the oscillating flow field induced by the gravity and buoyance. The principle of momentum is carried out to analyze the oscillating flow field in the initial working period. Result of analysis indicates that the behavior of rising gas bubbles is remarkably influenced by gravity and buoyancy. Compared to the conventional solid rocket motor in the atmosphere environment, the phenomenon shows that the rising gas bubbles are more obvious in the water. It is concluded that thrust-gravity and buoyancy coupling can generate bigger roll moment in the vertical direction than the conventional solid rocket motor. Numerical calculations in conjunction with experiment in the references proves that the model contains gravity, buoyancy and the VOF model can be used to describe the jet structure and thrust oscillation of underwater solid rocket motor; gravity and buoyancy cannot be ignored in the research of underwater solid rocket motor performance characteristics.

Conflicting Design Issues Hallmark the Compound Rocket Motors

A series of conflicting requirements to which the design of the compound rocket motors are subjected are considerably diminishing the efficiency of that type of rocket motors. The necessity of delivering a high thrust level at lift-off and during the sustainer phase up to optimal velocity accommodation, coupled with the requirement for a slander body configuration for air drag mitigation during the atmospheric ascent are the dominant issues that end into a compromise that lower the overall efficiency to a great extent, in comparison the available theoretical performance of compound rocket motors. The optimal design is imposing a lower mass ratio that expected and is the main cause of efficiency reduction when very high velocity requirements are searched for, like in orbital launchers. However, the margins of propulsive efficiency build-up can be conveniently manipulated through geometrical optimization and attentive risk management of the propulsion system as shown by the experimental results protruded during the development of the NERVA-ORVEAL space rocket motor.

Influence of Initial Propellant Temperature on Solid Rocket Internal Ballistics

The influence of initial propellant temperature on solid propellant rocket motor performance is examined in the present study, with the focus primarily on predicted steady and nonsteady burning rate behavior. The present evaluation includes an examination of quasi-steady burning of a conventional composite propellant under pressure, core flow (erosive burning), and normal acceleration (say, due to motor spinning). The most direct and well-known influence of initial propellant temperature is on the pressure-dependent burning rate. The pressure-dependent burning rate, as the base burning rate, in turn influences the degree of augmentation seen under core flow and normal acceleration; in general, the lower the base rate, the higher the augmentation. The explicit presence of initial propellant temperature in various model equations for the flow- and acceleration-dependent burning components, and the corresponding effect of the initial temperature on the propellant’s burning surface temperature, can produce an overall compounding effect that adds to the nominal augmentation produced by the base burning rate value. The same considerations apply when one moves from quasi-steady to transient operational conditions, where the transient burning rate response of the solid propellant (which is a function of the quasi-steady burning rate components, as well as the initial temperature of the propellant) can be modified significantly by varying the propellant’s initial temperature.

Dependence of Nano-Aluminum and Water Propellant Combustion on pH and Rheology

Over the past few years, the combustion of nano-aluminum/water (nAl/H2O) propellants has been widely reported, but further progress has been slowed for the following reason: the loosely correlated trends in combustion data are insufficient in guiding further research efforts, and they cannot be used to significantly improve the Isp observed in static rocket motor tests. It was previously found that different mixing techniques (hand, planetary and resonant mixers, duration and temperature), or equivalence ratio gave rise to different burning rates, but the influence of pH and rheology on nAl/H2O propellants was not considered. We find that the effects of pH on nAl/H2O propellants are profound, and correlate well with viscosity, low pressure deflagration limits, burning rate exponents, and rocket motor performance. Our findings suggest that coagulation can influence the pressure exponent over a wide range of values (0.34–0.68). For particle diameters <1 µm, dispersion during mixing is affected more by electrostatic repulsion from charged ions than from mechanical agitation, and this is reflected through zeta potential and viscosity measurements at different pH levels. Additionally, we observe that pH has an influence on nAl/H2O reaction kinetics during ignition, as the propellant transitions from low temperature oxidation to high temperature combustion.

Factors Affecting Burning Rate of Hydroxy Terminated Polybutadiene Propellant

Burning rate is one of the most important characteristics of rocket propellants and it is affected by many factors including composition, pressure, initial temperature, oxidizer particles size and burning rate modifier. This research aims to study experimentally the effects of combustion chamber pressure, initial temperature and particles size of the oxidizer materialon the burning rate of solid composite propellant. Besides, study theoretically the effect of composition and catalyst used. In addition, this research aims to show the relationship between these factors and the burning rate in order to evaluate factor that is more effective on burning rate, so to calculate of burning rate as a function of this factor. Accordingly, thiscould be used to achieve ballistics characteristics required for rocket motor design and performance.

Robust Design Optimization of Dual Thrust Solid Propellant Motors due to Burning Rate Uncertainties

Solid rocket motor design becomes a clamant task where fusion of design and non-design parameters may lead to various design configurations, especially under uncertainties. This paper proposes a robust design optimization method for the performance of a dual thrust solid rocket motor under investigation of the propellant burning rate and grain geometry uncertainties. It was found that due to uncertainties, the burning rate varies erratically during burning surface area regression, which may lead to catastrophic failure. The present approach aims at uncertainty quantification associated with burning rate and its influence on dual thrust motor performance. A first-order orthogonal design is applied to estimate the worst case deviation coupled with motor internal ballistics. The robustness assessment is measured directly by a mean-variance and average difference approach. A sensitivity analysis of performance parameters is performed to analyze the effects of variation of the design parameters. A hybrid genetic algorithm and simulated annealing approach is used as optimizer. The robust design solution obtained in the form of insensitive design and performance parameters shows the effectiveness and efficiency of the proposed methodology.