Packaged Liquid Rocket Motors

Small, low-thrust, long-burn-time solid propellant rocket motors could provide propulsion for a new class of kilogram-scale, transonic, uncrewed aerial vehicles (UAVs). This paper investigates technological challenges of small, low-thrust solid rocket motors: slow-burn solid propellants, motors that have low thrust relative to their size (and thus have low chamber pressure), thermal protection for the motor case, and small nozzles that can withstand long burn times. Slow-burn propellants were developed using ammonium perchlorate and 0–20% oxamide (burn-rate suppressant), with burn rates of at 1 MPa. Using these propellants, a low-thrust motor successfully operated at a thrust/burn area ratio 10 times less than that of typical solid rocket motors. This kilogram-scale motor can provide 5–10 N of thrust for 1–3 min. An ablative thermal protection liner was tested in these firings, and a new ceramic-insulated nozzle was demonstrated. This paper shows that small, low-thrust solid motors are feasible and presents a baseline design for the integration of such a motor into a small UAV.

Hybrid propellant rocket engines (HPRE) have been under development and advancement since the beginning of rocket propulsion research, alongside liquid and solid propellant rocket motors. However, its pace was much slower than that of the other propulsion types. Several companies have proposed hybrid launch vehicles in the last few years, and research and publications on such technology have increased significantly. One of the most crucial parts of HPREs is the pre-combustion chamber. This device is positioned upstream of the combustion chamber and downstream of the injection head, and it greatly affects the engine’s performance, size, weight, and cost, and combustion stability. Therefore, propulsion engineers are intensely interested in understanding the processes occurring inside pre-combustion chambers. Unfortunately, there are few published works on the design of such systems. Additionally, many references only recommend empirical size relations, which may only serve as a design starting point. Observing this issue, this paper compiles all the methods for designing pre-combustion chambers and provides rich discussions on each. These methods are: 1) experiments, 2) empirical relations, 3) numerical simulations, 4) droplet vaporization models, 5) combustion instability models, 6) droplet collision outcome prediction, and 7) molecular dynamics. The positive and negative aspects of each method are also considered. Moreover, other components that affect the calculation are discussed, such as injectors and grain design. Finally, a development workflow recommendation was constructed by applying all the discussed methods to reduce related costs and obtain an optimized design.

In solid propellant rocket motors, the design of grain port geometry primarily dictates the motor ballistic performance. The flexibility offered by the spatial variation of the port shape on motor performance is often discounted due to the requirement of complex mandrel shapes along with their associated manufacturing issues coupled with propellant processing constraints. Literature data indicate improvement in motor volumetric loading and stability by planned distribution of burn surface area along motor length. This paper deals with the realization and testing of a subscale motor with central finocyl-based grain geometry. A novel method of casting complex grain shapes in solid rocket motors using the wax mandrel technique is discussed. The wax mandrel preparation, propellant casting procedure, and wax mandrel removal setup are elaborated. Static testing results of the subscale motor with central finocyl port geometry are also presented as proof-of-concept. This work has demonstrated the feasibility of realizing solid rocket motor grains using the wax mandrel technique, which further enables realization of complex grain designs necessary to meet various mission requirements.

Abstract: Solid propelled rocket motors (SRMs) static firing tests are crucial tests in the aerospace industry while developing new motors as well as to ensure the quality of a motor batch. The operator can use this sort of test to determine whether the motor performance meets the project criteria by obtaining the measured "thrust versus time of burning" graphic the motor produces while burning. In this paper over all thrust measurement uncertaintyof a solid propellant rocket motor test bed is studied. During static test of solid motors requires the dynamic characteristics of the Static test configuration. To find out the magnitude of thrust oscillation, characterization of test setup and force transfer characteristics from motor case to load cell are required.

In the present research work, the effect of the igniter charge mass on the initial pressure rise in single grain solid propellant rocket motor is investigated experimentally. The igniter in the rocket motor is used to ignite propellant in rocket motor. Experiments are done with varying igniter charge mass by approximately ±15% of the calculated igniter charge mass value i.e. ~35g. Maximum pressure noted at 35g calculated igniter charge mass is 8.5632 MPa. Various experiments with 25g, 30g, 40g and 45g of igniter charge mass is carried out to study the effect of igniter charge mass on maximum pressure in rocket motor keeping all other parameters i.e. nozzle throat diameter, nozzle end grid, head end grid etc. are kept same. For each igniter charge mass two static tests were conducted. As igniter charge mass is increased from 25g to 45g ignition transient pressure increased from 2.9 MPa to 5.65 MPa, by optimizing the igniter charge mass the initial pressure can be kept within required range of MEOP (Maximum Expected Operating Pressure) for safe working of rocket motor.

A CFD simulation of the flowfield within a rod-and-tube solid propellant rocket motor has been developed. It couples the fluid dynamics and heat transfer of the gas flowfield within the rocket port to the nozzle and the regression rate of the propellant. The propellant regression is initially described with an empirical power law relationship and is then described with a provision for erosive burning phenomena. The erosive burning model is based on a phenomenological heat transfer approach. By incorporating regression into the model, the information provided by the resulting time-accurate solution will enable a much improved understanding of the flow phenomena within this rod-and-tube grain motor, which in turn will help in the design of both the motor and the nozzle.

Large solid propulsion mode has been used since the beginning of space launch activities and is currently operated on numerous space launchers in the world. The paper reviews the current situation of large solid propulsion for applications on space launchers on the 2003-2008 period, giving an idea of what is this business in term of produced volume and industrial organization. An engineering overview of the currently operational solid rocket motors (SRMs) is presented reporting the technical characteristics of the products themselves. The main identified interests of this type of propulsion - large thrust level for take off within a small volume, possibilities to adapt thrust law versus trajectory constraints, etc …- are analyzed and discussed. A tentative 'segmentation' of the technical function at launcher level is proposed identifying three main missions for this propulsion mode: (i) stage one of heavy launchers, (ii) Strap-on boosters for launchers based on a main liquid core stage, (iii) propulsion systems of quite all main stages of small launchers. Specific technical issues and constraints regarding each type of applications are analyzed like overall motor general design and performances, motor integration and interaction issues with the launcher system, thrust transmission, thrust orientation function, etc…... Past decade worldwide new motor development activities and evolutions of existing motors are reported in order to 'detect' the main trends of solid rocket motors design evolution. Performance increases through overall motor size growth, inert mass reduction and specific impulse progress are reported, giving a better understanding of the technological potentials of this propulsion mode. Identified new worldwide projects using solid propulsion motors are also briefly reported to give a glance on the future.

The development of a parallel adaptive mesh refinement (AMR) scheme is described for solving the governing equations for multi-phase (gas–particle) core flows in solid propellant rocket motors (SRMs). An Eulerian formulation is used to describe the coupled motion between the gas and particle phases. A cell-centred upwind finite-volume discretization and the use of limited linear reconstruction, Riemann solver based flux functions for the gas and particle phases, and explicit multi-stage time-stepping allows for high solution accuracy and computational robustness. A Riemann problem is formulated for prescribing boundary data at the burning surface and a mesh adjustment algorithm has been implemented to adjust the multi-block quadrilateral mesh to the combustion interface. A flexible block-based hierarchical data structure is used to facilitate automatic solution-directed mesh adaptation according to physics-based refinement criteria. Efficient and scalable parallel implementations are achieved with domain decomposition on distributed memory multi-processor architectures. Numerical results are described to demonstrate the capabilities of the approach for predicting SRM core flows.

Covers advancements in spacecraft and tactical and strategic missile systems, including subsystem design and application, mission design and analysis, materials and structures, developments in space sciences, space processing and manufacturing, space operations, and applications of space technologies to other fields.

Erosive burning rate studies of composite propellants under high cross flow velocities

HERE are imperfections existing in solid propellants. These imperfections may be produced during the manufacturing and/or fabricating process of the solid rocket motor. In analyzing the strength of a solid propellant, the localized imperfection may be idealized as a crack in the material. Besides the idealized crack, large cracks can also develop in solid motors during storage, handling, firing, etc. Therefore, to determine the ultimate strength of solid rocket motors, studies should be conducted to determine the significance of flaw type, size, and rate of growth. In order to conduct these studies, a detailed knowledge of crack growth behavior in solid rocket motors is indispensable. It is well known, on the microscopic scale, that a highly filled composite solid propellant can be considered a nonhomogeneous material. The nonhomogeneous nature of solid propellants is mainly due to the nonuniform distribution of the filler particles, the different sizes of filler particles used, the variation of bond strength between the particles and the binder, and the different crosslink densities of polymer chains. The presence of the filler particles affects not only the bulk properties of the material but also the crack growth behavior. For instance, depending upon the bond strength between the particle and the binder, the fracture path may preferentially either follow or avoid the filler particles, resulting in different fracture surfaces and crack growth rate. In addition to the nonhomogenity of the material, the material may also contain randomly spaced microvoids, incipient damage sites, and microcracks with statistically distributed sizes and directions. Therefore, when this material is strained, highly nonhomogeneous local stress and strength fields can be produced. Hence, the failure location and the degree of damage induced in the material will vary in a random fashion. Because the crack growth behavior is closely related to the local stress and strength fields as well as to the damage state in the material, the variation of these parameters in the material will undoubtedly be a contributing factor to the variability or scatter of

In an effort to better understand the properties of rocket noise, the statistical properties of noise data from various-sized solid propellant rocket motors. Time waveform data sampled at 204.8 kHz using 6.35 and 3.18 mm microphones were collected near motors with nozzle exit diameters ranging from 0.13 to 1.22 m. Non-Gaussian features of the data have been explored by calculating estimates of the probability density functions of the data and various statistical moments, including skewness and kurtosis. This has been carried out for both the pressure waveform and its first order time difference to better understand the formation of acoustic shocks within the noise. The analysis shows greater similarity between the statistics for the pressure than for the time derivative estimates.

3-D grain burnback analysis of solid propellant rocket motors: Part 1 – ballistic motor tests

Five variants of calculating the burning rate of a solid propellant as a function of the pressure in a solid-propellant rocket motor are considered. Two variants of analytical expressions are proposed for approximating real dependences. In all variants, the pressure in the rocket motor can be presented by simple analytical expressions as a function of solid propellant parameters, charging conditions, and structural factors of the charge and motor.

The processes of coagulation and aerodynamic fragmentation of liquid particles of aluminum oxide in an accelerating gas flow in the Laval nozzle are analyzed. A formula obtained by an approximate analytical solution of equations of a two-phase flow is proposed to calculate the characteristic particle diameter at the nozzle exit. The limiting particle diameter in the nozzle throat calculated theoretically is close to the mean-mass diameter obtained by numerical simulation of polydisperse two-phase flows with particle coagulation and fragmentation. The formula proposed is in agreement with Hermsen's correlation dependences and is confirmed by numerous published data on measurement of the mean-mass diameter of aluminum-oxide particles in exhaust plumes of small-, medium-, and large-scale solid rocket motors. The formula contains physical parameters whose values are readily calculated and prescribed. The formula is tested by all the parameters that enter into it. Based on a comparison of theoretical calculations and numerous experimental data, the formula is recommended for prediction of the size of aluminum-oxide particles in exhaust plumes of various types of solid rocket motors. Key words: rocket motor, two-phase flow, aluminum oxide, fragmentation of drops.