Bond Stress Transducer Design for Solid Propellant Rocket Motors

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

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.

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.

Several launch vehicle solid rocket motor propellants contain ammonium perchlorate as a major ingredient, and so the potential exists for perchlorate release to the environment during launch operations and propellant manufacture. Solid composite base propellants containing four types of binder were studied: hydroxyl-terminated polybutadiene, carboxyl-terminated polybutadiene, polybutadiene-acrylic acid-acrylonitrile terpolymer, and polyurethane. The rate of perchlorate release from solid propellant during immersion in deionized and simulated seawater has been measured at 5, 20, and 29°C. The rate-limiting step is most probably diffusion, and the diffusion coefficient is given. Arrhenius terms are presented to facilitate calculation of diffusion coefficients at intermediate temperatures. The diffusion coefficients for all propellant types and conditions tested fall in the range from 3.6 X 10 - 1 2 to 1.1 × 10 - 1 3 m 2 s - 1 . The diffusion coefficients for each propellant type are proportional to temperature and inversely related to salinity.

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.

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.

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.

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.

A comprehensive analytical model which considers time and space development of the flow field in solid propellant rocket motors with high volumetric loading density is described. The gas dynamics in the motor chamber is governed by a set of hyperbolic partial differential equations, that are coupled with the ignition and flame spreading events, and with the axial variation of mass addition. The flame spreading rate is calculated by successive heating-to-ignition along the propellant surface. Experimental diagnostic studies have been performed with a rectangular window motor (50 cm grain length, 5 cm burning perimeter and 1 cm hydraulic port diameter), using a controllable head-end gaseous igniter. Tests were conducted with AP composite propellant at port-to-throat area ratios of 2.0, 1.5, 1.2, and 1.06, and head-end pressures from 35 to 70 atm. Calculated pressure transients and flame spreading rates are in very good agreement with those measured in the experimental system.

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.

The relative merits and possible applications of liquid and solid propellant rocket motors have long been topics for discussion and controversy. The arguments for and against have been sharpened by the fact that the development of the two types of motor are usually vested in quite different teams of people. At the same time, a fair comparison has often been confused because the ultimate choice has been made for reasons which are really not technical at all, but instead depend on availability and the utilisation of already deployed capital resources. Nevertheless, the merits and de-merits of the “conventional” (as opposed to the packaged) liquid motor and the solid motor can be simply defined with reasonable accuracy. The liquid motor is complicated, demanding at least some, and often a great deal, of attention before use. At the same time, it has a much greater performance potential, can be used repeatedly and can be designed to give variable thrust. The solid motor is simpler (although is getting less so), is cheaper (although this is getting less true) and demands less preparation. However, it has definite performance limitations, is inflexible in its output, and is often quite sensitive to its environment, whether during storage, carriage, or operation.

The performance prediction and grain burn-back analysis of rocket motor are important steps in the designing of a solid propellant rocket motor. The grain burn-back analysis of the solid grain identifies the burning surface area at each burning step in order to predict pressure-time history of the rocket motor. In this study, the shape of propellant grains was conveniently designed based on a solid modeling program of conventional purpose and the internal ballistics analysis was performed using a Matlab code which was developed to analyze the grain burn-back for this shape model. Upon carrying several analyses for rocket motors, it was confirmed that the developed code is suitable and useful.

The new solid propellants studied by SNPE Matériaux Energétiques (SME) are mainly based on: fillers, such as ADN, nitramines, RDX, HMX, HNIW, etc.; cross-linked binders formulated with energetic or non-energetic polymers, such as GAP and polyether polymer; and energetic plasticizers. The effect of nano-sized aluminium on the properties of HTPB/AP/Al is also assessed. In this paper, we will only discuss the effects of nano-sized aluminium on propellant properties and GAP-based propellants. Nano-sized aluminium leads to an increase of the burning rate of composite propellants. The effect greatly depends on the quality of the powder. For some formulations, a mesa effect is observed between 7 and 15 MPa. The use of GAP-based binder and HNIW brings an increase in volumetric specific impulse around 12% compared to current XLDB propellants. The mechanical properties are good. GAP-based binder and HNIW lead to an increase in the burning rate. In its industrial plant, SME has recently set up a new mixing facility equipped with a twin-screw machine. This workshop can produce current solid composite propellants or new energetic formulations. The hazards are greatly reduced owing to the small quantity of propellant mixed at the one time. Cryogenic solid propellants are a new concept of chemical propellants that use frozen products in solid rocket motors. The objective is to combine the high performance of liquids and the simplicity of solids. SME is developing a formulation with a freezing point near 0°C to limit the storage constraints; these are called Refrigerated Solid Propellants and are named Hydroxalane™. They offer opportunities to reach mid-term high-energetic performances. The compositions are based on water, hydrogen peroxide, polymer, and aluminium. BATES grains with a 90-mm diameter have been manufactured and tested successfully. The high energy density materials carry great potential in terms of increased energetic performance and are likely to introduce a long-term breakthrough for solid propulsion. These molecules require the use of a particular process for their synthesis study. The methodology used and the advance of research will be described herein.

Aging effects can have a significant consequence on the mission readiness of solid propellant rocket systems. A potential end result of grain cracking caused by aging is the catastrophic failure of the rocket motor. Propellant changes leading to grain cracking are an increase in cross -link density and/or induced dam age. The detection of this crack generating degradation within the solid propellant is critical. Initial manufacturing variability creates rocket motor assets with varying vulnerability to grain cracking. This significantly raises maintenance requirement s, reduces operational life times and threatens mission readiness. Several experimental methods are currently used to determine the cross link density of solid propellant. Unfortunately, these cross -link density measurement methods are all destructive te st determinations meaning the solid rocket propellant material must be dissected out of the main rocket asset for analysis. Dissection work is extremely expensive in terms of manpower, materials and money. Current non -destructive evaluation (NDE) and non -destructive inspection (NDI) techniques offer limited capability for the determination of cross -link density and/or induced damage. Photon Induced Positron Annihilation (PIPA) has demonstrated that non -destructive results on strain buildup and grain crac king in solid rocket propellant can be determined using the PIPA technique and that propellant damage could be detected using this technique through the rocket case of actual rocket systems, which could result in accurate and reliable assessments of compon ent readiness either at the manufacturing level or during field inspections. This NDI process will provide reliable material damage characterization including the buildup of defects and changes in the atomic structure of the propellant without disassembly and at detection levels below current NDI methods.

The paper investigates the numerical simulation of vortex shedding in the flow field of solid-propellant rocket motors. This phenomenon, resulting from the strong coupling between shear-layer instability and acoustic waves in the chamber, produces thrust and pressure oscillations. Numerical simulations are performed on the combustion chamber of the Ariane 5 MPS P230 (Solid Rocket Motor) with the commercial code Fluent CFD for conditions corresponding to 89 s of combustion time. The objective of the study is to reproduce the pressure oscillations frequencies and magnitudes, to compare the available experimental data and to capture the vortex shedding phenomena.