Stage Solid Rocket Motors Research Articles

Experimental estimation of the amount of residual slag remaining inside orbital insertion stage solid rocket motor

Experimental estimation of the amount of residual slag remaining inside orbital insertion stage solid rocket motor

Small Solid Propellant Launch Vehicle Mixed Design Optimization Approach

For a small country with limited research budget and lack of advanced space technology, it is imperative to find new approaches for the development of low-cost launch vehicles (LV), which is, among all possibilities, an interesting option for rapid access to space, focused on integration of acquired components complemented with indigenously developed subsystems. This approach requires the cooperation of developed countries with huge experience and knowledge in LV development and operations. The main objective is to develop a small three stage solid propellant LV capable of delivering a payload of 100 kg to a circular low earth orbit of 600 km altitude, with the first and second stage solid rocket motors (SRM) hypothetically acquired from different countries and the third one designed and produced domestically in accordance with the production and technological capability. This approach provides main advantages such as: reduction in total time to access the space and to master the basic knowledge of launch operations. For this purpose, an integer continuous genetic algorithm global optimization method was selected and implemented, the SRM characteristics of the first and second stage were considered as integer variables, whereas the design variables of the third stage SRM and the trajectory variable were considered as continuous. A multi discipline feasible (MDF) framework was implemented along with the propulsion, aerodynamic, mass and trajectory models. Despite their particular characteristics and constraints, the results show highly acceptable values, and the approach proved to be reliable for conceptual design level.

Chemical Erosion of Carbon-Phenolic Rocket Nozzles with Finite-Rate Surface Chemistry

Ablative materials are commonly used to protect the nozzle metallic housing and to provide the internal contour to expand the exhaust gases in solid rocket motors. Because of the extremely harsh environment in which these materials operate, they are eroded during motor firing with a resulting nominal performance reduction. The objective of the present work is to study the thermochemical erosion behavior of carbon-phenolic material in solid rocket motor nozzles. The adopted approach relies on a validated full Navier–Stokes flow solver coupled with a thermochemical ablation model, which takes into account finite-rate heterogeneous chemical reactions at the nozzle surface, rate of diffusion of the species through the boundary layer, pyrolysis gas and char-oxidation product species injection in the boundary layer, heat conduction inside the nozzle material, and variable multispecies thermophysical properties. The results obtained with the proposed approach are compared with two sets of experimental data: subscale motor tests carried out for the space shuttle reusable solid rocket motor and the static firing tests of the second and third stage solid rocket motors of the European Vega launcher, which use carbon-carbon for the throat insert and carbon-phenolic for the region downstream of the throat.

Design and Optimization of 3D Radial Slot Grain Configuration

Upper stage solid rocket motors (SRMS) for launch vehicles require a highly efficient propulsion system. Grain design proves to be vital in terms of minimizing inert mass by adopting a high volumetric efficiency with minimum possible sliver. In this article, a methodology has been presented for designing three-dimensional (3D) grain configuration of radial slot for upper stage solid rocket motors. The design process involves parametric modeling of the geometry in computer aided design (CAD) software through dynamic variables that define the complex configuration. Grain burn back is achieved by making new surfaces at each web increment and calculating geometrical properties at each step. Geometrical calculations are based on volume and change-in-volume calculations. Equilibrium pressure method is used to calculate the internal ballistics. Genetic algorithm (GA) has been used as the optimizer because of its robustness and efficient capacity to explore the design space for global optimum solution and eliminate the requirement of an initial guess. Average thrust maximization under design constraints is the objective function.

Transient thermal analysis of Vega launcher structures

A transient thermal analysis is carried out to verify the base cover thermal protection system of Vega 2nd stage Solid Rocket Motor (SRM) and the flange coupling of the inter-stage 2/3. The analysis is performed with a finite element code. The work has developed suitable numerical Fortran subroutines to assign radiation and convection boundary conditions. The thermal behaviour of the structures is presented.

Thrust misalignments of fixed-nozzle solid rocket motors

Major sources of thrust misalignments are described, and both ground test and flight data for several different upper stage solid rocket motors are presented. Special test data and flight data statistics are presented to show the known sources, such as 1) manufacturing mechanical tolerances, 2) motor case distortions and nozzle cant angles as a result of pressurization, 3) nonsymmetrical nozzle erosion, 4) transient gas flow phenomenon from internal ballistics, 5) center of mass radial offset as a result of static unbalance, and 6) static unbalance produced by propellant bore concentricity errors. Nozzle canting as a result of pressurization has been measured during hydrotest, and static firing was found to be the largest source of thrust misalignment in several composite case motors. Transient moments or side forces have been observed early in the burn period and during rapid thrust tailoff. A brief set of guidelines is presented for developing and testing new solid rocket designs to estimate the disturbing torque.

Results in orbital evolution of objects in the geosynchronous region

The orbital evolution of objects at or near geosynchronous orbit (GEO) has been simulated to investigate possible hazards to working geosynchronous satellites. Orbits of both large satellites and small particles have been simulated, subject to perturbations by nonspherical geopotential terms, lunar and solar gravity, and solar radiation pressure. Large satellites in initially circular orbits show an expected cycle of inclination change driven by lunar and solar gravity, but very little altitude change. They thus have little chance of colliding with objects at other altitudes. However, if such a satellite is disrupted, debris can reach thousands of kilometers above or below the initial satellite altitude. Small particles in GEO experience two cycles driven by solar radiation: an expected eccentricity cycle and an inclination cycle not expected. Particles generated by GEO insertion stage solid rocket motors typically hit the earth or escape promptly; a small fraction appear to remain in persistent orbits.